airframe & systems

021 AIRFRAME & SYSTEMS

© G LONGHURST 1999 All Rights Reserved Worldwide

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http://www.click2ppsc.com/

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CLICK2PPSC LTD

EDITION 2.00.00 2001

This is the second edition of this manual, and incorporates all amendments to previous editions, in whatever form they were issued, prior to July 1999.

EDITION 2.00.00 © 1999,2000,2001 G LONGHURST

The information contained in this publication is for instructional use only. Every effort has been made to ensurethe validity and accuracy of the material contained herein, however no responsibility is accepted for errors ordiscrepancies. The texts are subject to frequent changes which are beyond our control.




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NTENTS

risation

ics

TABLE OF CO


Airframe and Systems

Landing Gear Systems

Hydraulics

Air Conditioning & Pressu

Ice and Rain Protection

Fuel & Fuel Systems

Electrics-DC

Electrics-AC

Computer, Binary and Log

Basic Radio Theory



NTENTSd Construction

and Cooling

Starting Systems

entation and

e

smissions and

nd Handling

TABLE OF CO


Piston Engine Principles an

Piston Engine Lubrication

Piston Engine Ignition and

Piston Engine Fuel Supply

Piston Engine Power AugmPerformance

Piston Engine Performanc

Piston Engine Power TranPropellers

Piston Engine Operation a



NTENTSOperation

Part 1 – The Cold

Part 2 – The Hot

s

rmance and

roplane

TABLE OF CO


Gas Turbine Principles of

Gas Turbine ConstructionSection

Gas Turbine ConstructionSection

Gas Turbine Engine System

Gas Turbine Engine PerfoOperation

Auxiliary Power Units

Emergency Equipment-Ae



021 Airframe & Systems



Aircraft Structures



Chapte

usually comprises five major units. tailplane), the landing gear and thers, are illustrated at Figure 1-1.

FIGUAircraMajor Comp


r 1 Page 1 © G LONGHURST 1999 All Rights Reserved Worldwide

1Airframe and Systems

Aircraft Structures1. The structure of the aircraft is known as the airframe andThese are the fuselage, the wings, the stabilising surfaces (fin andflying control surfaces. These major components, plus many othe

RE 1-1ft Structure

onents



Chapte

trong to withstand the forces actinglso, they must be joined together by combination of methods is used, itse junctions will be subjected.

made of materials that are strongxcessively or twisting. However, asthe wings must be able to flex. The

pon them tend to twist or bend thet is important that the wings, whilstd.

isting force (torque) applied to theust be strong enough to resist thisr bend, otherwise it might snap likeder is deflected left or right.

ort the weight of the aircraft on theloads when the aircraft turns duringequally, of course, to the points of

produce an aircraft constructed ofthe airframe will be subjected. The

loads and rigid where necessary to



2. Each of these airframe components must be sufficiently supon it during all stages of flight without distortion or failure. Abolts, screws, rivets, welding and so forth. Whichever method ormust be of sufficient strength to withstand the loads to which the

3. The wings support the aircraft in flight so they must beenough to withstand the aerodynamic forces, without bending ethese forces vary at different flight speeds or during turbulence, same applies to the junction between wings and fuselage.

4. When the rudder or elevators are used, the forces acting ufuselage, which must be strong enough to resist this. Similarly, iable to flex up and down, do not twist when the ailerons are use

5. When the elevators are deflected up or down there is a twhorizontal stabiliser and its attachment to the fuselage. Both mtwisting force, but the stabiliser must be supple enough to flex, oa dry twig. The same requirements exist for the fin, when the rud

6. The landing gear must be strong enough not only to suppground, but also to withstand the shock of landing, the twisting taxiing and the bending loads at touchdown. All this applies attachment of the landing gear to the airframe.

7. The aircraft designer must consider all these factors andmaterials strong enough to withstand all of the loads to which aircraft must be flexible where necessary to absorb changing prevent twisting.



Chapte

designer arrives at a compromise,eight to a minimum. This normallyto fail at an ultimate load that is 1½ate load to maximum applied load

ces stress within that material. This this is called strain.

airframe are tension, compression,

a material apart. This is illustrated being subjected to tensile stress.Thehen the aircraft is stationary on the



8. Having calculated the maximum anticipated loads the which gives sufficient structural strength but keeps airframe wensures that each of the various parts of the structure is designed times greater than the maximum applied load. The ratio of ultim(1.5:1) is known as the safety factor.

Stress9. The application of force to a given area of material industress will cause the material to change its shape, or deform, and

10. The stresses that act upon the component parts of the bending, torsion and shear.

Tension11. Tension is the stress that resists the forces tending to pullat Figure 1-2. The cable supporting the weight is in tension, or ispylon from which an under-wing engine is slung is in tension wground with the engine stopped.



Chapte

FIGUTensio

squeezing force, as illustrated atWhen an aircraft is standing on thes.



RE 1-2n

Compression12. Compression is the stress that resists a crushing or Figure 1-3. The material beneath the weight is in compression. ground the landing gear struts are subjected to compression stres



Chapte

FIGUComp

a material is bent it is subjected toof the material is being increased insed. This is illustrated at Figure 1-4. the upper surface is in compression



RE 1-3ression

Bending13. Bending involves both tension and compression. When both tension and compression stress. This is because one side length, or stretched and the other is being shortened, or compresWhen an aircraft wing is bent upwards due to increased loadingand the lower surface is in tension.



Chapte

FIGUBendin

g force applied to a wing when there. This twisting force is known as

e layer of material to slide over anjoint with rivets. If the assembly ishis is illustrated at Figure 1-5. Wereterial would be in direct shear.

FIGUShear



RE 1-4g

Torsion14. Torsion is the stress that resists twisting. Thus, the twistinaileron is deflected sets up torsional stress in the wing structutorque.

Shear15. Shear is the stress that resists a force tending to cause onadjacent layer. Suppose two metal panels were joined by a lap placed in tension a shearing stress will be set up in the rivets. Tthe two components joined by an adhesive bond, the bonding ma

RE 1-5



Chapte

spread the loads such that stressesise be subject to failure. The major

rs, freight, systems and equipment,e are, typically, the torsion from theingle-engine aircraft), bending one wings in flight. There are three (or truss), monocoque and semi-

, in modern aircraft, of steel tubes.s, joined together by lateral braces.ith intermediate diagonal braces asa Pratt truss.



16. The components of the airframe are constructed so as toare not concentrated at any particular point, which would otherwflight loads are borne by the aircraft wings and fuselage.

Fuselage Construction17. Besides providing the accommodation for crew, passengethe fuselage must be able to withstand the stresses of flight. Thesempennage (rudder and elevators) and the propeller (in a stouchdown and tension and compression transmitted from thcommon forms of fuselage construction known as steel tubemonocoque.

18. The truss type of fuselage comprises a framework madeThe principal components are longitudinal tubes called longeronThe lateral members may be perpendicular to the longerons, wshown at Figure 1-6, in which case the construction is known as



Chapte

FIGUSteel TTruss LAircraConst

s the Warren truss is used, whiched at Figure 1-7.



RE 1-6ube or ight

ft ruction

19. In many aircraft an alternative type of truss, known aemploys only diagonal braces between the longerons, as illustrat



Chapte

FIGUWarre

ression and tension stresses, due toely carried by the truss componentsing of the truss members is reversed

ing concentration at any one point.



RE 1-7n Truss

20. The basic concept of truss construction is that the compthe bending that a fuselage is primarily subjected to, are alternatas shown at Figure 1-8. When bending loads are reversed the loadand so stresses are spread evenly over the whole structure, avoid



Chapte

FIGUAvoidiConce



RE 1-8ng Stress ntration



Chapte

ed in truss construction nowadays,, often with steel wire forming some is limited to light aircraft fuselages. it carries no load. In earlier aircraft

struction the strength to maintainhere are no bracing members, only

kin must take all the loads this typethe skin thickness necessary would limited to small, narrow fuselages.



21. Although steel tubing is the material most commonly uswood and aluminium have both been extensively used in the pastof the bracing members. As a general rule, truss type constructionThe fuselage skin is usually made of thin gauge aluminium, sincetypes the skin was often fabric or plywood.

Monocoque Type22. The name means ‘single shell’ and in this type of confuselage rigidity and withstand stress is all in the fuselage skin. Tformers to maintain the desired shape of the fuselage. Since the sof construction is unsuited to large diameter fuselages because incur a high weight penalty. Hence, monocoque construction isAn example is shown at Figure 1-9.



Chapte

FIGUMonocConst



RE 1-9oque Type

ruction



Chapte

truction is high strength aluminium

r most aircraft fuselages, especially this a form of semi-monocoquebearing skin material and take some

e longerons. Formers called frames,e of this form of construction is thatvent of considerable damage, since being concentrated in the frames origure 1-10.

ression stress due to bending whilstrs are also the attachment points for

rincipally metal, with high strengthircraft. In larger aircraft steel andts. Secondary and non load-bearingte-based compounds and compositeminium and fibreglass honeycomb



23. The material most commonly used for monocoque consalloy, 2024 duralumin being a typical example.

Semi-Monocoque24. Neither truss nor monocoque construction is suitable fowhere large, pressurised aircraft are concerned. Because ofconstruction is used which employs longerons to brace the load-of the loads.

25. Shorter longitudinal members call stringers supplement thrings and bulkheads maintain fuselage shape. The main advantagit is capable of maintaining its structural integrity even in the eloads and stresses are spread over the whole structure rather thanskin. An example of semi-monocoque construction is shown at F

26. The longerons and stringers absorb the tensile and comptorsional stress is taken up by the skin. The longerons and stringethe skin.

27. The materials used in semi-monocoque construction are paluminium alloy being the commonest, especially in smaller atitanium alloys are often used for major load-bearing componencomponents are increasingly made from fibreglass, kevlar, graphimaterials. Cabin floors, for example, are often made from alusandwiched between aluminium sheeting.



Chapte

FIGUSemi MConst

bination of structural methods mayuction for the forward fuselage andone.

ocoque construction and formed ofmprises a streamlined nose section

on to which the wings are attached



RE 1-10onocoque

ruction

28. In many aircraft fuselages, especially smaller types, a combe used. Some Cessna designs, for example, use steel truss constrcockpit area and semi-monocoque for the rear fuselage and tail c

29. Large transport aircraft fuselages are usually of semi-mona number of sections joined end-to-end. The simplest format coincluding the flight deck, a parallel-sided cylindrical cabin sectiand a tapered tail section carrying the empennage.



Chapte

g the length of the fuselage, some ofed together by many longitudinalost cases longitudinal strength andde base of the structure, known as arough the wing centre section area.selage construction principle of the



30. Strong circular frames are spaced at regular intervals alonwhich are reinforced to form bulkheads. The frames are joinstringers, to which the load-bearing outer skin is attached. In mrigidity is supplemented by a stout beam extending along the insikeel beam. The keel beam runs along the fuselage centreline thThe general concept is illustrated at Figure 1-11, showing the fuLockheed L-1011.



Chapte

FIGUKeel B



RE 1-11eam



Chapte

ect are bending stress, since all theraft, hoop stress. Hoop stress is due pressure and places the frames and

and impact damage, such as bird-d forward vision. To achieve theselly in larger aircraft. By assembling strength lie perpendicular to eachsparency of similar thickness. Thin,nations to maintain the windshields

cockpit windows must be strongltitude and must be recessed into a internal pressure.

ge frames and are strengthened byminium alloy frames. The latter areselage to withstand pressurisationconcentration, which could lead tosion and contraction of successivew openings is limited in pressurised



31. The major stresses to which the aircraft fuselage is subjweight is borne at the wing centre section and, in pressurised aircto the tendency of the fuselage to expand because of the internalskin in tension.

Flight Deck and Cabin Windows32. The cockpit windows must be strong enough to withststrike, and must remain clear to afford the pilots uninterrupterequirements they are usually of laminated construction, especiathe pre-stressed laminations so that the directions of principalother much greater strength is achieved than with a single trantransparent electrical heating mats are layered between the lamifree of frost or condensation.

33. As with all other windows in pressurised aircraft, theenough to withstand the force due to differential pressure at astrong framework to prevent them being blown outwards by the

34. Cabin window openings are centred between the fuselaaluminium doublers, or reinforcing plates, around the strong alurecessed so that the window panel is fitted from inside the fuforces. The openings have well-rounded corners to avoid stress stress cracking and fatigue failure under the repetitive expanpressurisation cycles. For the same reason the size of cabin windohulls.



Chapte

ial, such as Perspex, and this is alsokpit windows of pressurised aircraft

the aircraft airborne. The wings,ere is considerable upward bendingat the point of attachment to theng force about the lateral centrelineough to withstand the bending and

lage attachment points must be ablees acting on the wings and by theng to separate the wings from the

t construction, the loads are in partings are designed on what is knowned entirely by the wing structural

e carried by one or more transversee in which almost all the stresses areconstruction.



35. Cabin windows are usually made of strong plastic materused for the cockpit windows of un-pressurised aircraft. The cocare usually made from strengthened glass.

Wing Construction36. The wings generate virtually all of the lift that keepstherefore, support the remainder of the aircraft. Thus, in flight, thforce acting upon the wings and this is largely concentrated fuselage. In addition the ailerons, when deflected, apply a twistiof the wings. Consequently, the wing structure must be strong entorsional stresses, which are trying to deform the wing. The fuseto withstand the stresses imposed by the upward bending forctwisting forces applied by the ailerons, both of which are tryifuselage.

37. In some aircraft, where the wings are necessarily of lightaken by bracing struts and wires. In most cases, however, the was the cantilever principle, where structural rigidity is providmembers.

38. The bending stresses to which the wing is subjected may bbeams, known as spars, or by building the wing as a box structurcarried by the external skin. The latter is known as stressed-skin



Chapte

f the centre of pressure, is taken upde the aerofoil shape. Stringers run the skin and to provide additional

a single spar as the name suggests,ure 1-12 and a two-spar wing at

spars are uncommon.



39. Torsional stress, due largely to the effects of movement oby chordwise ribs that give greater rigidity. The ribs also provispanwise, between the spars, to provide attachment points forspan-wise rigidity.

40. Wings of spar construction are either monospar, having two-spar or multi-spar. A monospar wing is illustrated at FigFigure 1-13. Multi-spar wings, having more than two span-wise



Chapte

FIGUMono



RE 1-12Spar



Chapte

FIGUTwo Sp

med into a beam either by extrusionction are shown at Figure 1-14.



RE 1-13ar

41. The wing spars of modern aircraft are made of metal, foror by built-up construction. Some examples of spar beam constru



Chapte

FIGUTypes Spar C



RE 1-14of Metal onstruction



Chapte

inium alloy, although a few highlyebs. The attachment points betweensteel alloy in large aircraft. Typical



42. The material most commonly used is high strength alumstressed aircraft, particularly military types, use titanium spar wwing spar and fuselage centre section are often of titanium or attachment point design is illustrated at Figure 1-15.



Chapte

FIGUTypicaAttachDesign



RE 1-15l ment Point



Chapte

section of the wing and have to being. These forces are much less thanntly of relatively light construction.ght aluminium alloy sheet.

rs running parallel to the spars, to skin is usually of light aluminium

running span-wise to which the ribsstressed skin metal is riveted to the1-16. The stressed skin material is half a millimetre in small aircraft to



43. The wing ribs are the formers that maintain the aerofoilstrong enough to resist the torsional stress tending to twist the wthe bending loads carried by the spars and the ribs are consequeThey are typically either pressed as one piece, or built up from li

44. Spanwise stiffness is complemented by the use of stringewhich the wing skin is attached. In spar constructed wings thisalloy sheet.

45. Stressed-skin wings have no spars as such, but shear websare attached. In turn, stringers are attached to the ribs and the stringers to form a load bearing ‘box’ as illustrated at Figure usually high tensile aluminium, the thickness ranging from aboutas much as 16 millimetres in a large transport aircraft.



Chapte

FIGUStresseWing Const

leviated to some extent by applying achieved by wing-mounted enginesith fuselage-mounted engines, or tome aircraft are fitted with aileronsard force at the outer wings. These



RE 1-16d-Skin

ruction

46. The bending stress acting upon a wing in flight can be aldownward forces to oppose the upward force of lift. This can beand by the weight of fuel in outboard fuel tanks. In aircraft wcompensate when fuel in the outboard tanks has been used, sobiased toward the up position to provide a stress-relieving downwconcepts are illustrated at Figure 1-17.



Chapte

FIGUWing-Stress



RE 1-17Bending Relief



Chapte

maximum permissible weight of anorne with no fuel in the wings, thedesign limiations. Since there willeded.

ear likely to sustain damage but the the large forces as the wings move to excessive loading stress.

at the tail and are known as thetabiliser (tailplane) and elevator, for and rudder for directional stability



Maximum Zero-Fuel Weight47. The maximum zero-fuel weight (MZFW) is defined as theaeroplane with no usuable fuel. If an aeroplane were to be airbupward bending stress on the wing structure would exceed its always be fuel in the wings, then these limiations will not be exce

48. In a heavy landing for example, not only is the landing gwing spar attachment points are likely to sustain damage due torapidly downwards. In addition, the fuselage might be subjected

Stabilisers49. The stabilising surfaces of most aircraft are located empennage. Typically they comprise the horizontal surfaces of slongitudinal stability and control, and the vertical surfaces of finand control. These are illustrated at Figure 1-18.



Chapte

FIGUStabilis

sers and associated control surfacesional fin and stabiliser layout, withl by the elevators, the fin and rudderure 1-19.



RE 1-18ing Surfaces

50. In a few light aircraft the horizontal and vertical stabiliare combined in a Vee tail configuration. Instead of the conventdirectional control provided by a rudder and longitudinal controare omitted and the stabilisers are set at an angle as shown at Fig



Chapte

FIGUVee Ta

known as ruddervators, since they

h ruddervators move downward inddervators move upward in unison,

ove in opposite directions to giveder pedal, for right yaw, causes the

ve up. This will apply a force to the

tor will move down and the right



RE 1-19il

51. The control surfaces attached to the angled stabilisers areperform the functions of both rudder and elevators.

52. When the pilot's control column is moved forward, botunison and when the control column is moved rearward both ruthus providing longitudinal (pitch) control.

53. When the rudder pedals are moved, the ruddervators mdirectional (yaw) control. For example, pushing on the right rudright ruddervator to move down and the left ruddervator to moleft on the tail assembly, causing the aircraft to yaw to the right.

54. When the left rudder pedal is pushed, the left ruddervaruddervator will move up.



Chapte

) are operated together the controlffect. To initiate a pitch up and yawal forward) the left ruddervator willurface will depend upon the bias of right surface will move up also, butrface will remain stationary or only

same as that for the aircraft wings,s and their associated twisting and

cal stabilisers (fin and tailplane) arevator and rudder) hinged to the rearthe front spar, which is the opposite structures is primarily aluminium

d to exert little effort in steady levele, or stabilise, the lift/weight couple.hen the elevators are deflected. Theected.

ter, a form of structural vibration ineads to bending due to aerodynamicof aerodynamic bending and so on.hich ultimately results in structuralhe structure, to prevent the initialbe found in the Principles of Flight



55. When both controls (control column and rudder pedalssurfaces movement is differentially biased to achieve the desired eto the right, for example, (control column back, right rudder pedmove up (yaw right, pitch up) whilst the movement of the right sthe yaw/pitch requirement. If more pitch than yaw is required theless than the left. If more yaw than pitch is required the right sumove up a very small amount.

56. The construction of the stabilising surfaces is basically thebut on a considerably smaller scale, since far lower lifting forcebending stresses are involved. Typically the horizontal and vertiof twin-spar construction, with the associated flying controls (elespar. Because of this the rear spar is usually much stronger than of the usual wing construction. The material used in stabiliseralloy.

57. The horizontal stabilising surface (the tailplane) is requireflight, usually developing a relatively small negative lift to balancThe effort exerted by the tailplane only changes significantly wvertical stabiliser (the fin) exerts no effort until the rudder is defl

58. Horizontal stabilisers are susceptible to aerodynamic flutwhich the surface twists under load. The twisting of the aerofoil lforce, which leads to twist in the opposite direction, a reversal The cycle of twisting and bending usually increases with time, wfailure. The solution is to increase the torsional stiffness of ttwisting. A full explanation of the various flutter modes can book.



Chapte

ssure rapidly moves longitudinally,y the horizontal stabiliser. In largend therefore stress on the stabiliseraircraft required to operate at highis automatically adjusts horizontal incidence, to maintain longitudinalture.

tural or non-structural. Structuralanother, or absorb forces. Such

bs, fuselage, bulkheads etc. Non-d include such items as landing gear

load cycles imposed during take-of,the years by using different design.

and to allocate the structure, a safe



59. During flight at high Mach numbers the centre of preresulting in large changes in the aerodynamic force generated baircraft this could cause rapid changes in the stabilising effort astructure. To compensate for the centre of pressure movement, Mach number are equipped with a Mach trimming device. Thstabiliser effort, by adjusting either the elevators or the tailplanetrim and prevent rapid reversals of loading on the stabiliser struc

Design Philosophies60. Components of an aircraft structure are known as struccomponents transfer loads and forces from one location to components are those already described such as wing spas, ristructural components do not transfer loads or absorb forces, andoors, cabin doors and nose radar radomes.

61. The life of an airframe is limited by fatigue, caused by thelanding and pressurisation. This lift has been calculated over philosophies, these being safe-life, fail-safe, and damage-tolerant

Safe-Life62. This original design philosophy involved testing to failurelife of 25% of the average life when tested to destructive failure.



Chapte

t the loads of adjacent componentslure. However, precise inspectiontudies of crack growth. In addition,tions (as was fail-safe) and did not as climate conditions and aircraft

onents will have minor flaws and laid down in order that these flawserant design which is now used for

rvicing to various levels for generalwill thus be necessary to establish ance lines and station numbers forong the fuselage are referenced to aage. The station numbers are givenns are measured from the fuselage



Fail-Safe63. In this construction, components were designed to accepshould one of the latter fail. The philosophy anticipated faitechniques were not specified, and there were, for example, no stesting was done with new components under laboratory conditake into consideration the deterioration caused by such thingsutilisation.

Damage Tolerant64. This design philosophy accepts that production companticipates their growth. Precise inspection procedures are thenmay be identified before they become critical. It is damage-tollarge transport aeroplanes.

Station Numbers65. During operational service, an aeroplane will require semaintenance, and will also require the replacement of parts. It method of locating various components by establishing referevarious parts of the aeroplane structure. For example, points alzero datum which is usually at the front of, or close to, the fuselas distances forward (or aft) of the datum line. Wing statiocentreline.



Chapte

o provide control about one of thehe ailerons for control in roll, the

ard trailing edge of the wings andease lift at that point and the otherrodynamic forces produce a rollingperated in some cases by sidewaysoften by rotation of a control wheelwheel to the right deflects the righting the control wheel to the left has

edge of the horizontal tail surfaces.ncreased, creating positive lift aft ofout the lateral axis of the aircraft.. They are operated by fore and afteflects the elevators down for pitch



Primary Controls66. The function of the primary flying control surfaces is tthree primary axes of roll, pitch and yaw. These are usually televators for control in pitch and the rudder for control in yaw.

67. The ailerons, for roll control, are mounted on the outbomove differentially when deflected, one hinging upward to decrhinging downward to increase lift. The resulting differential aemoment about the longitudinal axis of the aircraft. They are omovement of the pilot’s control column (or side stick), but more attached to the top of the control column. Rotating the control aileron up and the left aileron down, for roll to the right. Rotatthe opposite effect.

68. The elevators, for pitch control, are hinged to the trailingWhen deflected downward the positive camber of the surface is ithe aircraft CG to produce a nose-down pitching moment abUpward deflection of the elevators produces the opposite effectmovement of the pilot’s control column. Forward movement ddown; aft movement deflects the elevators up for pitch up.



Chapte

ing edge of the vertical stabiliser, ormetrical aerofoil, creating lift in ays lift to the right, aft of the aircrafthich yaws the nose to the left. Rightoperated by the pilot’s rudder footd deflects the rudder to the left and

aces is directly proportional to the

ed in a single set of control surfaces. ailerons and which are mounted atlly for roll control and up or down

s elevons, except that they are tail- Vee, or butterfly tail aircraft and

nder stabilisers.

d the primary control surfaces is, inent of the pilot’s controls is directlys or wires passing through guiding

controls depends entirely upon theincrease as the square of the truequently airworthiness requirementsrols which should not have to beices may be used, which reduce the



69. The rudder, for directional control, is hinged to the trailfin. When deflected it produces camber on an otherwise symsideways direction. Left deflection of the rudder produces sidewaCG, creating a yawing moment about the aircraft normal axis, wdeflection of the rudder has the opposite effect. The rudder is pedals, attached to the rudder bar. Pushing the left pedal forwarvice versa.

70. In all cases the degree of movement of the control surfdegree of movement of the pilot’s controls.

71. On some aircraft the effect of two of the above is combinExamples are elevons, which combine the effects of elevator andthe wing tips of some swept-wing aircraft. They move differentiatogether for pitch control. Tailerons are essentially the same amounted rather than wing-mounted. Ruddervators are used oncombine the effects of rudder and elevator, as explained earlier u

72. The method of actuation between the pilot’s controls anmany aircraft, direct mechanical. In this type of system, movemtransmitted to the appropriate control surfaces by means of rodpulley wheels. The force exerted by the pilot in moving the aerodynamic forces acting upon the control surfaces, which airspeed. Individual pilot strengths will clearly vary and consestipulate specific load limits for conventional wheel type contexceeded. To assist the pilot a variety of control balancing devaerodynamic force to be overcome.



Chapte

er it will try to return the control tosurface to the neutral position is theendicular distance of the centre ofe hinge moment and is illustrated at

FIGUHinge

ol surface will vary directly as theo overcome the hinge moment andlowest airspeeds he or she could doorm of control balance devices.

e moment, or by setting up a force



Aerodynamic Balance73. When a control surface is deflected the airflow acting ovthe neutral position. The total force trying to return the control product of the lift force on the control surface and the perppressure of the control surface to the hinge line. This is called thFigure 1-20.

RE 1-20Moment

74. The magnitude of the lift force generated by any contrsquare of the EAS. The pilot is required to provide the force tdeflect the control surface (in a manual system). At all but the with some form of assistance. This assistance is supplied in the f

75. Control balancing is achieved either by reducing the hingthat acts against the hinge moment.



Chapte

. The hinge is set back towards thethe surface forward of the hinge anding the aerodynamic force aft of the

esign to ensure that the centre ofe is operated its centre of pressure the inset hinge is too small there is

reversing the direction of the hingeed at Figure 1-21.

FIGUInset H



Nose Balance or Inset Hinge

76. This is perhaps the simplest form of aerodynamic balanceCP of the control surface so that, when it is deflected, air strikes reduces the force needed to move the control by partially balanchinge.

77. With this type of balance, care must be taken in the dpressure is not too near the hinge line. When a control surfacmoves forward. If the margin between the centre of pressure anda possibility that the CP will move forward of the inset hinge, moment and is known as overbalance. An inset hinge is illustrat

RE 1-21inge



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. Although it is shown here on aevators. In this system a portion ofroduces a moment in opposition to

FIGUHorn B



Horn Balance

78. An example of a horn balance is shown at Figure 1-22rudder, horn balances can equally well be used on ailerons or elthe control surface acts ahead of the hinge line, and therefore pthe hinge moment.

RE 1-22alance



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lance panel, often referred to as amber to a fixed structure (eg. spar).ure differential between upper andes are fed to the chamber to provideed with no increase in exterior drag.

FIGUInterna



Internal Balance

79. A projection of the control surface in the form of a ba‘beak’, is connected by a flexible diaphragm within a sealed chaSee Figure 1-23. Control surface movement produces a presslower surfaces and these upper and lower surface pressure changa partial balancing moment. Internal balance is therefore achiev

RE 1-23l Balance



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serves the purpose of reducing thea given airspeed. A balance tab ise way, so the balance tab moves the of the primary control, the momentty in terms of drag, and diminishesit is fitted.



Balance Tab

80. Like the inset hinge and the horn balance the balance tabstick forces involved in moving a primary control surface, at shown at Figure 1-24. As the primary control surface moves onother. Since the tab is a considerable distance from the hinge lineproduced by it is large. The balance tab imposes a small penalslightly the effectiveness of the primary control surface to which



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FIGUBalanc



RE 1-24e Tab



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ving a control surface against thein very low control column loads, ato excessive deflection of the controlse to the centre of pressure of thee tab is fitted which operates in the

strated at Figure 1-25.



Anti-balance Tab

81. In some aircraft, far from requiring assistance in moaerodynamic loads, the hinge moment is too small. This results lack of feel and the possibility of over-stressing the airframe due surface. This often occurs because of the hinge being too clocontrol surface. In order to improve the situation an anti-balancsame direction as the control surface. An anti-balance tab is illu



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FIGUAnti-B

ered necessary at low airspeeds, buterodynamic loads. The spring tablance tab is shown at Figure 1-26.



RE 1-25alance Tab

Spring Tab

82. With many aircraft aerodynamic balancing is not considis progressively required as airspeed increases, and with it the asystem may then be fitted to deal with this situation. A spring ba



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FIGUSpring



RE 1-26 Tab



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lever pivoted on the main controle is through springs, and with lowitted to the main control surface

ometry between the primary controlncrease at high speed, in order tocontrol surface, the spring becomesnto operation the balance tab on thery control surface, thus assisting the

rol surfaces the loads involved, evenircumstances servo tabs are used to

section, once again attached to theated directly by the control column,main control surfaces. As with theves in the opposite direction to the.



83. The movement of the control column is transmitted to asurface but not directly operating it. Operation of this surfacaerodynamic loads the movement of this pivot arm is transmthrough the springs; consequently there is no alteration in the gesurface and the balance tab. When the aerodynamic loads itransmit the control column movement via the pivot arm to the compressed. This upsets the geometry of the system and brings itrailing edge, which moves in the opposite direction to the primapilot by reducing the stick forces involved.

Servo Tab

84. When manual controls are used to operate very large contwith balance tab assistance, may be unacceptable. Under these coperate the control surfaces. A servo tab is a small aerofoil trailing edge of the main control surface. The servo tab is operwith no direct connection between the control column and the balance tab and the trim tab (described later), the servo tab moprimary control surface. A servo tab is illustrated at Figure 1-27



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FIGUServo



RE 1-27Tab



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ircraft flying at high speeds makes ity manual controls. It is therefore

power-assisted controls and fully

ovement of the control surface is partly by the power system. Hereface loads. Should a fault or powerailable, and control will continue towill be relatively high. Trim controlthat for manually operated flying

r systems are provided which, whilehe force necessary for operation ofed to actuators, which provide the



Powered Controls

85. The force needed to move the control surfaces of a large avirtually impossible for satisfactory control to be exercised bnecessary that the primary control surfaces be power operated.

86. Powered controls may be divided into two categories,power operated controls.

Power Assisted Controls

87. With power-assisted controls, the force needed for mprovided partly by the physical force produced by the pilot andthen the pilot will have ‘feel’ which is provided by the control surfailure occur in the control system, a disconnect system will be avbe maintained by manual means alone, although control loading for a power-assisted control is provided in the same way as controls.

Fully Powered Controls

88. Where fully power operated controls are installed, poweindependent of each other, operate in parallel and provide all tcontrol surfaces. Movement of the pilot’s controls is transmittforce necessary to move the control surfaces.



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ed by manual reversion, or by more conventional trailing edge tabs willzero positions of the artificial feeled for the maintenance of servo andhown at Figure 1-28.

FIGUPowerContro

ns that there is no feedback ofthe pilot has no feel through the



89. Safeguards against faults or power failure may be providthan one system, each with its individual hydraulic circuit. Thusnot be included and trim would be obtained by altering the mechanism. Occasionally, where necessary, balance tabs are fitthinge loads. An example of a power operated control system is s

RE 1-28ed Flying l System

90. Power operated controls are irreversible, which meaaerodynamic forces from the control surface. Consequently, controls for the aerodynamic loading on the control surfaces.



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artificial feel rarely has any directl can be provided by a spring thatore the control is moved the greaterem is that the resistance to controlte force at low airspeed would behigh airspeed would be too great at

e loading which varies in directal to dynamic pressure (q). Thet and static pressure is sensed and

ement proportional to airspeed andnual controls. In its simplest form ad to one side and static pressure to dynamic pressure. The piston is



91. Feel is provided by artificial methods and, in fact, the relationship to the forces working on the control surface. Feeexerts a constant load for a given control position, so that the mthe spring force to be overcome. The disadvantage of this systmovement is the same irrespective of airspeed. A proportionainadequate at high speed, alternatively a proportionate force at low speed.

92. A much more satisfactory arrangement is to providrelationship to airspeed. This loading is a force proportionarrangement is therefore commonly referred to as q feel. Pitoapplied to the pilot’s controls to produce a force resisting movtherefore representative of the force the pilot would feel with maq feel system comprises a large piston with pitot pressure appliethe other. The overall force acting on the piston is thus due toconnected to the pilot’s controls as shown at Figure 1-29.



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FIGUSimple

overcome by the utilisation of aressures are fed to either side of aalve provides a metered hydraulic small jack which opposes control

shown at Figure 1-30.



RE 1-29 ‘q’ Feel

93. The bulkiness of the simple ‘q’ feel system can be hydraulically enhanced system in which the pitot and static pdiaghragm attached to a hydraulic servo-valve. The servo-vpressure, which is an amplified value of dynamic pressure, to acolumn movement. An example of a hydraulic ‘q’ feel system is



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FIGUHydra

t's primary controls are transmitteds known as fly-by-wire. The systemessing which alters the response tocess control surface movement, ory to improve aircraft performance,f control surface movements too

inputs is by means of the primaryd slip indicator for yaw. Angle ofhe case of excessive pitch up.



RE 1-30ulic ‘q’ Feel

Fly-By-Wire

94. In many present-generation aircraft signals from the piloelectrically to actuators which move the control surfaces. This ilends itself to the incorporation of sophisticated electronic proccontrol inputs by the pilots to avoid stalling, over-rapid or exunstable flight regimes. Fly-by-wire not only has the capabilitefficiency and safety, it can also incorporate co-ordination ocomplex for a pilot to achieve unaided.

95. Indication of the aircraft response to the pilot’s controlflight instruments; artificial horizon for roll and pitch, turn anattack sensors may give aural stall warning, or stick shaking, in t



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wind when the aircraft is stationaryernal mechanical locking devices towith manual controls, or internall warning that such locks have been

ockpit at the operating control. On are in place.

is not to provide control about onet circumstances. Examples are flaps,ming tail-plane and trim tabs. All of

dges of the wings and are extendedapproach and take-off. Spoilers aresome of the lift generated. They arension of the landing roll and can ben and wheel spin-up, provided the reverse thrust is selected above a

upper or lower surfaces (or both) orht, during descent for example. Aence, adjusted to trim the aircraft

minimal stick force. Trim tabs arete the stick force needed to hold a



96. To prevent control surface damage due to strong gusts ofon the ground, control locks are fitted. These can be either extprevent control surface movement, usually fitted to aircraft hydraulic locks incorporated with powered flying controls. Visuafitted are provided externally at the control surface and in the csome aircraft, warning lights indicate when control surface locks

Secondary Controls97. The principal function of the secondary control surfaces of the primary axes, but to adjust the lift or trim in specific flighslats, spoilers (except when used for roll control), airbrakes, trimthese are described in detail in Principles of Flight.

98. Flaps and slats are attached to the trailing and leading esymmetrically to increase wing lift at low flight speeds, such as hinged to the wing upper surface and, when extended, destroy used during and after touch down to prevent ‘floating’ and exteselected manually or automatically on landing after touchdowlever is in the armed position, or for a rejected take-off, whenparticular speed (typically 60kt).

99. Airbrakes may be extended symmetrically from the wing from the fuselage to create drag and slow the aircraft in fligtrimming tailplane is a horizontal stabiliser with variable incidlongitudinally so that aircraft attitude can be maintained withattached to the primary control surfaces and adjusted to elminagiven control position.



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may be direct mechanical, hydraulicsize and operating speed. The pilot’snd spoiler operation and trimming

g surfaces.

indicated by the operating controlnic indicating systems.

s (flaps and slats) should always beate a strong rolling moment difficultnnecting actuators are designed to with warning systems to alert the

ing-mounted airbrakes is similarlyse controls.

e flaps from excessive air loads bylap position whenever the airspeedaps automatically return to the fully



100. The mode of actuation of the secondary control surfaces or electric powered, or fly-by-wire, depending upon the aircraft controls usually take the form of levers for flap, slat, airbrake awheels or electrical trim switch on the control wheel for trimmin

101. The position of flaps, slats, airbrakes and spoilers is position and, on larger aircraft, by position gauges and/or electro

102. It is essential that wing-mounted lift augmentation deviceextended symmetrically, since asymmetric deployment would creor impossible to counteract with the primary controls. Intercoprevent asymmetric deployment and many aircraft are equippedpilot to this potential hazard. Asymmetric deployment of wundesirable and warning devices are often incorporated with the

In addition, a flap load limiter system protects the trailing edgautomatically retracting flaps from the fully extended landing fexceeds a predetermined speed. When airspeed is reduced, the flextended position.






Main Components of Landing Gear

Landing Gear Bracing

Landing Gear Warning System

Nosewheel Steering

Wheels and Tyres

Tyre Checking Procedures

Brake Energy Capacity



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ight of the aircraft while it is on thehe main gear provides the principall installation, almost invariably the

eration are:

hard braking.

.

take-off or landing in cross-wind

the landing forces and the effects ofntrolling the aircraft on the ground.ipped with nose-wheel steering forn wheel steering (known as bodyhe landing gear is retracted into themance aircraft this is less important.



2Landing Gear Systems

1. The landing gear of a fixed wing aircraft supports the weground and is made up of the main and auxiliary landing gear. Tsupport. The auxiliary gear is in the form of a nose or tail wheeformer, especially in large aircraft.

2. The advantages of tricycle landing gear during ground op

(i) No risk of the aircraft ‘nosing over’ during

(ii) Better visibility for the pilot during taxiing

(iii) Less likelihood of ground looping duringconditions.

3. The landing gear contains shock absorbers to withstand taxiing over uneven surfaces, and brakes to stop and assist in coMost modern aircraft with tricycle (nose-wheel) layouts are equground manoeuvring, some very large aircraft also have maisteering). In aircraft designed for flight at high speeds/altitudes tfuselage or wings in flight, to reduce profile drag. In low perforand the landing gear is often fixed in the extended configuration



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rn a large aircraft.



Main Components of Landing Gea4. Figure 2-1 shows typical main landing gear components i



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FIGUMain LComp



RE 2-1anding Gear onents



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s are designed to absorb the shockhem from being transmitted to the the airframe and contains a lowerhe case of the nose strut) within theressurised with compressed air oriing and balances the weight of theinder takes up an approximate mid- cylinder moves up, shortening therence of oil from lower to upperubber valve as shown at Figure 2-2.ith the upward movement, the gas

craft bounce on landing, the shockmpression.



Oleo-Pneumatic Shock Struts The landing gear shock strutloads of landing and taxiing over uneven ground, preventing tairframe. The upper (outer) cylinder of the strut is attached to(inner) cylinder which is free to slide up or down (and rotate in touter cylinder. The cylinders are partially filled with oil and pnitrogen. This compressed gas absorbs the shocks of normal taxaircraft when it is stationary on the ground, so that the inner cylstroke position. Under the increased shock of landing the innerstrut length. To prevent excessive upward movement, transfecylinder is progressively restricted by either a metering pin or a snAs the volume of the gas space in the upper cylinder decreases wpressure increases to balance the upward force. To prevent airabsorber damping on rebound is greater than the damping on co



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FIGULandinShock



RE 2-2g Gear Struts



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as pressure. Too high a pressure andstrut extension will be inadequate,under shock loading. Checking thef the amount of the inner cylinderion to the pilot on a walkaround or graphs may be available whichn weight and loading configuration.

subject to considerable side-for/afte leg, additional support is providedly. See Figure 2-1.

gitudinal axis of the aircraft. Theympress for damping.

onft over a wide area to achieve anf landing gear wheel arrangement -



5. It is important that the struts are inflated to the correct gthe shock absorption is reduced, too low a pressure and the leading to ‘bottoming’ and complete loss of shock absorption strut gas pressures is a job for an engineer, however a check owhich is visible (the amount of extension) is a good indicatinspection that the gas pressure is approximately right. Tablesenable the pilot to determine the appropriate extension for a give

Landing Gear Bracing6. During take-off, landing and taxiing, the landing gear isloads. To prevent damage, breaking or possible collapsing of thby fitting side load struts/stays and drag struts to the gear assemb

7. Torsion links maintain the wheel alignment with the lonjoin the inner and outer cylinders but allow the shock strut to co

Main Landing Gear Wheel Configurati8. Heavy aircraft need to spread the weight of the aircraacceptable pavement loading. The four basic configurations osingle, double, tandem and bogie are illustrated at Figure 2-3.



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FIGUMain LWheeConfig



RE 2-3anding Gear l uration



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e locks are engaged by spring force

ng gear to reduce drag in flight. Thedraulically operated and sequenced

rcraft is, almost invariably, achievedlanding of the aircraft, it is of vitalg the gear in the event of hydraulicem can be selected to actuate thegear in an emergency is a gravityssure to the retraction system is shutnd door up-locks are mechanically

e down locks are engaged by spring of its operation follows.



Landing Gear Locks9. Landing gear up locks and down locks are provided. Thand broken during retraction/extension by hydraulic pressure.

Landing Gear Doors10. Landing gear doors are used to enclose the retracted landidoors may be mechanically operated by gear movement or hywith the landing gear.

Extension and Retraction11. Raising and lowering the retractable landing gear of an aiby hydraulic systems. Since the landing gear is essential to safe importance that there should be an alternate means of extendinsystem failure. In many aircraft an emergency pneumatic systlanding gear. An alternative method of lowering the landing extension or free-fall system. In this type of system, hydraulic preoff, the raise and lower lines are open to return and the gear areleased. The gear extends under the influence of gravity and thaction. Figure 2-4 shows a landing gear system and a description



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FIGULandinSystem

e appropriate side of the operating of the jacks to the reservoir returnsequence of operations when UP is



RE 2-4g Gear

12. A selector valve directs hydraulic system pressure to thjacks (up or down), at the same time connecting the other sideline. Figure 2-4 shows the gear extended, let us first follow the selected.



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recting hydraulic pressure to the upgear (NLG) down locks. Pressure isgears. As the NLG reaches the fully-a spigot on the torsion links. As the

up lock engages the detent on thence valve (SV1). This directs up line door.

directing hydraulic pressure to the operating jack to extend the nose

opened by the door jack. When thelve 2 (SV2), opening it and allowingain gear. It will be noted that a one-lve and the MLG jack. This restrictsg MLG extension, thus limiting theto hydraulic force, would otherwisell fluid flow.

ent UP selection with the aircraft ononly called squat switches, (whichs) or by mechanical locking devices



Up selection. The selector valve spool moves to the right, dilines, releasing the main landing gear (MLG) and nose landing also applied to the MLG and NLG operating jacks to retract the retracted position the spring-loaded NLG up lock engages with MLG reaches the fully-retracted position the spring-loaded MLGMLG shock strut and a pintle on the main gear opens the sequepressure to the inner door jack, which operates to close the inner

Down selection. The selector valve spool moves to the left,down lines, releasing the NLG up lock and pressurising the NLGgear. Before the main gear is extended the inner doors must be door reaches the full open position a pintle contacts sequence vahydraulic pressure to release the MLG up lock and extend the mway restrictor valve is fitted in the up line between the selector vathe flow of fluid returning from the jack to the reservoir durinrate of travel of the heavy main gear which, with gravity added be excessive. During retraction the one-way restrictor permits fu

13. Accidental retraction of the landing gear, due to inadvertthe ground, is prevented by weight-on safety switches, commisolate the selector switch when aircraft weight is on the wheelinserted by ground maintenance staff.



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with a knob in the shape of a wheeln and locked, travelling (unlocked)en the gear is down and the aircraftt is airborne and its weight is off theitions (airborne or on the ground) a/OFF/DOWN) are often used forhe landing gear retraction system in

ator for each gear (nose, left main,ate green when the gears are down,

from UP to DOWN, or vice versa, atakes the form of a flashing red lighthts are extinguished. An alternativers show:

light.



Landing Gear Control and Indication14. The landing gear control panel contains a selector handlefor UP or DOWN and indications of landing gear position - dowand up and locked. The gear operating handle may be locked whweight is on the wheels. It will only be unlocked when the aircrafwheels. In order to compare operator demands with aircraft condlogic circuit is provided. Three-position selector handles (UPadditional safety. The OFF position permits depressurisation of tflight.

15. Landing gear position is displayed by means of one indicright main) and often takes the form of three lights which illuminand locked in the down position. Whilst the gears are travelling red light illuminates to indicate ‘gear unlocked’. This sometimes in the gear operating handle. When the gear is locked up, all ligtype of indication is shown at Figure 2-5. The three gear indicato

(a) up and locked

(b) unlocked (travelling)

(c) down and locked

16. The unlocked condition is accompanied by a red warning



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FIGULandinControIndicat



RE 2-5g Gear l &

ion



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that an unsafe landing conditiong horn sounds whenever the gear is

certain conditions the warning horn

es are retarded. The warning horned.

a steering mechanism for the nose-ave main (body) gear steering also.

tandem.

mands are transmitted hydraulicallyt. Light aircraft are steered by push



Landing Gear Warning System17. The warning system provides visual and aural warningexists. The red landing gear warning lights come on and a warninnot down and locked and the aircraft is not safe to land. Under can be cancelled.

18. The landing gear red lights illuminate when the throttlsounds with a combination of flap extension and throttles retard

Nosewheel Steering19. Practically all tricycle undercarriage aircraft incorporate wheel. Larger aircraft invariably do, and some very large ones hThe latter is necessary when the main gear uses several wheels in

20. In all but the lightest aircraft the nose-wheel steering comto a yoke, or steering arm, attached to the nose-wheel shock strupull rods connected to the rudder pedals.



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hown at Figure 2-6. Rotation of the which, through a system of pulleysraulic pressure to the steering jacks shock strut. As the yoke rotates, itsontrol valve to the neutral positionheel. In many transport aircraft thell inputs being transmitted throughheel must be able to caster freely

ock strut cylinders, provided by the is an essential pre-taxiing check to

ndency to ‘caster’, that is a natural oscillation about the centre line, a

the steering mechanism and makeby means of a shimmy damper. This and filled with hydraulic fluid. Thelinder is attached to the nose-wheelm one side to the other, dampening

n large transport aircraft, consisting

he neutral (mid) position as shown,e nosewheel is free to caster. This is

whilst on the ground.



21. A schematic diagram of a nose wheel steering system is spilot's steering wheel is transmitted by cables to a steering drumand linkages, moves the hydraulic control valve. This directs hydwhich rotate the yoke attached to the lower cylinder of the NLGmotion is transmitted through the pulley system to return the cwhen the desired degree of turn has been reached by the nose wsteering wheel or tiller is only used for large steering inputs, smathe rudder pedals. For aircraft ground towing, when the nosewthrough wide angles, the connection between upper and lower shtorsion links, is broken by the removal of a quick-release pin. Itensure that this pin is engaged.

22. Nose-wheels, especially single wheels, have a natural teself-centring stability. Damping is required however to preventcondition known as ‘shimmy’.

23. Nose-wheel shimmy would exert considerable force onsteering the aircraft difficult. The castering tendency is snubbed is usually in the form of a hydraulic cylinder containing a pistonpiston rod is connected to a fixed part of the airframe and the cyleg. A small orifice in the piston allows restricted flow of fluid fropiston movement. More complex dampers are sometimes used iof a system of rotary vanes and known as a vane-type damper.

24. Note from Figure 2-6 that when the control valve is in tthe steering jacks are connected to the return line and therefore thnecessary to permit directional control by rudder at high speeds



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FIGUNosewSteerin



RE 2-6heel g



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g is its effect upon aircraft turning being turned by means of the nose-ead of the track of the nose-wheel,f the nose-wheel. Consequently the

to accommodate this. In the Boeingater than nose radius. This is shown

FIGUAircraRadius



Turning Radius An important feature of nose-wheel steerinradius, and particularly wing tip clearance. When the aircraft iswheel, the path followed by the outer wing tip will pass well ahand therefore of the pilot, whose seat position is also forward opilot must allow sufficient clearance from ground obstructions 757 for example, both outer wing and tailplane tip radius are great Figure 2-7.

RE 2-7ft Turning



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are taxi speed, gross weight, centre nose and main wheel centres and

castings made from aluminium orpart of a wheel is the bead seat areato protect from tensile loads appliedlit-hub wheels. An example of such



25. Factors which will affect the turning radius of an aircraftof gravity position, nosewheel steering angle, distance betweentrack width between main wheels.

Wheels and Tyres26. Aircraft wheels are usually constructed from forgings ormagnesium alloy to minimise aircraft weight. The most critical which is rolled, pre-stressing its area, thus increasing its strength by the tyre. Most large aircraft use tubeless tyres mounted on spa wheel is shown at Figure 2-8.



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FIGUSplit-H



RE 2-8ub Wheels



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and removal of the tyre. Obviouslyessure. The surface condition of the the tyre. The high pressures used to the wheels extremely important.Aircraft wheel hubs are made fromrone to rapid corrosion.

tons at speeds up to 250 mph. loads whilst the tyre is constantlys shown at Figure 2-9.



27. The two halves of the wheel are separated for installationwhen assembled they make an airtight seal to contain the tyre prwheel flanges is also a vital factor in preventing air escaping frominflate large aircraft tyres makes the structural integrity ofCorrosion and cracking are conditions to be guarded against. aluminium alloy, but in some cases magnesium alloy, a material p

28. Aircraft tyres must withstand aircraft loads of manyConsequently their construction is designed to withstand theseflexing during wheel rotation. Typical aircraft tyre construction i



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FIGUAircraConst



RE 2-9ft Tyre ruction



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rm ‘ply rating’ is used to identify a index of the tyre strength and doesonstruction. The marking may beR. The speed rating is included for

tact with the ground and may beon pattern is the ribbed tread. Thebs enable the tyre to disperse water,

erned variety for aircraft tyres. Theer. The most popular tread patternes around the tyre. A ribbed tread

lity mostly suited for hard surface

or unpaved airfields.

ine) on the upper sidewall to deflect

to prevent shimmy.

und, are causes of tyre wear andhstand the dynamic and static loadsir usable life, is caused by:



29. Tyres are classified by load, ply and speed rating. The tetyre with its maximum recommended load and pressure. It is thenot necessarily represent the number of cord piles used in its cimprinted in full, e.g. 10 PLY RATING or abbreviated, e.g. 10Ptyres used above 160 mph.

30. The tyre tread rubber is the abradable material in conpatterned to achieve particular characteristics. The most commribs provide directional stability and the grooves between the rireducing the risk of aqua-planing.

31. Tyre tread pattern is generally limited to a ribbed or patttread of the tyre refers to the area forming the crown and shouldis the ribbed variety which is formed from circumferential groovprovides good traction, long tread wear and directional stabirunways.

32. Patterned (diamond) tread tyres are particularly suitable f

33. Some nose wheels are fitted with a water deflector (or chwater away from rear mounted engines.

34. Twin contact tyres are used on nose wheels or tail wheels

35. Tyre flexure, and friction due to contact with the grodeterioration. The condition of the tyre must be sufficient to witof supporting the aircraft. Damage to aircraft tyres, reducing the



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-down the tyres may ‘creep’ around the inner tube in tubed tyres (fitted on tubed and tubeless tyres and the inch in width for tyres of up to 24res over 24 inches outside diameter.newly fitted, and/or when the tyre

prolonged braking could result inssure which could result in explosive aircraft wheels. These plugs melt at

re 2-10.



Creep. During braking and when the wheels spin up on touchthe wheel, which leads to wear of the tyre bead and can damageto some light aircraft). White creep indication marks are paintedwheel rim (flange) in line with each other. These marks are oneinches outside diameter and one and a half inches in width for tyThe tendency of the tyre to creep is greater when the tyre is pressure is too low.

Temperature. Build up of high wheel temperature during overheating of the tyre bead, and an increase in tyre inflation prefracture of the wheel. Fusible plugs are fitted in high performancea predetermined temperature and release tyre pressure. See Figu



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FIGUFuseab

rown. If the tyre is under-inflated it the spin up on touch-down cause when wear has reached the limits

e tread is worn to the depth of the

used until worn to the top of the tie



RE 2-10le Plug

Wear. If a tyre is over-inflated it suffers excess wear on the cwears on its shoulders (rim of the crown). Locked wheels andscuffing of the tread. It is recommended that tyres be removeddefined below:

(i) Patterned tread tyres may be used until thpattern.

(ii) Ribbed tyres with marker tie bars may be bars.



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y be worn to within 2mm of the

centre of the crown shows sign of

tyre tread.

oils) and solvents onto tyres will

anual. This is always the inflationhen tyre pressure is adjusted with

d to the rated inflation pressure. Aenerally specified and tyre pressures nitrogen for safety.

rolonged taxying or heavy braking,sult in under-inflation at normal



(iii) Ribbed tyres without marker tie bars mabottom of the wear indicator groove.

(iv) Twin contact tyres may be used until thecontacting the ground.

Cuts. Foreign objects on the runway/taxiway can cut into the

Contamination. Leakage of oil (especially some hydraulicdestroy the rubber casing.

Tyre Checking Procedures

Tyre Inflation36. Tyre inflation pressure is given in the aircraft operating mpressure with the wheel not supporting the aircraft weight. Waircraft weight on wheels an allowance of 4% should be addetolerance of 5% to 10% above this loaded inflation pressure is gup to this maximum are permitted. Tyres should be inflated with

37. If tyre pressures increase as a result of heating, due to pthe excess pressure should not be released as this could retemperatures.



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bient temperature, any tyre which isd, together with the companion tyre loaded inflation pressure should becheck; if the pressure is again more

are still hot following a landing the with the pressures of the other tyresmore below the maximum recorded, but should be rejected if a similar commercial aircraft is in the rangeockpit) may include a tyre pressure

g during manufacture or by normalare marked by green or grey dots on

e disc type, only a few light aircraftdisc brakes, whereas light generalid disc distortion due to heat, very

own as a segmented rotor brake.



38. When checking the pressure of cold tyres which are at ammore than 10% below loaded inflation pressure should be rejecteon the same axle. Any tyre which is between 5% and 10% belowre-inflated to the correct pressure and checked at the next daily than 5% low the tyre should be rejected.

39. When it is necessary to check the pressure of tyres that pressure of each tyre should be checked and noted and comparedon the same undercarriage leg. Any tyre with a pressure 10% or on the same leg should be re-inflated to that maximum pressureloss is apparent at the next check. A typical tyre pressure for a150-250 PSI. Aircraft with electronic instrument systems (glass cindication system.

Tyre Venting40. Tubeless tyres are vented to release air trapped in the casinpermeation through the inner liner. The awl hole vent positions the lower side wall.

Wheel Braking Systems41. Almost without exception, aircraft wheel brakes are of thremain fitted with drum brakes. Larger aircraft use multiple aviation aeroplanes often require only single disc brakes. To avolarge aircraft often use a variation of the multiple disc system kn



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FIGUUnbooSystem



RE 2-11sted Brake



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ystem may be used. A brake pedalrential braking. The applied brake locked on by applying a parking

h is keyed to the landing gear wheel.high friction pads to clamp on eitherking friction.

parallel to each other with mores is necessary when stopping a large,

e plate. The brake plate is splined toing the brake operating pistons is



42. On light aircraft an independent or unboosted brake sattached to each rudder pedal permits the application of diffepressure is proportional to pedal pressure. The brakes may bebrake. See Figure 2-11.

43. The single disc brake comprises a polished steel disc whicWhen the brake is applied, hydraulic pressure on a piston forces side of the disc. The greater the force applied, the greater the bra

44. Multiple disc brakes use a number of discs mounted hydraulically operated pads, to increase the braking friction. Thiheavy aircraft.

45. Figure 2-12 shows a disc brake with a single disc, or brakthe aircraft wheel and rotates with it. The torque plate carryattached to the axle and is stationary.



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FIGUSingle



RE 2-12Disc Brake



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e the piston to the right. The brakeich in turn moves on its splines totationary brake pads. Release of thepads axially out of contact with the

tant that the clearance between the, otherwise brake control authority

Figure 2-13. When brake hydraulice brake pads to grip the brake disc.ing retainer contacts the head of therake cylinder housing. When brakeback until it contacts the head of theme the friction of the grip collar, sopad linings wear, the more the pin isrance between pad lining and disc.pad wear.



46. Application of braking causes hydraulic pressure to movpad attached to the piston pushes against the brake plate, whcontact the fixed pad, effectively clamping the disc between the shydraulic pressure allows the return springs to move the brake disc.

Automatic Brake Adjusters47. As the friction linings of the brake pads wear it is imporlining and the disc, with brakes released, is maintained constantwould be progressively reduced.

48. An example of a single-disc brake adjuster is shown at pressure is applied to the piston it moves to the right, forcing thThe piston movement compresses the return spring until the sprindicator pin and pulls the pin through the grip collar in the bhydraulic pressure is released the return spring moves the piston indicating pin. The return spring is not strong enough to overcothe pin limits the return travel of the piston. The more the brake pulled through the grip collar, thus maintaining a constant cleaThe amount of pin external protrusion is an indication of brake



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FIGUSingle-Adjust

must be converted into heat energyion between brake pads and wheely aircraft. A multiple disc brake is



RE 2-13Disc Brake er

Multiple Disc Brakes49. The heavier the aircraft the greater the kinetic energy thatby the wheel brakes. This conversion takes place through frictdiscs, thus more pads and more discs are necessary with heavshown in the diagrams at Figure 2-14.



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FIGUMultipBrake



RE 2-14le Disc-



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s that retract inwards towards thefore gear retraction takes place, the

an autoretract brake system. Nosentacted by the tyres when the gear

on a rejected take-off involves theunits and main wheels. This energy a brake limitation chart to providehow to deal with hot brakes safely



50. The majority of large aircraft have main landing gearaircraft centre-line. To reduce the stress on the wheel spinning bemain wheels are braked. This braking function is achieved by gears that retract fore/aft usually have de-spin friction pads coreaches the up position.

Brake Energy Capacity51. Stopping a high speed aircraft either after landing or conversion of considerable kinetic energy into heat at the brake may be expressed in foot-pounds or joules. An aircraft may haveflight crew and maintenance personnel a means of determining and effectively.



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FIGUBrake Chart



RE 2-15Limitation



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ires and to ensure adequate brakeines the amount of energy to beht, indicated airspeed and densitya braking event may be classified as

pecial requirements.

ft.lb. Allow a brake cooling time as

ft.lb.

will blow 2 - 30 minutes after stop.

til fusible plugs have released tyre



52. The specific purpose of the chart is to avoid in-flight fcapacity at all times for a rejected take-off. The chart determabsorbed by the brakes by considering the aircraft gross weigaltitude at the time the brakes are applied. A condition zone for normal, caution or danger.

The zones may be summarised as follows:

(a) Normal zone, zone I: Below 1.0 million ft.lb, no s

(b) Normal zone, zone II: 1.0 million to 2.05 millionshown before attempting a take-off.

(c) Caution zone, zone III: 2.05 million to 4.0 million

(i) Move aircraft clear of active runway.

(ii) Use brakes sparingly to manoeuvre.

(iii) Do not set parking brake.

(iv) Allow brake cooling time as shown.

(d) Danger zone, zone IV: over 4.0 million ft.lb.

(i) Clear runway immediately, as fusible plugs

(ii) Do not apply dry chemical or quench unpressure.



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le plugs have blown.

emoval.

tic energy of the aircraft into heatting (locks) and the tyre skids on thensfer is now only between tyre and aircraft may be lost.

ircraft braking systems include skiduch a system requires a device to to apply the brakes in proportion.tly to just prevent wheel-locking.

C generator mounted in the wheel the case of AC) will be directlyskid control unit which compares it developing, the braking action is is developing, the skid control unit prevent the skid developing further.



(iii) Do not approach for 30 mins or until fusib

(iv) Allow 2 to 3 hours cooling to permit safe r

(v) Replace tyres, wheels and brakes.

Anti-Skid Systems53. The function of the wheel brakes is to convert the kineenergy, through the friction in the brakes. If the wheel stops rotarunway, the brakes have ceased to function and the energy trarunway. Furthermore, on a wet runway directional control of the

54. To prevent the wheels locking during braking, transport acontrol, or anti-skid systems. The principle of operation of smeasure wheel rotational speed (the skid-control generator) andAs rotational speed diminishes, braking force is reduced sufficien

55. The skid-control generator consists of a small DC or Aaxle. The voltage output of the generator (and frequency inproportional to wheel rotary speed. This is fed as a signal to the with the pilot's braking demands. If there is no wheel skidproportional to the pilot's pressure on the brake pedals. If a skidactivates valves to release some of the brake actuating pressure toThis is called pressure bias modulation.



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the wheels are rotating in the firstel brakes applied, the wheels would so the wheels would remain locked.m being applied during the landingown control.

s can happen on a patch of ice, thel it spins back up. This is known ass above 15 to 20 mph.

light or caption is activated on the

t Figure 2-16. At Figure 2-17 is a Boeing 757.



56. Clearly, the anti-skid sensing system can only function ifplace. If, for example, the aircraft were to touch down with wheskid and the anti-skid system would have no way of sensing this,A protection circuit in the control unit prevents the brakes froapproach. This circuit is called touch down protection or touch d

57. Should any wheel lock fully when the aircraft is rolling, aanti-skid system will release the brakes fully on that wheel untilocked-wheel skid control and is only functional at aircraft speed

58. In the case of failure of the anti-skid system a warning flight deck and the brake system becomes fully manual.

59. A schematic diagram of an anti-skid system is shown aschematic layout of the wheel braking and anti-skid system for a



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FIGUAnti-S



RE 2-16kid System



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FIGUAnti-S(B757)



RE 2-17kid System






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ion rate controller, which compares value. If the wheel deceleration rateontroller and a permanent magnetc pressure from the spool valve totains equal pressure at either end of

ure from the pilot's brake pedals ision rate exceeds the preset value theanent magnet field and causes the

2-16. This results in an increasedhe right, which causes the spool toe brakes, thus preventing the wheelsack within limits the controller will

ft stationary while the pilot is notrake hydraulic pressure to the wheelre is available.

rocedure is to depress the toe brakeake control valves in the fully ON of the parking brake (even partialne from the anti-skid valves.



60. The output of the wheel speed sensor is fed to a deceleratactual deceleration rate of wheel rotation with a preset reference(spin-down) is within limits there is no output from the rate cholds a flapper valve in its mid-position. This allows hydrauliescape equally from the jets on either side of the flapper and mainthe spool valve, centralising it. In this position, hydraulic presstransmitted directly to the wheel brake cylinders. If the deceleratrate controller produces an output signal which biases the permflapper valve to tilt, as shown in the lower diagram in Figurepressure on the left side of the spool valve and a decrease on tmove over, relieving the brake cylinder pressure and releasing thdecelerating to a locked condition. Once the deceleration rate is brestore normal brake operation.

Parking Brake61. The purpose of the parking brake is to hold the aircraoperating the brake pedals. Applying the parking brake routes bbrakes to hold them firmly ON so long as brake hydraulic pressu

62. Operation varies with aircraft types, but in general the ppedals fully and apply the parking brake, which holds the brposition. It should be noted that, in many aircraft, applicationapplication) cuts out the anti-skid system by closing the return li







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inated on the flight deck, togetherperate main wheel brakes, and in wheel brakes are operated by the

sprovided by the aircraft hydraulicof the main hydraulic systems (fore system B).

lator and a non-return valve. In theessure loss from the brake system topressure for a number of brake is also available by connecting anf loss of main hydraulic systems. In

ure.

system (accumulator) pressure, andupply and the anti-skid system arewarnings. Aircraft with electronicm has a pressure transducer in eacha computer for display on a page



63. With parking brake applied a warning indication is illumwith anti-skid failure warnings. Parking brakes usually only oaircraft with multiple main wheels often only some of the mainparking brake. See Figure 2-17.

Normal, Alternate and Reserve System64. The power for operating the wheel braking system is system. Normal braking hydraulic pressure is supplied by one example system A) and alternate braking by another (for exampl

65. The brake hydraulic system(s) always include an accumuevent of loss of hydraulic supply the non-return valve prevents prmain system and the accumulator holds sufficient reserve applications. In many modern transport aircraft reserve brakingelectric hydraulic pump to a reserve supply of fluid in the event osome cases emergency brake operation employs pneumatic press

Indications66. Flight deck brake system indications are usually of brakebrake temperature. Warning indications of failure of normal salways provided and may include aural as well as visual instrument systems may have tyre pressures displayed. The systewheel which sends a signal corresponding to tyre pressure to showing landing gear information.







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ear indicator pin shows brake life indicator is provided, wear of theeen brake piston and disc with thes are usually numerical, increasingindication is activated, since brake

perature monitoring system is fitted a central monitor and warning unitand an illuminating selector button temperature level and the gaugeny of the individual wheel brake button associated with that wheele gauge to indicate the temperature

king facility which enables the pilotke deceleration rate is less than thatd during auto-brake operation. The

y apply the brakes when the enginens the selected deceleration rate inversers. It will continue to provide



67. Brake wear is indicated at the brakes. A protruding wremaining, the pin retracts as wear progresses. Where no wearbrake pads can be determined by measuring the distance betwbrakes applied (see Figure 2-12). Brake temperature indicationwith increased temperature. Above a certain value a warning efficiency decreases with increasing temperature.

68. On some aircraft with ‘conventional’ displays a brake temwhich includes temperature sensors at each wheel, which feed toon the flight deck. The monitor has a single temperature gauge for each wheel. The monitor is calibrated to a predeterminednormally displays the highest of the brake temperatures. If atemperatures exceed the predetermined temperature the selectorwill illuminate. Pressing any of the selector buttons will cause thof the brake unit on that particular wheel.

Auto Brakes69. Modern large transport aircraft incorporate an auto-brato pre-select various deceleration rates. The maximum auto-braavailable from manual braking. Anti-skid protection is maintainesystem is armed by selecting a deceleration rate, but it will onlthrust levers are at IDLE. The auto-brake system then maintaiconjunction with the aerodynamic speed brakes and thrust rebraking to a complete stop or until disarmed.







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to-brakes disarm automatically andication by the pilot will also disarm

selection, which can only be armedke system applies maximum brake ifft speed (typically 80 to 90 knots).

kes when they are hot and is due tore designed to resist brake fade, but

ake-off.

ed by a variety of factors includinge hydraulic system.

friction the brake friction varies to come from the segmented brakes emit a squealing noise. The causesal on the discs, leading to uneven



70. If an anti-skid or auto-brake system fault develops, the aua warning light is displayed on the instrument panel. Brake applthe auto brake system.

71. The auto-brake system includes a rejected take-off (RTO)with the aircraft on the ground. With RTO selected, the auto-brathe engine thrust levers are retarded to idle above a certain aircra

Brake Symptoms

Fading. Loss of braking action. This occurs in drum type braexpansion of the drum away from the brake shoes. Disc brakes athis may occur with a fully worn brake during a severe rejected t

Dragging. Failure of the brakes to release completely. Causweak or broken return springs, distorted discs and air in the brak

Chattering or Squealing. Instead of maintaining an evenduring one revolution of the wheel. This causes a ‘chatter’ sounddiscs. If the frequency of this chattering is high enough the brakeare warped or glazed discs or deposits of brake lining materifriction.








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energy into heat energy. The greaterted in the brakes. The major singlerakes may cause the disc to warp orircraft with very hot brakes shouldat pack.

re, which can build up into a wedgegating the effects of braking. It candent upon aircraft speed and tyrehe minimum speed for initiation of

eed

d off of the runway and completely

es a very thin film of water which at, or persist down to, much lowerng is particularly associated withch is smoothed by rubber deposits.

n lb in2⁄

9 200×( )is 127 knots



Overheating. The function of the brakes is to convert kineticthe kinetic energy to be converted, the greater the heat generacause of brake overheating is high taxiing speeds. Overheating bcause the friction material to break up and adhere to the disc. Anot be parked with the brakes applied to prevent fusing of the he

Aquaplaning72. Aquaplaning is caused by a layer of water beneath the tyand lift the tyre away from contact with the runway, thereby neoccur in water depths as little as 0.1 of an inch and is depenpressure. A simple formula has been derived which states that taquaplaning is approximately:

Thus, for a tyre pressure of 200 psi the aquaplaning sp

73. There are three distinct types of aquaplaning:

Dynamic. This is due to standing water where the tyre is liftesupported by the water.

Viscous. This occurs when the runway is damp and providcannot be penetrated by the tyre. Viscous aquaplaning can occurspeeds than simple dynamic aquaplaning. Viscous aquaplanismooth surfaces such as the touch-down zone of the runway whi

9 the square root of the tyre pressure i×








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s the affected tyre(s) become tackyly the consequence of a long skide tyre and the wet surface boils the which delays water dispersal. Thethe runway surface.



Reverted rubber. When reverted rubber aquaplaning occurand take on the appearance of uncured rubber. It is normaloccurring on a wet runway, during which the friction between thwater and reverts the rubber. As a consequence a seal is formedresulting steam then prevents the tyre from making contact with









Hydraulics

Basic Principles

Hydraulic Fluids

Basic Aircraft Hydraulic Systems

Hydraulic Components

Hydraulic System Indications

Hydraulic Systems







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greater or lesser degree. The use ofral advantages when compared withitted is very large in comparison toan easily be routed into inaccessible pose engineering problems in terms sufficiently strong anchor points in

fluid is directly proportional to therea of the fluid container alters thish 3 metres tall and filled with fluid. 2 cm in diameter and the third anse. The fluid pressure at the base of

Hydraulics


3Hydraulics

Basic Principles1. Most modern aircraft incorporate hydraulic systems to afluid (hydraulic) pressure to provide power transmission has sevea mechanical system. With a hydraulic system the power transmthe size and weight of the equipment required. Hydraulic lines cparts of the aircraft where the use of mechanical linkages wouldof 90° turns in the transmission train, and the establishment ofrestricted and/or non-reinforced areas.

Hydrostatics2. The pressure produced at the base of a static column of height of the column. Neither the shape nor the cross-sectional afundamental principle of hydrostatics. Imagine 3 containers, eacOne is cylindrical and 3 metres in diameter, the second is a pipeinverted cone 3 metres in diameter at the top and 2 cm at the baeach would be identical.







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ssionlied to a confined fluid the resulting

the fluid is not compressed by the and every direction is undiminished

s not in motion. The situation is

FIGUFluid M

re in the hydraulic system shown ate of 1000 gram is acting on a piston

Hydraulics


Principles of Hydraulic Power Transmi3. Thanks to Mr. Blaise Pascal we know that if a force is apppressure is transmitted equally in all directions. Providing thataction of the applied force, then the pressure transmitted in eachby distance.

4. Pascal's law applies only to a confined fluid which iillustrated at Figure 3-1.

RE 3-1otion

5. Since pressure is equal to force divided by area the pressuFigure 3-1 is 100 gm per sq cm throughout, since the applied forcof 10 square centimetres.







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square centimetres, and since, byng on this piston is 100 kg.

ge, in this case of 100:1, since an

t nothing moves. Obviously we arent movement, and now we discover

ant to displace a 100 kg mass byhe piston at A by 10 cm in order to

Hydraulics


6. Since the piston at the other end of the system is 1000transposition, force is equal to pressure times area, the force acti

7. What we have in fact achieved is a mechanical advantaapplied force of 1 kg has achieved a resultant force of 100 kg.

8. Strictly speaking Pascal's law only applies providing thainterested in a system whereby an applied force causes a resultathat we have not in fact achieved something for nothing.

9. Reproducing Figure 3-1, but now assuming that we wapplying a 1 kg force, we will find that it is necessary to move tmove the piston at B by 1 mm. This is shown at Figure 3-1.







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FIGUMovingAdvan

k achieved at B (assuming that noming friction in the system).

done at A (1 kg x 10 cm = 100,00000,000 gm-mm).

other) hydraulic system. By movinglarge load through a small distance.

Hydraulics


RE 3-2 Fluid with

tage

10. At Figure 3-2 the work applied at A will equal the worwork has been done either in compressing the fluid, or in overco

11. Now, since work is equal to force times distance, the workgm-mm), must equal the work achieved at B (100 kg x 1 mm = 1

12. This is precisely the principle behind an airborne (or anythe small applied force through a large distance you can move a






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system is work wasted. One of theould be effectively incompressible

r waxing.

the elastic material used in the sealsnents.

o aid in the detection of leaks.

Hydraulics


Hydraulic Fluids13. Any work which is done in compressing the fluid in theessential qualities of a hydraulic fluid is therefore that it shthroughout its entire range of approved operating pressures.

14. Other required properties of the fluid are listed below.

(a) Low viscosity

(b) Good lubrication properties.

(c) Non flammable.

(d) Non toxic.

(e) Low freezing point.

(f) High boiling point.

(g) No foaming.

(h) Stable. The fluid should show no decomposition o

(i) Compatibility. The fluid must be compatible with and with the metal in the hydraulic system compo

(j) Coloured. This is for easy identification and also t






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ay be used in an aircraft hydraulic in small aircraft. It is red in colourIt has good lubricating properties, isorrosion. There is little change ofased hydraulic fluid has seriouse leak, there is a serious fire hazard.bstance must not be used with it.l based fluid is designated DTD 585

fluids having a large temperatureydrol 500B which is light purple in and have an operating range ofydrol is highly fire resistant and theilicone, fluorocarbon or butyl. Theric contaminants and thus a systemen forms a sludge. Additionally, anyg made of polyvinylchloride (PVC).and must therefore be handled with

Hydraulics


Types of Fluid15. Only fluid of the type specified by the manufacturer msystem. Mineral or petroleum based oil is most frequently usedand should be used with synthetic rubber, leather or metal seals. chemically stable and has additives to prevent foaming and cviscosity with change of temperature. However, mineral bflammability limitations such that, in the event of a high pressurMineral oil is corrosive to natural rubber and seals of this suSynthetic rubber seals are used with mineral based fluid. Mineraor MIL-H-5606.

Synthetic Based Fluids16. In high performance aircraft there is a requirement forrange. This is provided by SKYDROL. The common grade is Skcolour. Other grades of Skydrol are coloured green or bluetemperatures varying from about -55°C to more than +105°C. Skseals used in association with it are made of materials such as sdrawback of Skydrol is its susceptibility to water and atmospheusing it must be sealed. If overheated, skydrol turns acidic and thfluid leak will attack the insulation of electrical wiring or anythinSkydrol is also very harmful to the skin and especially the eyes, great care.






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o leakage. Moving parts within theepage of fluid from one side of theommon of which are illustrated at

friction free, so as to prevent any

FIGUHydra

Hydraulics


Hydraulic Seals17. To be of any use a hydraulic system must suffer little or nsystem, such as pistons, are fitted with seals, which prevent sepiston to the other. Various shapes of seal are used, the more cFigure 3-3. It is important that, as far as possible, the seal besignificant decrease of efficiency within the system.

RE 3-3ulic Seals






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rgy for activating a wide range ofakes, flaps and flying controls, andn a pump to provide the pressure, arate the various components listed

led.

ydraulic system without a pressure.

FIGUPassiveSystem

Hydraulics


Basic Aircraft Hydraulic Systems18. An aircraft hydraulic system is a source of pressure eneaircraft systems and components such as landing gear, wheel brnosewheel steering. All hydraulic systems must, therefore, contaireservoir of fluid for the pump to draw upon, actuators to opeabove and selectors with which the desired functions are control

19. The hydraulic system shown in Figure 3-4 is a passive hpump. The system is not pressurised unless the pump is operated

RE 3-4 Hydraulic






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hows a basic hydraulic system, withe pump when the actuators are notis unloaded, and a hand pump for

ces. Pumps may be engine driven,lic pipe-lines are normally stainlessturn lines.

Hydraulics


20. As an introduction to the following sections, Figure 3-5 sthe added refinements of an automatic cut-out valve to unload thbeing used, an accumulator to store pressure whilst the pump emergency operation. This is an active hydraulic system.

21. Hydraulic pumps may be powered from various sourelectric motor driven or powered by an air motor. The hydrausteel for pressure lines and may be light alloy for low pressure re






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FIGUActiveSystem

draulic system are described in thetions that they serve.

Hydraulics


RE 3-5 Hydraulic

Hydraulic Components22. The major components which are to be found in a hyfollowing pages, together with an outline of the function or func






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draulic fluid within a system. Thelow for slight losses through minor

Hydraulics


The Reservoir23. The reservoir is effectively a storage vessel for the hyreservoir will also contain an additional quantity of fluid to alleakages. A reservoir is illustrated at Figure 3-6.






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FIGUHydraReserv

Hydraulics


RE 3-6ulic oir






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ill hold the excess volume of fluidg gear is raised or lowered, varyingeir jacks. The reservoir will contain

es may become trapped in it. Theised and any air in the fluid will be a de-aerator.

nt foreign matter from entering theluid level rises or falls (assuming anluid. A sight glass enables the fluidally etched onto the sight glass. Theverfilling the system should also be

, or possible rupture of the system ifa high fluid volume return from a

fluid stored in the reservoir wouldthe system forcing all of the fluidlevel of fluid drops to the top of therve of fluid is still available howevern the form of a hand pump on small

Hydraulics


24. Under conditions of thermal expansion the reservoir wwhich results. When units are actuated, for example the landinvolumes of liquid are required as the piston rods move inside ththe surplus fluid, or alternatively supply an increased demand.

25. During the passage of fluid round a system air bubblreturning fluid is directed in such a way that foaming is minimswirled out or extracted. The device which does this is known as

26. The reservoir normally contains a screened filter to prevesystem, a vent to allow air to leave or enter the tank when the funpressurised system), and baffles to prevent splashing of the flevel to be checked. Minimum and maximum level lines are normresults of allowing the system to become depleted are obvious. Oavoided, since this could result in an overflow of hydraulic fluidthe vent could not cope with the fluid surge in the event of retracting piston.

27. In the event of a major leak in the system, the level of diminish. In order to prevent the power driven pump within through the leak, a standpipe is fitted in the reservoir. Once the standpipe the supply of fluid to the powered pump ceases. A resefor use with the emergency back-up system, which is frequently iaircraft.






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ric pressure is correspondingly low,psi depending on the manufacturer.compressor and the desired pressurer ensures that the system receives aitation occurs (typically on the inlet

is so low that cavities form due totside air pressure falls to a low level.

of hydraulic fluid at the designedvery part of the system.

f the pump is maintained in one of

a constant displacement pump, is pump will move a given amount ofe of fluid needs to be moved at aressure has been attained a cut-out,r increase in pressure. A hydraulicnt displacement pump.

Hydraulics


28. For aircraft that fly at high altitude where the atmosphethe reservoir is pressurised, typically to between 10 psi and 30 Pressurisation is provided by a filtered air bleed from the engine is controlled by a pressure relief valve. Pressurising the reservoiconstant supply of fluid and that pump cavitation is avoided. Cavor suction side of a hydraulic pump) when the fluid pressure entrapped gas expansion. Foaming can be prevalent when the ouAfter entering the reservoir, the fluid will be de-aerated.

The Power-Driven Pump29. The powered pump in the system ensures that a supplysystem pressure is always available when required, at each and e

30. The required system pressure at all points downstream otwo ways.

31. In the constant delivery system the pump, known as operating continuously and is driven by the engine. This type offluid for each revolution and is used when a fairly large volumrelatively low pressure. It means, though, that once the required por pressure relief valve, will be required to prevent any furtheaccumulator will normally be found in any system using a consta






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pFIGUVane T

by a spacer, are free floating in thethe inlet side of the pump will be the rotor, the volume between theh the outlet port. This type of pump

Hydraulics


Vane Type Constant Displacement PumRE 3-7ype Pump

32. The vanes, which are held against the wall of the sleeve rotor. As the rotor rotates, the volume between the vanes on increasing, so that fluid will be drawn in. On the outlet side ofvanes will be decreasing and thus fluid will be forced out througis illustrated at Figure 3-7.






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me of fluid needs to be moved atriven by the engine via the enginer within the pump. As the two gearslet to the outlet side of the pump.

Hydraulics


Spur Gear Displacement Pump33. This type of pump will be used when a medium volupressures up to approximately 1500 psi. One of the gears is daccessory gear box and this gear will, in turn, drive the other gearotate, the space between the gear teeth carries the oil from the in



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FIGUSpur G

Hydraulics


RE 3-8ear Pump



Chapte

amount of fluid is allowed to leake low pressure relief valve opens at being drawn into the pump in the

ssure builds up. On some pumps,a cavity behind the bushing flanges minimise leakage by decreasing sider pump is illustrated at Figure 3-8.

ceases to operate once the requirede pump again delivers fluid into the no cut-out valve or accumulator for prevent over-pressurisation of the

required line pressure is achieved.

displacement pump. In the variableried to maintain constant discharge

uch systems. This type of pump is

Hydraulics


34. For purposes of cooling, sealing and lubrication, a smallpast the gears. This oil is kept within the hollow shafts until thapproximately 15 psi. This so-called case pressure prevents airevent of shaft or seal wear.

35. There is also a tendency for the case to distort as pretherefore, high pressure oil is fed past the check valve and into which are thus forced hard against the side of the gears. This willclearance and will also compensate for bushing wear. A spur gea

Variable Displacement Pumps36. In a constant pressure live line system the pump physicallyline pressure is achieved. As and when the line pressure falls thsystem until the pressure is restored. This type of system requirespressure control, however a pump by-pass valve is essential, tosystem in the event that the pump fails to cease delivery once the

37. The constant pressure live line system employs a variabledisplacement system, pump output volume, or displacement, is vapressure. An axial piston swash plate pump is often used in sillustrated at Figure 3-9.



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FIGUVariabDisplaPump

Hydraulics


RE 3-9le cement



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g contact with the stationary swashrings acting upon the pistons. Since pistons will move backwards and engine-driven shaft.

sh plate angle, and this in turn isd control piston. When the outlete maximum volume of fluid will bering pressure, reducing the angle ofn stroke. A point will eventually beoccurs the pump is said to be idling.

ulic pressure is supplied to one sideline to reservoir, the pressure acting hydraulic inlet pressure, will begled swash plate. Alteration of theply pressure) and reversal of swash

ing a small amount of oil to escapeence back to the reservoir. This isn be routed via a heat exchanger in the aircraft tank is used to cool the

Hydraulics


38. The pistons in the rotating cylinder block are held in slidinplate through shoes, pivoted to the ends of the piston rods, by spthe shoes slide around the angled, stationary, swash plate, theforwards in their cylinders as the cylinder block is rotated by the

39. The extent of the piston stroke will depend upon swaadjusted by pump outlet pressure acting upon a spring-loadepressure is zero or low, the stroke will be at its greatest and thdisplaced. As pressure increases, the control piston overcomes spthe swash plate and progressively shortening the length of pistoreached where the pistons will not be pumping at all. When this

40. The same device is often used as a rotary motor. If hydraof the ported plate and the other side is connected to the return upon the ‘retarding’ pistons, that is the pistons extended byconverted into rotary motion by the shoes sliding around the answash plate angle will alter the speed of rotation (for a given supplate angle will reverse the direction of rotation.

41. Lubrication and cooling of the pump is achieved by allowfrom the pistons into the central chamber of the pump and thgenerally referred to as the case drain cooling flow. The fluid cathe aircraft fuel tank before returning to the reservoir. The fuel inhydraulic fluid.



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r in a small light aircraft hydraulic

running engine.

so on, can be operated.

s non-return valves and also a relief pump or the power unit driving theering the gear and flaps, as well as

larly, if a major leak occurs, and the reservoir, the hand pump will be

reservoir and fed to the essentialcuits. A hand pump is illustrated at

Hydraulics


The Hand Pump42. While the hand pump may be the only source of powesystem, in larger aircraft it will be installed:

(a) As an emergency standby unit.

(b) To allow ground servicing without the need for a

(c) So that pressure joints and lines can be tested.

(d) So that cargo doors, undercarriage bay doors, and

43. The hand pump is normally double acting. It incorporatevalve. In the air, should a failure occur of either the power drivenpump, the hand pump will provide an alternative method of lowproviding vigorous exercise for the most junior of the pilots. Simisystem depletes to the level of the top of the standpipe in therequired. In this case fluid is drawn from the bottom of thehydraulic services via the hand pump and duplicate hydraulic cirFigure 3-10.



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FIGUDoublHand P

ically on large modern aircraft, it isssociated services. This will provideirements of the JAR and FAA.

include:

wn under the influence of gravity,ve 'g'.

Hydraulics


RE 3-10e Acting ump

Additional Safety Features44. Because of the large number of services operated hydraulnecessary to duplicate, or even triplicate pumps and some of the aredundancy in the event of system failure whilst meeting the requ

45. Other means of emergency operation of essential services

(e) Landing gear designed for free fall and lockdopossibly supplemented by the application of positi



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t a hand pump to supplement them.

a separate selector valves to selected

mp.

an the actuators are housed in one

em:

ation of a system component causes pump control responds.

large fluid demands are made, byump.

ic pump is disconnected from the

ited operation of hydraulic services example, flap operation for ground the engines are not running.

Hydraulics


(f) Electrically driven back-up pumps, with or withou

(g) Air, nitrogen or CO2 bottles providing pressure viservices.

(h) A drop out ram air turbine driving a hydraulic pu

(i) Power pack units, where all components other thunit.

Accumulators46. The accumulator serves four functions in a hydraulic syst

(a) It absorbs pressure surges, which occur when opera pressure drop, followed by a pressure rise as the

(b) It provides supplemental system pressure when supplementing the fluid flow from the hydraulic p

(c) It maintains system pressure when the hydraulsystem by the Automatic Cut-Out Valve.

(d) It serves as a pressure storage unit to permit limwhen the pump is not operating. This allows, forservicing or brake operation during towing, when



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phragm Type (the bladder type isd at Figure 3-11.

de of the diaphragm, or piston, and hydraulic fluid can flow freely intoer like a spring to absorb pressure

emaining three functions are met by forcing hydraulic fluid out of the

ervicing, typically to a pressure of the hydraulic pumps pressurize theepressing the diaphragm, or piston, hydraulic system is operating theame value.

sure flexes the diaphragm, or moves freedom of movement of the pistonon of the lower (gas) sealing ring.

e gas side for use during charging.hich reads system hydraulic fluid

umulator to return to the reservoir,

Hydraulics


47. The most common types of accumulator are the Diaessentially similar), and the Piston Type. Both types are illustrate

48. Hydraulic fluid at system pressure is connected to one sithe other side is charged with nitrogen gas under pressure. Systemor out of the accumulator. Since gas is compressible it acts rathsurges, the first of the accumulator functions listed above. The rvirtue of the gas pressure acting on the diaphragm or piston,accumulator as required.

49. The accumulator is charged with gas during ground sapproximately half system pressure. When the engines are startedsystem and fluid enters the fluid chamber of the accumulator, dand compressing the gas in the gas chamber. Thus, when theaccumulator gas pressure and system hydraulic pressure are the s

50. When the hydraulic pumps are not operating the gas presthe piston, to displace fluid into the system as required. To ensurein the piston-type accumulator a drilling is provided for lubricati

51. Accumulators usually incorporate a pressure gauge on thSystem hydraulic pressure is displayed on a pressure gauge wpressure. This is illustrated at Figure 3-12.

52. A bypass valve may be provided to allow fluid in the accthus permitting the accumulator gas charge to be checked.



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FIGUDiaphrAccum

Hydraulics


RE 3-11agm Type ulator



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FIGUAccumPressu

stant delivery system to divert fluidFigure 3-13 and Figure 3-14 show atem, including the accumulator. At see then that an automatic cut-out

Hydraulics


RE 3-12ulator re

Cut-Out Valves53. As already described, the cut-out valve is used in the conback to the reservoir once the required line pressure is achieved. cut-out valve. At Figure 3-13 the pump is pressurising the sysFigure 3-14 the accumulator is pressurising the system. We canvalve of the type shown will:



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FIGUPump the SyAccum

Hydraulics


RE 3-13Pressurising stem and ulator



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FIGUAccumPressuSystem

tem when pressure falls below a set

mulator pressure control.

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RE 3-14ulator rising the

(a) Provide a return line to the hydraulic reservoir.

(b) Close the return line and direct fluid into the sysfigure.

(c) Give a smooth transition between pump and accu



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ermit an idling circuit from pump tomption of engine power and, at theuid.

d pressure is required to operate it.ydraulic pressure).

and seal off the system to maintain

valve provides reliable informationer internal or external, will cause a

re will be insufficient fluid in thehis inadequate quantity of ‘stored’ice operations, causing the cut-outnly to be quickly exhausted again assulting in a rapid on/off cycle.

fluctuations in pressure since theure surges. The accumulator pistontic cut-out valve frequently leads to

Hydraulics


(d) Automatically cut out to off-load the pump and preservoir. This will reduce pump wear and consusame time, prevent overheating of the hydraulic fl

(e) Cut in automatically when a service is selected an(Cut-in will also occur if a leak causes a drop in h

(f) Act as a non-return valve during cut-out periods pressure when the hydraulic pump is idling.

54. The interval between cut-out and cut-in of the automaticconcerning the condition of the hydraulic system. A leak, whethreduction in the period between cut-out and cut-in.

55. If the accumulator gas charge pressure is too high theaccumulator when system and gas pressures have equalized. Tfluid under pressure will be rapidly exhausted after a few servvalve to cut the pump in. Pressure will be immediately restored, othe fluid contents of the accumulator are once again depleted, re

56. Too low an accumulator pressure will cause rapid accumulator will no longer serve its function of absorbing presscontacts an internal stop and rapid on/off cycling of the automa‘hammering’, a loud knocking noise, in the system.



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irection only. A non return valve issociated with the cut-out valve.

FIGUSimpleRelief

Hydraulics


Non-Return Valves57. These valves permit the flow of hydraulic fluid in one dshown at Figure 3-13 and Figure 3-14 as part of the plumbing as

Pressure Relief ValvesRE 3-15 Pressure Valve



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used by excess pressure, a pressureas a safety device and is designed totly higher than the intended systemainst its seat by a spring. The spring

to the return line to the reservoir. A

ich a pressure relief valve begins to designed to operate at 3000 psi, the0 psi and to reset at 3190 psi. Thelues of the fully open and the fully

Hydraulics


58. In order to prevent damage to the hydraulic system carelief valve is invariably incorporated into the system. It is used open when system pressure reaches a preset value which is slighpressure. The valve comprises a simple ball valve which is held agtension is adjustable and is set to relieve a small amount of fluidsimple pressure relief valve is shown at Figure 3-15.

Cracking Pressure59. Cracking pressure is the term given to the pressure at whopen. For example, in a typical aircraft hydraulic system which ispressure relief valve might be designed to be fully open at 365cracking pressure will therefore be somewhere between these vareseated pressures.



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FIGUFull FloValve

er-pressurisation due to a variableump stuck on maximum flow. Thisp output while limiting the pressure

full flow relief valve is shown at

Hydraulics


Full Flow Relief ValvesRE 3-16w Relief

60. A full flow relief valve protects the system against ovdisplacement pump pressure control failure which leaves the ptype of relief valve must be capable of passing the maximum pumto approximately 10% above normal working pressure. A Figure 3-16.



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ves are designed to open when the increase in pressure is caused by anid in a part of the system where theating jack. Thermal relief valves arel pressure relief valves in the samethe normal operation of the system.

ump, shut-off valves are normallye used to facilitate ground servicing

ulic fluid whilst connecting another. pump would become useless as thee. It would therefore be necessary to

p. The necessary re-routing of the shuttle valve becoming stuck in itsapable of connecting the emergency

Hydraulics


Thermal Relief Valves61. Just like normal pressure relief valves, thermal relief valpre-set pressure at the valve is exceeded. In this case however theincrease in the temperature and subsequent expansion of the flufluid is trapped, typically between a non return valve and an operdesigned to operate at a higher pressure than the conventionasystem. The thermal relief valves do not therefore interfere with

Shut-Off Valves62. Positioned between the hydraulic reservoir and the poperated electrically to cut off fluid supply to the pump. They arand to isolate the fluid supply in the event of engine fire.

Shuttle Valves63. Shuttle valves are used to disconnect one source of hydraAs already described, following a leak in the system the poweredlevel of fluid in the reservoir dropped to the level of the standpipsupply fluid from the bottom of the reservoir via the hand pumcircuit would be achieved using shuttle valves. In the event of a‘normal’ position (as illustrated at Figure 3-17) it would be incsupply to the system.



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FIGUShuttle

quired for a particular service, for

Hydraulics


RE 3-17 Valve

Pressure Reducing Valves64. These will reduce the main system pressure to that reexample the brake system.



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FIGUPressuValve

operating condition. The spring isigh pressure (HP) supply will, byessure. When main high pressure is compresses, allowing the piston to

where the pressure of fluid enteringsure (LP) output will therefore beo uncover the return port to route

Hydraulics


RE 3-18re Reducing

65. Figure 3-18 shows a pressure reducing valve in a non designed to resist the required service pressure. The main hdefinition, deliver a greater pressure than the required service prsupplied to the valve the piston is forced against the spring whichmove. This movement closes off the HP supply port to a positionthe valve is balanced by the spring pressure and the low presrequired system pressure. The movement of the piston will alsexcess fluid back to the reservoir.



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nstream of the fuse. They are ofteny operate when pressure drop acrosslic fuse in the normal and closed

FIGUHydra

some essential services by isolatingreset value.

Hydraulics


Hydraulic Fuses66. Hydraulic fuses prevent fluid loss when a leak occurs dowincorporated in braking, flap and thrust reverser systems and thethe fuse exceeds a preset value. Figure 3-19 shows a hydraupositions.

RE 3-19ulic Fuse

Pressure Maintaining Valves67. These are designed to maintain a desired pressure for supply to non-essential services if system pressure falls below a p



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FIGUPressuMainta

rmal operating configuration. Main spring which is designed to balancelied to both primary and secondaryr example by component failure, theollar onto the seating, cutting off theual pressure to be channelled to the

Hydraulics


RE 3-20re ining Valve

68. Figure 3-20 shows a pressure maintaining valve in the nosystem supply pressure is acting on the piston and depressing thethe main supply pressure. There will therefore be pressure suppservices. In the event that the main system pressure is reduced, fospring will overcome this reduced pressure and force the piston cpressure supply to the secondary services. This will allow all residprimary services, for example the flying controls.



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f a pressure maintaining valve. Thessume that the system fluid content

ed in the reservoir will be activated.ses off the supply line to the non-

is shown at Figure 3-21, where theercarriages and gear doors can bepply sufficient energy to stop theonsequently these are considered to

Hydraulics


Priority Valves69. On some aircraft systems a priority valve is used instead oend result is the same but is achieved by different means. Let us ais falling due to a leak. At a pre-determined level, a switch locatThis will transmit a signal to the priority valve which then cloessential services. An example of a system using a priority valvepowered controls are necessarily the essential services. The undlowered under free-fall, the wheel brake accumulators will suaircraft on the ground and the aircraft can land without flaps. Cbe non-essential services.



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FIGUPriorit

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RE 3-21y Valve



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r valves are needed to control fluidoses a line as required. More usuallyh cases it is necessary to incorporatector valves may be of rotary, poppetat Figure 3-22.

Hydraulics


Selector Valves70. In a complex hydraulic system, control valves or selectoflow. The simplest type of all is an on/off valve which opens or clhydraulic jacks are required to operate in both directions. In suca four-way valve which permits fluid flow in either direction. Seleor piston type. Rotary and piston type selector valves are shown



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FIGUSelecto

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RE 3-22r Valve



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vesFIGUOrificeValve oRestric

hydraulic fluid in one direction andrictor valve. Flow in the direction Astricted flow in that direction. Flowe the valve onto the seating, therebye of the most common applicationsndercarriage is heavy it will tend toting its movement is used. Since thetion, any restriction in this line willulically operated flap systems a one

f the flaps during flap extension to

he flaps during flap retraction when

Hydraulics


Orifice Check Valves, or Restrictor ValRE 3-23 Check r tor Valve

71. An orifice check valve is designed to provide free flow ofrestricted flow in the opposite direction. Figure 3-23 shows a restto B compresses the spring, which opens the orifice to allow unrein the direction B to A acts together with the spring force to movclosing the orifice and restricting fluid flow in that direction. Onof this device is in the up line of a landing gear system. Since the ufall too rapidly when being lowered, unless some means of restricup line is the return line for the hydraulic fluid during the operalimit the speed of movement of the gear. Similarly in some hydraway restrictor is fitted in both the flap up and flap down lines.

The restrictor in the flap up line controls the lowering speed oreduce the effects of aircraft trim change.

The restrictor in the flap down line controls the raising speed of tthe slipstream tends to blow the flaps up.



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ictor, serves the same purpose as ann one direction and full flow in the

draulic operations into the requiredin the landing gear system. The

elf is extended, and these operationsce valves will be seen in the section

lves. They allow system pressure tod:

he accumulator, and

omponent.

two components, thus avoiding ather.

Hydraulics


Metering Check Valves72. A metering check valve, sometimes called a one-way restrorifice check valve, but is adjustable to permit a metered flow iopposite direction.

Sequence Valves73. A sequence valve (or timing valve) organises a series of hysequence. A common example of the use of this valve is undercarriage doors must be opened before the undercarriage itsare properly sequenced by this type of valve. Examples of sequendealing with landing gears.

Manual Pressure Relief Valves74. Alternatively known as off loading valves or dumping vabe released back to the hydraulic reservoir and would be operate

(a) to exhaust system pressure prior to gas-charging t

(b) to permit servicing or exchanging of a hydraulic c

Flow Dividers75. Their function is to divide equally the flow between situation where there is pressure in one line but cavitation in ano



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tion is shown at Figure 3-24. Theses:

his is achieved by the springs withinnges.

urst pressure gauge. Under normalhe piston and compress the piston the open position. If the pressuret failure the higher system pressure spring pressure and seat the valve,eed line.

auge. Disconnecting the gauge willsure relay fitted in the line the gauget the risk of loss of hydraulic system

Hydraulics


Pressure Relay Valves76. A pressure relay valve in the normal operating configuravalves are fitted into pressure gauge lines and have three function

(a) To protect the gauge from pressure fluctuations. Tthe valve, which dampen out random pressure cha

(b) To prevent hydraulic fluid loss in the event of a boperating conditions back pressure will act on tsprings. The valve spring will retain the valve indownstream of the valve drops due to componenwill move the piston in order to overcome valvethereby closing off the flow to the burst gauge or f

(c) To allow servicing/replacement of the pressure ghave the same effect as a burst gauge. With a presmay be removed for servicing/replacement withoucontents.



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FIGUPressuValve

h axle mounted anti-skid units. Onto match the capacity of the brake,essure. The reduced flow conservesxhaust operating pressure when the

produced by hydraulic pressure intocks are commonly used as described

Hydraulics


RE 3-24re Relay

Modulator Valves77. These will be found in brake systems in association witinitial brake application, an unrestricted fluid flow is provided but then fluid flow will be modulated to only allow reduced prmain system pressure and allows the brake unit to completely eanti-skid unit comes into operation.

Hydraulic Jacks78. The purpose of any hydraulic jack is to convert the force a linear movement of a piston rod, or ram. Two basic types of jabelow:



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ressure in one direction, and in thedraulic force, such as a spring. Ading gear and door locks.

either direction is due to hydraulicalve. The jack may be compensatede area on either side of the piston isth sides of the piston. In the secondmote from the rod or ram is greater

is normally used for landing gearin raising the gear or extending theps.

-compensated) jacks are shown at

Hydraulics


(a) Single acting jack. These move under hydraulic pother direction under the influence of a non hycommon application of the single acting jack is lan

(b) Double acting jack. In this case the movement inpressure, and is controlled by means of a selector v(balanced) or non-compensated. In the first case thidentical, since there is a piston rod or ram on bocase the area on the side of the piston which is rethan on the other side. A non-compensated systemand flap systems, where a greater force is needed flaps than in lowering the gear or retracting the fla

79. Single-acting and double-acting (compensated and nonFigure 3-25.



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FIGUTypes Jacks

Hydraulics


RE 3-25of Actuator



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ich are automatically engaged onceautomatically disengaged before thecarriage, which is well protected byilot, otherwise taxiing can prove a

Hydraulics


Mechanical Locks80. Locking devices are mechanical latches, braces or pins whthe required hydraulic function has been achieved. They must be process can be reversed. An obvious example of this is the underlocking devices to prevent its retraction until required by the plittle difficult.



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FIGUPressu

accumulate in a hydraulic system prevent this, filters are installed in

Hydraulics


FiltersRE 3-26re Filter

81. The tiny particles of metal, dust and seal material thatwould cause significant damage if allowed to circulate freely. Tothe pressure and return lines.



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sually made of cellulose material,e bowl and must pass through the. A typical return filter is shown at

ass valve. As the element becomesn this differential reaches a pre-set direct to the hydraulic system.

temperature (0°C) to prevent false

ed protruding pin, or button, aboutr element requires changing befores are non-bypass. This is to ensurealves.

nel on the flight deck, containingoss of hydraulic pressure, excess activate warning lights or captions.gined jet aircraft.

Hydraulics


82. A hydraulic filter consists of a renewable ‘element’, usituated in a bowl-shaped container. Hydraulic fluid enters thcylindrical filter element in order to reach the hydraulic systemFigure 3-26.

83. Return filters normally incorporate a spring-loaded by-pclogged the pressure differential across the filter increases. Whevalue, the by-pass valve opens and allows unfiltered fluid to pass

84. The clog indicator pin on some filters is restrained at lowwarnings due to viscous fluid.

85. In some cases this actuates an indicator, which extends a r5 mm (3/16 inch) to alert the operator to the fact that the filteflight. On aircraft with powered flying controls, pressure filterthat no debris is permitted to pass which could jam the control v

Hydraulic System Indications86. Larger aircraft normally have a hydraulic services paindications of fluid quantity, pressure and temperature. Ltemperature and loss of fluid flow (indicating pump failure) mayFigure 3-27 shows a typical hydraulic control panel for a twin-en



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FIGUTypicaContro

re as follows:

s of fluid, pump failure or pressure

r low air pressure. Low air pressure

failure or drive shaft failure.

Hydraulics


RE 3-27l Hydraulic l Panel

87. The implications of the hydraulic system warning lights a

(a) System low pressure. This may be caused by losfilter blockage.

(b) Reservoir Low Level. This may be low fluid level omay cause pump cavitation.

(c) Pump Low Pressure. This may be caused by pump



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low position. Blockage of case drain

lying one hydraulic system. Duringd and the other pump is off loaded.nding gear, flaps and slats the otherperate.

ngined executive jet type of aircraft. pressure to an automatic cut-out

0 and 1500 psi. Standby hydraulic pump, for emergency or ground air conditioning system. A system

r passenger aircraft where system (A, B and standby), provide powering gear extension and retraction,

is pressurised by two engine-driven outboard flight spoilers and ground/retraction and nose-wheel braking.

Hydraulics


(d) Pump Overheat. Pump swash plate stuck in high fcooling flow.

88. Some large aircraft have two engine driven pumps suppnormal operation, one pump is capable of supplying the demanDuring periods of high demand i.e. simultaneous operation of lapump may either automatically come on-line, or be selected to o

Hydraulic Systems89. Figure 3-28 shows a hydraulic system as fitted in a twin-eThe engine-driven constant displacement pumps (EDPs) supplyvalve (ACOV), which maintains system pressure between 125pressure can be provided by an electrically-driven auxiliaryoperations. The reservoir is pressurised to 10 psi from the cabinrelief valve is set to begin opening at 1700 psi.

90. Figure 3-29 shows the concept usually found in largeredundancy is important. Three independent hydraulic systems,to operate the aircraft's primary flying controls, spoilers, landwheel braking systems, nose-wheel steering and wing flaps.

91. System operating pressure is typically 3000 psi. System Apumps and supplies hydraulic power for operation of wing flaps,spoilers, ailerons, elevators, lower rudder, landing gear extensionIt is the alternate pressure source for mainwheel braking.



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s and operates ailerons, elevators,oor and aft passenger stairs. System

en pump powered from the essentialting pressure for a standby rudder

sections of the hydraulic system asAn internal leak will cause a fluid

n at Figure 3-29 is merely a typical

and independent hydraulic systems,d by one engine driven pump (EDP)y two electric pumps with, in some

rmal operation all three systems areome spoiler surfaces. In the event ofwer from the right system to retractnsfers energy from one system toump (RAT) may be provided as an

ngine or engine driven pump failure.und by maintenance. This system is

Hydraulics


92. System B is pressurised by two electrically-driven pumpupper rudder, inboard flight spoilers, mainwheel brakes, cargo dB can be interconnected on the ground to pressurise System A.

93. The standby system is pressurised by one electrically-drivAC bus. It is the standby to system A and will supply operaactuator and the wing leading edge flaps.

94. Maintenance isolation valves may be provided to isolatean aid to trouble-shoot internal leaks in the hydraulic system. temperature rise in the system.

95. It must, of course, be appreciated that the system showsystem and that specific aircraft systems will vary in detail.

96. In the Boeing 757, for example, there are three separate left, centre and right. The left and right systems are each powereand one electric pump. The centre hydraulic system is powered baircraft, a ram air turbine (RAT) power pump as back-up. In noin use. Each system powers all the primary flying controls and sleft engine failure on take-off, a power transfer unit supplies pothe landing gear and lift devices. The power transfer unit traanother. There is no fluid transfer. A ram air turbine powered pemergency source of hydraulic power to be used in the event of eOnce deployed in flight, the RAT is normally stowed on the groillustrated schematically at Figure 3-30.



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FIGUHydraTwin E

Hydraulics


RE 3-28ulic System - xecutive Jet



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FIGUHydraTypica(B727)

Hydraulics


RE 3-29ulic System - l System



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FIGUHydraB757

Hydraulics


RE 3-30ulic System -



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tems of light aircraft, where only ally operated, either an open centre

n no components are being used the flow freely from the engine-drivenn to the reservoir. Thus the pump iscks) in a state of hydraulic lock.

ild up to move the actuating jack. and returns the selector valve to its a system is that, with the selectore. An open centre hydraulic system

Hydraulics


Light Aircraft Hydraulic Systems97. In order to minimise the complexity of the hydraulic sysfew components such as flaps and landing gear are hydraulicasystem or a self-contained hydraulic power pack may be used.

Open Centre Systems98. These use selector valves in series with one another. Wheselector valves are in a neutral position and hydraulic fluid canpump, through the open centre of each selector valve, and returoperating on virtually no load, with the component actuators (ja

99. Only when a selection is made does pump pressure buWhen jack travel is completed, pump pressure builds up furtherneutral (open centre) position. The main disadvantage of suchvalves in series, only one actuator can be operated at any one timis illustrated at Figure 3-31.



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FIGUHydraOpen

Hydraulics


RE 3-31ulic System - Centre



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ydraulic system requirements is top, reservoir and selector valves) intower pack as fitted in some Cessna

pack are of the poppet type.

Hydraulics


Hydraulic Power Packs100. The trend in modern light aircraft with only limited hincorporate the whole of the hydraulic power system (hand pumone easily-serviced unit. A schematic diagram of a hydraulic polight aircraft is shown at Figure 3-32. The selector valves in this



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FIGUHydraLight APower

Hydraulics


RE 3-32ulic System - ircraft

Pack



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ding gear.

Hydraulics


Self Assessed Exercise No. 1

QUESTIONSQUESTION 1.

What is the function of the landing gear oleo/shock strut.

QUESTION 2.

What is the function of the bogie unit.

QUESTION 3.

What is the purpose of the drag strut on the undercarriage unit.

QUESTION 4.

What is the function of the torque link.

QUESTION 5.

How is the gear prevented from collapsing on the ground.

QUESTION 6.

What methods are used to provide alternative lowering of the lan



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ystems.

Hydraulics


QUESTION 7.

What is meant by the term ‘shimmy’ relating to a nosewheel.

QUESTION 8.

What is the function of an autoretract brake system.

QUESTION 9.

Describe the operating principle of the anti-skid system.

QUESTION 10.

List the properties of the ideal hydraulic fluid.

QUESTION 11.

Identify the types of hydraulic fluids in use in aircraft hydraulic s

QUESTION 12.

Explain the working principle of passive hydraulic system.

QUESTION 13.

Explain the working principle of an active hydraulic system.

QUESTION 14.

Describe the function of the Hydraulic System reservoir.



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system and the Fluid Cooler.

ircraft.

Hydraulics


QUESTION 15.

Describe the function of the Hydraulic System accumulator.

QUESTION 16.

Describe the working principle of a pneumatic pump.

QUESTION 17.

Describe the working principle of a Hydraulic Pump Case Drain

QUESTION 18.

Describe the working principle of a hydraulic motor.

QUESTION 19.

Describe the function of a check valve.

QUESTION 20.

What is the function of a Priority Valve.

QUESTION 21.

What is the normal hydraulic system pressure of Jet Transport A

QUESTION 22.

Describe the operation of a single acting jack.



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stem.

patterns.

Hydraulics


QUESTION 23.

What indicating instruments are used to monitor the hydraulic sy

QUESTION 24.

Describe the operation of the anti-skid system.

QUESTION 25.

What is meant by tyre creep.

QUESTION 26.

Why are some tyres limited to a maximum ground speed.

QUESTION 27.

What is the function of the brake wear indicator.

QUESTION 28.

Identify a tyre tread and describe the advantages of various tread

QUESTION 29.

What is meant by a tyres ‘ply rating’.

QUESTION 30.

List the stresses acting upon an aircraft in flight.



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to withstand stress is in.

rne by the …… and the torsional

Hydraulics


QUESTION 31.

In flight, the lower surface of a wing is subjected to stress called.

QUESTION 32.

In a monoque type fuselage construction, the strength and ability

QUESTION 33.

In a cantilever wing construction, the bending stresses are bostresses by the _________ ribs.

QUESTION 34.

List the different types of fuselage construction.

QUESTION 35.

What do you understand by damage-tolerant design.

QUESTION 36.

Define the maximum zero-fuel weight.

QUESTION 37.

What limits the airframe life.



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rised aeroplane are subjected.

ressurised aeroplane.

Hydraulics


QUESTION 38.

What are the two major stresses to which the fuselage in a pressu

QUESTION 39.

How is stress concentration avoided in the cabin windows of a p

QUESTION 40.

What type of construction is used in large jet transports.



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ound.

re than one axle. This enables the

ndercarriage, especially on landing.

ircraft.

An alternative to this would be the

it can be allowed to free fall under

Hydraulics


ANSWERSANSWER 1.

To absorb the shock loads of landing and taxiing over uneven gr

ANSWER 2.

A bogie unit consists of a multi – wheels unit design using moaircraft weight to be spread over a greater surface area.

ANSWER 3.

A drag strut will absorb the drag (fore/aft) loads imposed on the u

ANSWER 4.

To maintain wheels alignment with the longitudinal axis of the a

ANSWER 5.

Down locks prevent the gear from collapsing on the ground. provision of a geometric lock associated with the sidestays.

ANSWER 6.

Pneumatic pressure can be used to lower the gear alternately orgravity.



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about the centreline of a naturally

action takes place, the wheels are

ircraft without the wheels skidding

Low freezing point. Non foaming.

ed red and generally used in lightrple and used in high performance

by exerting a force on a small areager force on a piston having a larger

Hydraulics


ANSWER 7.

Nosewheel shimmy is a situation which occurs when oscillationcastering nosewheel becomes excessive.

ANSWER 8.

To reduce the stress on the main wheels spinning before retrautomatically braked.

ANSWER 9.

The anti-skid system will provide maximum retardation of the afor varying runway conditions.

ANSWER 10.

Low viscosity. Good lubrication. Non flammable. Non toxic. Stable. Compatibility.

ANSWER 11.

Mineral based oil, designated DTD 585 (MIL-H5606) colouraircraft. Synthetic oil (phosphate ester based) coloured light puaircraft.

ANSWER 12.

Fluid pressure is transmitted in an enclosed hydraulic fluid systempiston and moving it over a set distance. This will produce a lararea and will move it over a shorter distance.



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Flaps, Flying controls etc. A highm pressure pump and with suitable

ll fluid losses. Allow for thermaluid displacement. Provide a reserve

ure surges. Provided supplementalpressure when pump is idling (in

, is coupled to a hydraulic pump vias available when required.

Hydraulics


ANSWER 13.

In order to operate such aircraft components as Landing Gear,pressure energy source is required. This is provided by the systerouting and control, the various components are actuated.

ANSWER 14.

The reservoir will: Store Hydraulic fluid. Make up for smaexpansion. De-aerate the fluid. Allow for varying volumes of flof fluid.

ANSWER 15.

The accumulator serves the following functions; Absorbs presssystem pressure on high fluid demand. Maintains system association with ACOV). Stores Hydraulic pressure.

ANSWER 16.

An air driven motor, supplied from the aircraft pneumatic systema drive shaft. The hydraulic pump ensures that system pressure i



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ng part of the pump and routed tomally situated in the lower part of a

pressure applied to the pistons will, to convert the resulting force into

aulic fluid flow in one direction and

tial hydraulic services by isolatingre-set value.

Once hydraulic pressure is released

Hydraulics


ANSWER 17.

Hydraulic fluid is designed to flow through the internal workireturn to the reservoir. It will pass through a heat exchanger, norfuel tank, which will cool the hydraulic fluid.

ANSWER 18.

A hydraulic motor is the reverse of a hydraulic pump. Hydrauliccause the shoes, in contact with the variable angle swash platerotary motion.

ANSWER 19.

A check valve (or restrictor valve) is designed to provide full hydra restricted flow in the opposite direction.

ANSWER 20.

A priority Valve will maintain desired system pressure to essensupply to non essential services if system pressure falls below a p

ANSWER 21.

Normal system pressure is 3000psi.

ANSWER 22.

The jack moves under hydraulic pressure in one direction only. the jack is returned by spring operation.



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ntents. In addition, warning lightspresent.

computer and is compared to thel speed diminishes, braking force is

e wheel. It can be identified by the and tyre.

ational speed could limit take-off

nal protrusion, the amount of brake

f the tyre. A ribbed tread is most and directional stability. Patterned

Hydraulics


ANSWER 23.

Gauges are used to indicate system pressure, temperature and cocan be used to alert the operator when abnormal conditions are

ANSWER 24.

Wheel rotation speed is measured by a transducer, passed to abarking demand and aircraft de-acceleration rate. As rotationareduced to prevent wheel locking.

ANSWER 25.

Tyre creep is the rotational movement of the tyre in relation to thchange in position of corresponding marks painted on the wheel

ANSWER 26.

It forms part of a tyres rating. It means that the tyre rotperformance.

ANSWER 27.

A brake wear indicator will indicate, by the amount of pin exterpad wear.

ANSWER 28.

The tread refers to the area forming the crown and shoulder ocommonly used, which provides good traction, long tread wear(diamond) treads are more suitable for unpaved surfaces.



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cessarily reflect the number of plies

and anticipates their growth, which

Hydraulics


ANSWER 29.

Ply rating is an index of the strength of that tyre. It does not neused in the construction.

ANSWER 30.

Bending. Tension. Compression. Torsion. Shear.

ANSWER 31.

Tension.

ANSWER 32.

The fuselage skin.

ANSWER 33.

Spar’s, chordwise

ANSWER 34.

Steel tube (or truss). Monoque. Semi-monoque.

ANSWER 35.

It is assumed that production components will have minor flowsare then monitored by precise inspection procedures.



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e fuel.

.

Hydraulics


ANSWER 36.

The maximum permissible weight of an aeroplane with no usabl

ANSWER 37.

Fatigue due to the load cycles on T/O, landing and pressurisation

ANSWER 38.

Bending and hoop stress

ANSWER 39.

Well-rounded edges

ANSWER 40.

Semi-monocoque





Air Conditioning & Pressurisation

Compressed Air Sources and Uses

Compressor Bleed Air Pneumatic Systems

Air Conditioning

Pressurisation



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urisation

sd air may be used as an emergencydiscussing gas turbines, air startingerate the gyroscopic instruments in

r conditioning system. In the case ofed to pressurise the cabin. In somend a number of aircraft currently in

clusively to operate services. Boths. The most significant difference inompressible whereas air is highly

y engine driven compressors areare used to operate services such as:



4Air Conditioning & Press

Compressed Air Sources and Use1. When discussing braking systems we saw how compressepressure supply in the event of hydraulic system failure. When systems are covered. We know that low pressure air is used to opsome light aircraft.

2. In addition, air is supplied under pressure to the cabin aiaircraft operating above 10,000 feet altitude, this air is also usaircraft compressed air is used to operate ice protection systems aservice have air-operated leading and trailing edge flaps.

3. Some piston engined aircraft use high pressure air expneumatic and hydraulic systems are completely enclosed systemthe two systems being that hydraulic fluid is practically inccompressible.

Uses4. Extensive high-pressure pneumatic systems powered bgenerally fitted on the older types of piston engined aircraft and

(a) Landing Gear



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umatic system pressure. WorkingDepending upon required workingntercoolers are used to cool the airto dissipate heat in the compressor

Figure 4-1.



(b) Wing Flaps

(c) Wheel Brakes

(d) Radiator Shutters

(e) Aerofoil De-icing

Power Sources5. Engine driven compressors provide the required pnepressures range from 450 psi to 3500 psi (typically 3000 psi). pressure, compressors will have several stages of compression. Ibetween each stage of compression. Finned tubes may be used delivery line.

System Description and Components6. A typical high pressure pneumatic system is illustrated at



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FIGU



RE 4-1

The main components are:

(a) Engine driven compressor

(b) Pressure regulator

(c) Main storage bottle

(d) Pressure reducing valve



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nd provides a means of off-loadingregulator valve failure.

ressed air, giving a reserve of power the event of engine or compressor

in system pressure (e.g. brakes) a

om the compressed air passing fromm compressor lubrication and the



(e) Oil and water trap

(f) Indication and warning

Pressure Regulator7. A pressure regulator valve controls the system pressure athe compressor. A relief valve protects the system in the event of

Main Storage Bottle8. The main storage bottle acts as a reservoir to store compfor short bursts of heavy service operation, or emergency use infailure.

Pressure Reducing Valve9. For actuators requiring a pressure lower than the mapressure reducing valve is fitted.

Oil and Water Trap10. The oil and water trap is fitted to remove oil and water frthe compressor to the air storage bottle. The oil originates frowater is precipitated from the air in the compression process.



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tion of pressure in the main and the illumination of a warning light.

Systemsigh by-pass ratio engines, have apressors. Air is bled from the laterigh to satisfy the requirements of allFigure 4-2.

sor and supplied to the system at an be supplemented from the HPpply. The very high temperature ofhrough a pre-cooler heat exchanger,

can be supplied from the auxiliaryused to supplement the pneumaticson.



Indication and Warning11. Pressure gauges, located in the cockpit, provide indicaemergency storage bottles. Failure of a compressor is indicated by

Compressor Bleed Air Pneumatic12. Gas turbine powered aircraft, especially those using hsuperfluity of compressed air produced by the gas turbine comstages of the HP compressor, where the pressure is sufficiently hthe air-driven services. An example of such a system is shown at

13. Air is bled from the 5th stage of the engine HP comprescontrolled rate by the engine bleed air valve. This bleed cacompressor 9th stage when system demand exceeds 5th stage suthe bleed air is reduced to manageable proportions by passing it tcooled by air bled from the turbo-fan outlet (by-pass air).

14. When main engines are not operating, the bleed air systempower unit (APU) compressor bleed. This source can also be system in flight in the event of loss of engine bleed air for any rea



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rating the main engine air turbinen of the hydraulic system reservoirsioning packs. The system is divideduld normally be closed, with the lefte number two engine. Either of thetic requirements of the aircraft andair conditioning and pressurisation.in air back to the mix manifold tod to extract any contaminants from



15. The bleed air (pneumatic) system supplies air for opestarters, wing leading edge thermal anti-icing (TAI), pressurisatioand the domestic water tank, and for the two aircraft air conditinto two halves, separated by an Isolation Valve. In flight this wosystem supplied by the number one engine and the right by thengines, or the APU, is capable of supplying the normal pneumaone air conditioning pack is capable of meeting the demands of The air conditioning load is reduced by recirculating some cabsupplement incoming fresh air. A simple replaceable filter is fittethe recirculation flow before it passes back to the cabin.



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FIGUAir DrService



RE 4-2iven s



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rakes is often provided for by aires in the event of APU failure.

eing lighter than hydraulic systems,here) is virtually inexhaustible. For have employed pneumatic powered stages of flight (flaps, landing gear weight saving advantage.

the same as hydraulic systems, withs, or jacks. The disadvantages ofs of air and the dangers of its waterers are a potential explosive hazard.

sed briefly below:

e of 1lb per person per minute inan half this rate following a failureioning system.

tained within the range +18°C to

ust be maintained at, or close to,the relative humidity is only 1 to



16. Emergency pneumatic operation of landing gear and bstorage bottles. These may also be fitted for air starting the engin

17. Pneumatically powered systems have the advantage of bsince no return lines are required and the reservoir (the atmospthis reason some aircraft (the F27 is the only surviving example)services throughout. In others, those systems used only at certainand brakes) are pneumatically operated, primarily because of the

18. In construction and actuation, these systems are basicallycontrol effected through selector valves and actuation by rampneumatically operated systems are the non-lubricating propertiecontent freezing at altitude. Also, high pressure air storage cylind

Air Conditioning19. The requirements of an air conditioning system are discus

Provision of Fresh Air. Fresh air must be provided at the ratnormal circumstances, or at not less thof any part of the duplicated air condit

Temperature. Cabin air temperature should be main+24°C (65°F to 75°F).

Relative Humidity. The relative humidity of the cabin air m30% (in the atmosphere at 40,000 ft 2%).



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apable of maintaining an adequate temperature and relative humidity.

he operating characteristics of theill vary.

he cabin air must not exceed one

ded on the ground and during

duplicated to the extent that no provision of fresh air to fall to an per minute.

cabin is maintained at a constantisation when cruising at maximum

cabin is controlled by mixing hotmaintain the cabin air temperature

air supply at a controlled rate, toy.



20. An aircraft air conditioning system must therefore be csupply of air for ventilation and pressurisation at a comfortableThese requirements are met as follows:

21. Depending upon the type of power unit fitted, and taircraft, the source of cabin air and the method of conditioning w

Contamination. Carbon monoxide contamination of tpart in twenty thousand.

Ventilation. Adequate ventilation must be proviunpressurised phases of flight.

Duplication. The air conditioning system must be single component failure will cause therate which is lower than 0.5lb per perso

Adequate Supply. The mass flow of air into the aircraft value, sufficient to achieve cabin pressuroperating altitude.

Temperature. The temperature of the air supply to theand cold air in variable proportions to within prescribed limits.

Humidity. Atomised water is added to the cabin maintain a comfortable level of humidit



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atmospheric air is admitted to the as ram air systems. Some of the ramroportions to achieve a comfortable

r through a muffler surrounding theding the combustion chamber of a and combustion heater systems are



Ram Air Systems22. In light, unpressurised, piston-engined aircraft ambient cabin by way of intakes facing into the airflow. These are knownair can be heated and added to the cold ambient air in variable pcabin temperature.

23. Heating is achieved either by passing some of the ram aiengine exhaust system, or by passing it through a duct surrounfuel-burning heater. The operating principles of the exhaust muffshown in Figure 4-3 and Figure 4-4.



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FIGUMuffler

re within the combustion chamber.pplied from the normal fuel systemsolenoid may be effected by duct



RE 4-3 Heating

24. The combustion heater works by burning a fuel/air mixtuAir for combustion is supplied by a blower or fan and fuel is suvia a solenoid operated fuel valve. Control of the fuel valve temperature sensors, or by manual override.



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tially no possibility exists that anyund the chamber. The system mustin the combustion chamber. Safety

air outlet temperature becomes too

tion.

lure of the structural integrity of the



25. The combustion chamber must be so designed that essenleak can occur from within the chamber into the air flowing aroalso be strong enough to completely contain any explosion withdevices must include:

(a) An automatic shut-off controller, to operate if thehigh.

(b) Automatic fuel cut-off in the event of any malfunc

(c) Totally adequate fire protection in the event of faicombustion chamber.



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FIGUCombHeatin

m a gas turbine compressor is not the engine accessory gearbox. Such in some turbo-prop aircraft. These

f the centrifugal type.



RE 4-4ustion g

Engine-Driven Blower Systems26. For pressurised aircraft, where a supply of bleed air froavailable, cabin air supply is provided by blowers driven througha system was necessary in piston engined airliners and is still usedblowers are either of the positive displacement (Roots) type, or o



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low necessary in order to maintainuise rpm. Consequently, at sea leveligher mass air flow than is required.e excess mass air flow is dumpedhis system is shown schematically at

e-driven blower is dependent upon. If this excess mass flow is restrictedr discharged by the blower to a levelrge air is sensed and used to controle valve will be progressively openedure will cause the valve to partiallyl often used for this is a wheatstone



27. The blower must be capable of supplying the mass air fcabin pressurisation at maximum altitude with the engine at crwith the engine at full power the blower will be delivering a far hIn order to prevent over-pressurisation of the air ducting thoverboard by spill valves, controlled by a mass flow controller. TFigure 4-5.

28. In such a system, the mass flow produced by the enginblower rotary speed, which is limited only by the drive gear ratiothe resulting back pressure will increase the temperature of the aisufficient for cabin heating. The temperature of the blower dischathe position of the choke heat valve. As temperature increases thto prevent excess pressure/temperature. A reduction of temperatclose, increasing pressure and temperature. The method of controbridge circuit.



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FIGUMass FContro



RE 4-5low ller



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onditioning pack. This chilled air isir supply to the aircraft cabin at a

erfly valves, the setting of which isemperature is sensed and the bridgecrease or decrease the quantities of mixing unit is shown at Figure 4-5.gh 180° to drive the valves from then. This principle is illustrated

FIGUButterCold CValves

aft. Air for the aircraft pneumatich pressure stage (HP) of the enginepressure of air supplied at reducedlly bled from the high pressure stageailable at all thrust settings.



29. Some of the hot air from the blower is cooled in the air cmixed with hot air, which has by-passed the pack, to provide acomfortable temperature. The mixing unit consists of two buttcontrolled by a wheatstone bridge circuit. The cabin supply air tcircuit used to position the ‘hot’ and ‘cold’ butterfly valves to inhot or cold air supplied to the cabin. Such a temperature controlThe two valves are geared to a single spindle which rotates throu‘maximum cold’ position to the ‘maximum hot’ positiodiagrammatically at Figure 4-6.

RE 4-6fly Heat/ontrol

Engine Bleed Air Systems30. This system is widely used on modern jet-engined aircrsystem is generally bled from an intermediate stage (IP) and highigh pressure (HP) compressor. In order to supplement the lowthrust settings (for example during the descent), air is automatica(HP) of the compressor. This ensures that adequate pressure is av



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FIGUFlow CValve

d by an electro-pneumatic pressureorporates a shut-off facility which ise illustrated at Figure 4-7 is in effectf the air flow passing through the

erse flow of air back into the enginey a duct pressure relief valve which

system is illustrated at Figure 4-8.



RE 4-7ontrol

31. The pressure and flow rate of the bleed air is controlleregulating valve or flow control valve. This valve additionally incused in the event of a system malfunction. The flow control valva variable orifice device, the orifice being set by the pressure ovalve. A non return valve (NRV) is used in order to prevent a revcompressor. Overpressure of the pneumatic system is prevented bis located downstream of the pressure regulating valve. A simple



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FIGUPneum

ressure gauges located on the flight leak from a pneumatic duct couldt to duct) without a noticeable lossbly) will be indicated by a loss ofs are fitted adjacent to the ducting or a continuous detection sensing



RE 4-8atic System

32. The pneumatic system is monitored by reference to duct pdeck and warning lights indicating a sensed duct leak. A smallresult in the activation of a thermal sensing device (fitted adjacenof duct pressure. A severe duct leak (rupture or broken assempressure on the associated duct pressure gauge. Thermal sensorthroughout its length. The detectors can be either a spot typeelement system.



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system to filter the air used for air

FIGUWater

, a water extractor and a humidifier.of humid air, in order to preventsituated downstream of the coolinger droplets in it than the warm airater extractors to prevent blockagee or routing a warm air bleed to therated at Figure 4-9.



33. Some systems incorporate an air cleaner in the bleed airconditioning before it enters the pneumatic system.

Humidity ControlRE 4-9 Separator

34. The humidity control system consists of two componentsThe water extractor is used to reduce the moisture content precipitation in the ducting and in the cabin itself. The unit is unit, since the cooled air is more likely to have condensed watupstream of the cooling unit. Protection devices are fitted to wdue to formation of ice. They may take the form of a bypass valvextractor should ice begin to build up. A water separator is illust



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FIGUHumid

ry dry and consequently the relativeion, would be less than the requiredrmal domestic water supply is used,onditioning supply duct. A pressures the supply of water if the duct air

icient atomization and precipitation unit is shown at Figure 4-10. Thetitudes by an altitude switch.



RE 4-10ifier Unit

35. The humidifier is used at high altitude, where the air is vehumidity of the air introduced to the cabin, without humidificatminima, resulting in crew/passenger discomfort. The aircraft's nothe water being atomised by air which is bled from the main air cswitch, which is located downstream of the humidifier itself, stoppressure drops below a preset level, since this would prevent effwould occur in the cabin. A schematic diagram of a humidifierhumidifier is automatically prevented from operating at lower al



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l invariably need to be cooled, sincery hot. The most obvious source ofat from the cabin fresh air supply.subsonic speeds, ambient air alonee level. Furthermore, even in lowersignificant weight and drag penalty. temperature of the fresh air supplybin temperature can be maintained.ration.

erature must also be reduced. An to make the air drive a turbine. Thetemperature to drop. The more the greater the temperature drop. Thises is the means by which the turbinee turbine is in excess of 100°C.



Air Cooling Systems36. The air which is supplied to the cabin and flight deck wilthe source of supply is hot or, in the case of engine bleed air, vecooling is ambient air, used in a heat exchanger to extract heHowever, largely because of the ram air heating effect at high cannot reduce the cabin supply air temperature to an acceptablspeed aircraft the size of such a heat exchanger would impose a Some method of refrigeration is necessary in order to reduce theto a level where, by mixing chilled and hot air, a comfortable caMost modern passenger aircraft use an air cycle system of refrige

Air Cycle Cooling37. If the pressure of a volume of gas is reduced, its tempefficient means of reducing the pressure of the cabin air supply isexpansion of the air as it passes through the turbine causes its work done by the turbine, the more the air must expand and theprinciple is used in all air cycle refrigeration systems, all that variis loaded (made to do work). Typical temperature drop across th



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FIGUBrake



RE 4-11Turbine



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e simplest air cycle cooling unit inM).

ssure, so the brake turbine air cyclee compressors. Engine bleed air is

ome of its heat without significantlye a turbine, which in turn drives ale in the heat exchanger outlet andger.

typically less than 12 lb (5 kg). An, as a brake load on the turbine. Thee high supply pressure necessary to

ans that brake turbine cold air unitswered aircraft, where HP bleed air



38. A brake turbine is illustrated at Figure 4-11 and is thcommon use. Commonly referred to as an air cycle machine (AC

39. In order to drive an air turbine the air must be at high premachine uses bleed air from the high pressure end of the enginpassed through a ram air-cooled heat exchanger, which extracts sreducing its pressure. The high pressure air is then used to drivcompressor. The compressor discharge is ejected through a nozzthis nozzle induces the ambient air flow through the heat exchan

40. The unit has the benefit of simplicity and light weight, alternative form of brake unit uses a fan, instead of a compressorfan is used to pump ambient air through the heat exchanger. Thachieve a large pressure/temperature ratio across the turbine meare only suitable for use in multi-spool turbo-jet or turbo-fan popressure is high.



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FIGUBootstSystem

here the air supply pressure is notrap system can be used. The namets pressure by its own bootstraps.



RE 4-12rap Cooling

41. A bootstrap cooling system is illustrated at Figure 4-12. Whigh enough to satisfactorily operate a brake turbine a bootstbootstrap derives from the way in which the air apparently lifts i



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is pressure-boosted by the turbine-te pressure/temperature drop acrossrcooler, is necessary to remove the electrically-driven fan provides theen ram air is not present. (Aircraft

nd pressure as air flows through the

FIGUBootstPressuTempeRelatio

lex than the brake turbine, but itgh pressure bleed air is not availabletilising high by-pass ratio engines.



(a) The low pressure bleed air (or blower output) driven compressor in order to produce an adequathe turbine. A secondary heat exchanger, or inteunwanted temperature rise in the compressor. Ancooling air flow through the heat exchanger whparked).

42. Figure 4-13 shows the relationship between temperature abootstrap system.

RE 4-13rap re & rature nship

43. The bootstrap cold air unit is heavier and more comprequires less power to operate. It is favoured in aircraft where hior is undesirable, for example in small turbo-prop and aircraft u



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ure 4-14 is of the type fitted to the



44. The air conditioning system shown schematically at FigBoeing 737.



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FIGUAir CoSystem



RE 4-14nditioning



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hich are completely independent of conditioned air to maintain a cabinaximum certified ceiling. Normallye right pack from the number twomanifold. The right system only is

by its pack valve. The flow of airwo air mix valves. In the automaticve the requisite mix of hot and cold

set according to the position of theair mix valves are driven to the fullial cabin overheat on pack start-up.

the ram air flow is automaticallylouvres, according to the sensed airund, or during slow flight (flaps

o-fan (pack cooling fan) operates toontrolled by the turbo fan air valvedepending on aircraft position and

machine turbine, to remove water sensor in the separator controls anious temperature sensors are located



45. The aircraft has two air cycle air conditioning packs weach other and either of which is capable of supplying sufficientaltitude of 8000 feet or less when the aircraft is operated at its mthe left pack uses bleed air from the number one engine and thengine. The output of the two packs is combined in the mix illustrated at Figure 4-14.

46. Flow of air to the air conditioning pack is controlled through the pack, or by-passing the pack, is controlled by the tsetting these are positioned by the temperature controller to achieair in the mixing chamber. In the manual mode the valves are temperature selector control. When the pack valve is closed the cold position (by-pass closed, air cycle inlet open) to prevent init

47. The heat exchangers are cooled by ram air in flight, controlled by adjustment of the ram air inlet door and exhaust cycle machine compressor discharge temperature. On the groextended) the inlet door is full open and the bleed air driven turbaugment the ram air flow. The pack cooling fan (turbo fan) is cwhich passes pneumatic air to the turbo fan at varying flows configuration i.e. ground/airborne or slow/high speed flight.

48. A water separator is located downstream of the air cyclewhich has condensed during the cooling process. A temperatureanti-ice by-pass valve to prevent icing in the water separator. Varthroughout the air conditioning system. They function to:



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or control and indication of cabin

he flight deck. Some systems includetically.

a warm air bypass to remove any

or in some cases auto shutdown of

supplement the air temperature in zone. Individual zone temperaturef the packs (ACM). However, trim

.e. demand for most cooling).

entilate the aircraft in flight in thems failure.

t distributions system. There is noow may be provided.

ntilation flow.

, the aircraft must descend to below



(a) sense cabin temperature and provide a signal ftemperature on the flight deck.

(b) sense duct temperature and provide a warning to tshutdown of system on overheat detection automa

(c) warn of hot air leaks from the ducting.

(d) sense blocking of water separator and provide formation of ice.

(e) sense compressor outlet temperature for indicationsystem following overheat.

49. On some large jet transport aircraft it is necessary to individual zones to satisfy the comfort of the occupants of thatrequirements are satisfied by adding hot trim air to the output oair is not added to the zone controlling the output of the packs (i

50. On all pressurised aircraft, provision must be made to vunlikely event of total pneumatic supply or air conditioning syste

51. A means is provided to admit ram air into the aircraftemperature control of this ram air, but a means of varying the fl

52. The cabin pressure outflow valve is opened to ensure a ve

53. Prior to operation of the unpressurised ventilation system10,000ft, terrain clearance permitting.



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FIGURam AVentila



RE 4-15ir tion



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refrigerator) air cooling systems areft is on the ground. Vapour cyclen, a liquid absorbs latent heat from

latent heat removed from the cabinpply being chilled. The refrigerantnt of about 3°C at ISA msl pressure.6.

f the evaporator (a heat exchanger)evaporator as a liquid, however thetor causes the refrigerant to boil and the cabin air supply.

ing point, of the refrigerant before itm air passing over the tubes of the

liquid, giving up latent heat to the

ve which is an integral part of therant to start evaporating the instantit leaves the coil. A thermal elemente in temperature at the suction lineemperature change in the thermaltrols the refrigerant entering the

nk acts as a reservoir.



Vapour Cycle Cooling54. In some large passenger transport aircraft, vapour cycle (installed to supplement the air cycle systems whilst the aircrarefrigeration systems make use of the fact that, during evaporatioits surroundings. By using a liquid with a low boiling point, thefresh air supply results in the temperature of the cabin air sutypically used in vapour cycle systems is freon, with a boiling poiA vapour cycle cold air unit is shown schematically at Figure 4-1

55. Refrigerant at low pressure is drawn through the tubes oby the air turbine-driven compressor. The refrigerant enters the heat of the cabin air supply passing over the tubes of the evaporathe latent heat absorbed during this process of evaporation cools

56. The compressor raises the pressure, and therefore the boilenters the second heat exchanger in the cycle, the condenser. Racondenser cools the refrigerant, which condenses back into a ambient ram air as it does so.

57. The pressurised liquid passes through an expansion valevaporator situated at the inlet side. The valve causes the refrigeit enters the evaporator and to be completely evaporated before is attached to the suction side of the evaporator and any changcauses a corresponding change in the thermal element. The telement is signalled to the expansion valve which then conevaporator according to temperature demand. The refrigerant ta



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utput from the evaporator by meanse cabin at a controlled temperature

ven, the evaporator being capable ofrbine.

be used on the ground without thed do not require a high pressure air



58. Turbine outlet and inlet air is mixed with the chilled air oof a thermostatically operated series of valves. Air then enters thof about 20°C.

59. In many aircraft systems the compressor is electrically-drichilling the cabin air supply without the need of a pre-cooling tu

60. A major advantage of vapour cycle cooling is that it canneed to run the engines. Units are lighter than air cycle units ansupply.



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FIGUVapouCold A



RE 4-16r Cycle ir Unit



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n. Insufficient oxygen, or hypoxia,ase in altitude both air pressure and

ases in the air will remain constant,he amount of oxygen inhaled will

during normal respiration, and thise lungs. The lower the pressure, theefficiently.

t air pressure and oxygen to enabletificial method must be employed tothe conditions of air pressure andy achieves conditions equivalent tomstances the aircraft is said to be



Pressurisation61. In order to stay alive the human body requires oxygeresults in loss of consciousness and, eventually, death. With increair density decrease and this has a twofold effect:

(a) With decreased air density the proportions of the gbut will quite simply be reduced in quantity. Ttherefore reduce.

(b) The body absorbs oxygen through the lung tissuesprocess is assisted by the pressure of air within thmore difficult it becomes for the lungs to perform

62. Up to an altitude of 10,000 feet (3.3 km) there is sufficienthe human body to perform efficiently. Above this level some arproduce an environment inside the aircraft which equates to density at or below 10,000 feet. Cabin pressurisation normallthose of about 8,000 feet (2.6 km) or less. Under these circuoperating at a cabin altitude of 8,000 feet or less.



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the aeroplane must be able to 00ft in event of any probable

s, controls and indicators, for

e positive pressure differential to w delivered by the pressure must be large enough so that the ciable rise in the pressure hen the internal pressure is

their equivalent) to automatically damage the structure. However,

ably precludes its malfunctioning.

rapidly equalised.

the intake or exhaust airflow, or re and airflow rates.

ifferential, the cabin pressure ltitude.

when the safe or pre-set pressure altitude of 10,000ft is exceeded.



JAR-OPS 23.841 Pressurised Cabins

(a) If certification for operation over 25,000ft is requested, maintain a cabin pressure altitude of not more than 15,0failure or malfunction in the pressurisation system.

(b) Pressurised cabins must have at least the following valvecontrolling cabin pressure.

(1) Two pressure relief valves to automatically limit tha predetermined value at the maximum rate of flosource. The combined capacity of the relief valvesfailure of any one valve would not cause an appredifferential. The pressure differential is positive wgreater than the external.

(2) Two reserve pressure differential relief valves (or prevent a negative pressure differential that wouldone valve is enough if it is of a design that reason

(3) A means by which the pressure differential can be

(4) An automatic or manual regulator for controllingboth, for maintaining the required internal pressu

(5) Instruments to indicate to the pilot the pressure daltitude and the rate of change of cabin pressure a

(6) Warning indication at the pilot station to indicatedifferential is exceeded and when a cabin pressure



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plexity of the pressurisation system,ithstand the considerable pressureen the inside and the outside of theas hoop stress) which the hull must

re pressurised, whilst areas such ashis reduces the number of potentialn points where, for example, the

undergo significant stress each timecle). Just as continually bending andreak, so the repetitive stresses oftunately, those clever engineers canement takes place in good time. Thee fatigue life of an aircraft.

not designed for pressure ng in combination with landing

divert airflow from the cabin if pressor or continued flow of any function occurs.



63. The main problem with pressurised aircraft is not the combut the fact that the airframe must be strong enough to wdifferential. Pressure differential is the pressure difference betwepressurised hull of the aircraft. It imposes large stresses (known be capable of withstanding.

64. The flight deck, passenger areas and main cargo holds athe wheel wells and the tail cone are normally unpressurised. Tleakages of cabin pressure, but introduces stress concentratiofuselage meets the fore and aft pressure bulkheads. These pointsthe aircraft is pressurised and de-pressurised (a pressurisation cystraightening a piece of wire will cause it to fatigue and bpressurisation cycles will eventually lead to fatigue failure. Forcalculate the fatigue life of an aircraft structure and ensure replacnumber of pressurisation cycles is a major element in assessing th

(7) A warning placard for the pilot if the structure is differentials up to the maximum relief valve settiloads.

(8) A means to stop rotation of the compressor or tocontinued rotation of an engine-driven cabin comcompressor bleed air will create a hazard if a mal



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through the first few thousand feet,t altitude. For this reason, a cabinn mean sea level pressure. By thist within the bounds of engineering

sure whilst the aircraft was flying at

ould be 12 lbs/in2. By maintaining a

8.3 lbs/in2.

ccommodate a maximum operating

t, by governing the rate at which airrate at which cabin air is released to which passes a signal to discharged to maintain the required cabin air

s/in2)

in2)

2)



65. The rate of change of pressure from the surface upwards,is far greater than for the same number of thousands of feet apressure altitude of 8,000 feet or less is maintained, rather thameans the pressure differentials the hull must withstand are kepfeasibility. Table 1 illustrates this point.

66. If the cabin were to be maintained at mean sea level pres

40,000 feet, the pressure differential across the skin of the hull w

cabin altitude of 8,000 feet the pressure differential is reduced to

67. Modern aircraft pressure hulls are normally designed to a

pressure differential of 8.6 lbs/in2 to 8.9 lbs/in2 .

68. Cabin pressurisation is not controlled, as one might expecis supplied from the air conditioning units, but by governing the atmosphere. This is achieved by means of a pressure controller,valves. These valves restrict the rate at which cabin air is releasepressure (cabin altitude).

Altitude Air Pressure

Mean Sea Level 1013 mb (14.7 lb

8,000 feet 758 mb (11.0 lbs/

40,000 feet 186 mb (2.7 lbs/in



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vel flight, the pressure controller/ of change of cabin altitude as the per minute and 500 feet per minutedetermined by the controller settingescent of the aircraft. Maximum

alves, pressure limiting safety valvesl integrity of the hull in the event of

ut they basically consist of pressureh cabin and external pressures) andtude and its rate of change. As the senses the change relative to thes. This signal adjusts the dischargecient to achieve the desired pressure

When activated, this will signal aller. At the same time, all compressorxits would be impeded if the cabin a plug type door measuring 6 feet

uld be 1.16 tons.



69. As well as maintaining a constant cabin altitude in ledischarge valve arrangement is designed to ensure a steady rateaircraft climbs or descends. A rate of change of between 300 feetis normally considered to be comfortable. The rate of change is selected by the flight deck crew and by the rate of climb/dpermissible rates of change of cabin altitude are discussed later.

70. In addition to the pressure controller and the discharge vand inward relief valves are also fitted to safeguard the structurafailure of the basic components.

71. Pressure controllers vary in construction and operation, bsensing capsules and/or diaphragms (which are subjected to botmetering valves and controls for setting the required cabin alticabin pressure changes, the pressure controller automaticallyexternal pressure and transmits a signal to the discharge valvevalve opening until the release of air from the cabin is just suffidifferential.

72. On some pressure controllers a ditching control is fitted.discharge valves to fully close in order to minimise inflow of watoutput will be dumped, otherwise operation of the emergency eremained pressurised. Appreciate that the force required to open

by 3 feet, with a positive differential pressure of just 1 lb/in2, wo



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atic dumping facility linked to theround, the safety (dump) valve will aircraft. Alternatively, the pressuretflow valves. In both cases manuale aircraft weight (squat) switch.

ol panels, warning operators not to. The restricted negative differentialft type. This is to ensure that undue occur during take-off and landing

he following list:

te which will, at the maximum ratedifferential to a given value. (If themb it is permitted to exceed this2).

shall be such that failure of one halffferential.

negative differential pressure from



73. Some pressurisation systems incorporate a manual/automsafety valve. This is to ensure that, whilst the aircraft is on the gbe held in the open position preventing the pressurisation of thedumping facility on some systems is linked to the discharge/ouswitching is available to the pilot or automatic operation is via th

74. Placards are displayed on flight deck pressurisation contrexceed nominated pressure differentials on take-off and landingpressures typically range from .1 psi to .5 psi depending on aircrastresses are not imposed on the airframe structure which couldwith excess negative differential pressure.

Safety Devices75. The requirement for various safety devices is outlined in t

(a) The aircraft must be fitted with devices in duplicaof flow, automatically limit the positive pressure aircraft is climbing at its maximum rate of cli

maximum pressure differential value by 0.25 lb/in

(b) The capacity of each half of the duplicated systemwould not cause appreciable fall in the pressure di

(c) Devices must be fitted in duplicate to prevent the

exceeding 0.5 lb/in2.



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ch permits the crew to reduce the

automatic regulator (the pressure rate of change of cabin pressure.

tem controls are not situated at thede must be fitted at the pilot station.ly shall have an indication of cabinximum differential pressure shall be

means of equalising the pressureould result from a rapid loss of

cause collapsing of floors, panels,ample would be the cabin floor and



(d) A manually operated device must be fitted whipressure differential to zero.

(e) The system must have a suitable manual or controller) for controlling both cabin pressure and

(f) Suitable instruments must be provided. If the syspilot station, an instrument indicating cabin altituThe crew member in charge of the cabin air suppaltitude and differential pressure. The value of maindicated on or near the instrument.

76. Wide bodied jet transport aircraft are fitted with a differential between sections inside the pressure hull which wpressurisation. A rapid loss of pressurisation in one area couldwalls etc. separating other areas in the pressure hull. A typical excargo department (see Figure 4-17).



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FIGUPressuPanels



RE 4-17re Balance



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rther discussion, namely the concept to withstand the forces imposed byitive differential pressure). If, duringft descended to mean sea level, theside (negative differential pressure).ich the pressure hull is designed. Inoors would be extremely hazardous

rom occurring, inward relief valvest a differential pressure of, typically,

essure.

ensure that the absolute maximumlly combined with the inward reliefslightly higher then the controlled lightly spring-loaded in the shut on the ground to ensure pressures



77. Of the items on the above list perhaps one is worthy of fuof negative differential pressure. The pressure hull is strengthenedrelatively high pressure acting outwards on the pressure hull (posa descent, the cabin altitude stuck at 8,000 feet whilst the aircrapressure outside the aircraft would be greater than the pressure inHence the pressure loading would be the reverse of that for whthis condition, on the ground, the operation of inward opening dand the operation of outward opening doors impossible.

78. To prevent a condition of negative differential pressure fare fitted to the pressure hull. These valves open automatically a

0.5 lb/in2, whenever the outside air pressure exceeds cabin air pr

79. The action of the pressure limiting safety valves is to pressure differential is never exceeded. This safety valve is usuavalve and is spring-loaded to open at a differential pressure maximum differential pressure. In normal circumstances it isposition. In many aircraft the safety valve may be selected OPENare equalised. It is selected CLOSED prior to take-off.



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pressurisation control system of arised by a continuous inflow ofd cabin pressure is controlled by proportional relationship betweendescent, and a constant differentialfety valves (one only shown) and an

erated main outflow valve. During situations a DC motor takes over.ol:

the static vents, with barometricfficer's altimeter (STBY). In standby (ADC) is also used, in the systemers, together with signals from theitch.

of AUTO failure.

e operation of the outflow valve.



Pressurisation System Description80. Figure 4-18 shows the schematic arrangement of the modern passenger transport aircraft. The aircraft is pressuconditioned air through the overhead distribution ducts, anregulation of the outflow of air. The normal procedure is for aambient and cabin pressure to be maintained in the climb and pressure to be maintained in the cruise. Two pressure-relieving sainward-relieving negative relief valve are fitted.

81. Pressure control is achieved with the electric motor-opnormal operations an AC motor drives the valve, in emergencyThe cabin pressurisation operates in one of three modes of contr

82. Aircraft altitude is sensed barometrically direct from connections fed from the Captain's altimeter (AUTO) or First Omode an electrical altitude signal from the Air Data Computerillustrated. Cabin pressure is also fed to the pressure controllundercarriage squat switch and the flight/ground (FLT/GRD) sw

AUTO The normal mode of operation.

STANDBY A semi-automatic system used in the event

MANUAL Manual control of cabin pressure by remot



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l panel is ‘set up’ by inserting thert altitude (LAND ALT). Whilst on continuously.

ercarriage squat switch transmits as the controller from initiating cabinny cabin pressurisation by holding



AUTO Mode Operation The AUTO pressurisation controintended flight or cruise altitude (FLT ALT) and the landing airpothe ground the departure airport altitude (cabin altitude) is fed in

83. Whilst the aircraft weight is on the main wheels the undsignal to the AUTO pressurisation controller. This signal preventclimb. The FLT/GRD switch, when selected to GRD, prevents athe main outflow valve full open.



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FIGUPressuContro



RE 4-18risation l System



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FLT and the controller moves the.1 psi positive differential pressureansition to pressurised flight will beressurisation controller switches tortional to aircraft rate of climb, butuld the aircraft be required to holde of climb is proportional to aircraft

selected flight altitude (FLT ALT) ahes from proportional to isobaric by maintaining a constant pressure feet) between cabin and ambient

s tripped (whilst still in proportional rate proportional to aircraft rate ofclimb, the pre-set value of FLT ALTfacility is lost.

eases the controller will allow themaintain constant cabin altitude. Ifaintained and the cabin will climbfor minor excursions from cruisen must be reset. If the differential7.45 psi at 28,000 feet or lower) theet below the selected LAND ALT.



84. When taxiing commences the FLT/GRD switch is set tomain outflow valve towards close, pressurising the cabin to 0(equivalent to airport altitude -200 feet). This ensures that the trgradual. Once airborne (signalled by the squat switch) the pproportional control. Cabin altitude is increased at a rate proponot greater than the maximum cabin rate of climb selected. Shoduring the climb, the cabin altitude will hold also, since cabin ratrate of climb.

85. When the outside air pressure is 0.25 psi above that of thecruise relay is tripped and the pressurisation controller switccontrol. The controller now maintains a constant cabin altitudedifferential (7.45 psi up to 28,000 feet, 7.8 psi above 28,000pressures.

86. If the aircraft begins to descend before the cruise relay hacontrol), the pressurisation controller reduces cabin altitude, at adescent, to land at the departure airfield elevation. If, during the is changed, this automatic reduction to departure field elevation

87. During isobaric (cruise) control, if aircraft altitude incrpressure differential to reach a maximum of 7.9 psi in order to the aircraft climbs further this maximum differential will be mmaintaining maximum differential. This feature is adequate altitude, but for a higher cruise altitude the FLT ALT selectiopressure between FLT ALT and LAND ALT is less than 7.8 psi (controller maintains the cabin altitude during the cruise at 300 fe



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ssure is 0.25 psi above that of theurisation controller to proportionalraft rate of descent, but not greaterouch-down, the squat switch signal,ntroller to adjust cabin pressure to

e controller drives the main outflowressures.

to STANDBY mode in the event of:

t per minute.

l mode is shown at Figure 4-19.



88. At commencement of the descent, when outside air preselected FLT ALT, the descent relay trips and switches the presscontrol. Cabin rate of descent is now kept proportional to aircthan the maximum rate of change of cabin altitude selected. On twith the FLT/GRD switch still in the FLT position, causes the co0.1 psi above ambient pressure (airfield altitude -200 feet).

89. During taxi in the FLT/GRD switch is put to GRD and thvalve to the full open position to equalise cabin and outside air p

90. The pressurisation control system will automatically trip

(a) Loss of AC power supply.

(b) Cabin rate of altitude change in excess of 1800 fee

(c) Excess cabin altitude (14,000 feet).

91. A typical flight profile in the AUTO pressurisation contro



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FIGUAuto PressuContro



RE 4-19

risation l Mode



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the main outflow valve will be heldurisation controller drives the mainBIN ALT should therefore be set to

l panel is used in selecting a cabinerential. Cabin rate of climb/descente of settings from 150 fpm to 1800

elow destination airfield altitude, to

fail, the main outflow valve opening, by reference to the valve position

liar to the system illustrated atues and it is important that you areh are relevant to your aircraft. For

ressurised at all times whilst on the. Other systems allow for slight In any system, manual operation ofck crew, especially during the climb workload is already high.



Standby Mode Operation92. On the ground, with the FLT/GRD switch set to GRD, full open. With the switch in the FLT position the standby pressoutflow valve to attempt to achieve the selected CABIN ALT. CA200 feet below departure airfield altitude.

93. In flight a reference placard on the pressurisation controaltitude based upon the intended flight altitude and pressure diffis selected using the cabin rate selection knob, which has a rangfpm, and is detented to 300 fpm.

94. In the descent the cabin altitude is selected to 200 feet bensure that the cabin is slightly pressurised on touch-down.

Manual Mode Operation If both AUTO and STBY modes can be modulated by means of the main outflow valve controlindicator and cabin altitude gauge.

95. The operating procedures described above are pecuFigure 4-18. Different systems require different operating techniqtotally familiar with the normal and abnormal procedures whicinstance, some operating systems require the aircraft to be depground, with pressurisation commencing following take-offpressurisation prior to take-off and depressurisation on landing.the main outflow valve imposes a high workload on the flight deand descent, and particularly during a stepped descent, when the



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inimum instrumentation:

or descent.

ed with an aural warning) which

tween cabin pressure and ambient

brated to display this in terms of

ocedures for the various modes of of rapid cabin depressurisation can

ibility to the common cold by virtuecabin rates of change are kept tomon for passengers to be tested fore carefully monitored and wherever



Instrumentation96. All aircraft pressurisation systems contain the following m

Vertical speed indicator. This shows the rate of cabin climb

Cabin altitude warning light. A red light (often associatilluminates if cabin altitude reaches 10,000 feet.

Differential pressure gauge. This shows the difference beatmospheric pressure.

Cabin altimeter. A gauge reading cabin pressure, but caliequivalent altitude.

Pressurisation System Procedures97. The operations manual for the aircraft contains the prpressurisation control. The procedure to be followed in the eventbe found in the emergency procedures section.

Limitations98. Because of the structure of the human ear and its susceptof inflammation of the eustachian tubes, it is important that moderate values, especially in the descent. Because it is not comrunny noses before flight, cabin altitude rates of change should bpossible descent rates in particular kept to a minimum.



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through uncontrolled means, anyust be adequately pressure sealed,

times to ensure that there are nodevices are functioning correctly.

uctural strength of the pressure hull.

ak rate tests and functioning tests isf aircraft. A general guide is given

g the engines or by using an externalen functional testing, as this checksrate testing, since this removes the jet pipes and propellers. It also has



Ground Testing99. In attempting to limit the escape of pressurised air mechanical linkage which passes through the pressure hull mirrespective of the type of linkage.

100. Pressurisation systems must be checked at appropriateserious leaks and that the pressure control equipment and safety

101. The occasions on which tests are carried out are:

(a) When specified in the maintenance manual.

(b) After repairs or modifications which affect the str

(c) After suspected damage to the fuselage.

102. The exact procedure to be followed when carrying out lelaid down in the maintenance manual for the particular type obelow.

103. The functional testing can be carried out either by runninsource of supply. It is generally preferred to run the engines whthe whole system. An external supply is preferred when leak hazard of checking for leaks in the proximity of engine intakes,the advantage that leaks can often be heard.



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valve/squat switch and ensure that

y the manufacturer). This includes

oning controls set as specified by the

imum differential working pressure

k each one separately.

als.

ly to prevent moisture precipitation



Functional Test(a) Refer to the maintenance manual. Override dump

the dump valve is fully closed.

(b) Close all external doors and windows.

(c) Open internal doors (unless otherwise specified bcupboards, galleys and so on.

(d) Electrical power on, pressurisation and air conditimanufacturer.

(e) Introduce air supply at the appropriate rate.

(f) Watch pressure increase until it stabilises at maxand remains stable.

(g) If more than one pressure controller is fitted, chec

(h) Operate flying controls to check the integrity of se

(i) Check automatic action of Safety Valves.

(j) On conclusion of test, allow pressure to drop slowin the cabin.



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te. Allow pressure to stabilise, then pressure to drop over the range maintenance manual.

eat.

r, but a soapy water solution can be

efore opening doors and windows

the condition they were in before

pen and close all hatches, windows,movement.

transparencies for crazing.

ed in the cabin. A minimum of twosting. If the test is being carried out the engines. All operators must be.



Leak Rate Check(a) Introduce air into the cabin at the appropriate ra

shut off air supply and check time taken forappropriate to the aircraft type. Check against the

(b) If leak rate is excessive, investigate, rectify and rep

(c) Detection of leaks can usually be carried out by eaused to produce bubbles at the point of leakage.

(d) On conclusion of the test shut off air supply. Bensure pressure differential is zero.

(e) Restore all pressurisation control components totesting.

(f) Examine fuselage for deformation and damage. Ointernal doors and so on, to check for freedom of

(g) Examine pressure bulkheads for deformation, and

104. It should be noted that in both tests observers are requiroperators should be inside the pressurised area during pressure tewith the engines running, a third person is required to supervisecertified fit, this includes freedom from colds and sinus problems






Conditions Conducive to Aircraft Icing

Ice Warning

De-Icing Systems

Anti-Icing

Pitot Static Pressure Sensors and Stall Warning Devices

Rain Removal



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airframe and the power units are

Icingcumulation of ice accretions underflight can encounter a variety ofn produce ice accretions on various

ed water droplets (below 0°C) thatter droplets will freeze upon impactin in the liquid state at ambientcretion on an aircraft component is size, cloud liquid water content, and velocity.

y cold temperatures where moistureme icing does not occur in these

res below 0°C containing a mixture



5Ice and Rain Protection

1. In this chapter the systems employed to ensure that theprotected from the effects of ice and rain are considered.

Conditions Conducive to Aircraft 2. Aircraft on the ground or in flight are susceptible to acvarious atmospheric and operational conditions. Aircraft in atmospheric conditions which will individually or in combinatiocomponents of the aircraft. These conditions include:

(a) Supercooled clouds. Clouds containing supercoolhave remained in the liquid state. Supercooled wawith another object. Water droplets can rematemperatures as low as -40°C. The rate of ice acdependent upon many factors such as dropletambient temperature, and component size, shape,

(b) Ice Crystal Clouds. Clouds existing usually at verhas frozen to the solid or crystal state. Airfraconditions.

(c) Mixed Conditions. Clouds at ambient temperatuof ice crystals and supercooled water droplets.



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within clouds or below clouds atlets remain in the supercooled liquid

operations, are susceptible to manyn to conditions peculiar to ground

aining moisture, slush, or snow.

ings, or other ground structures.

round support equipment.

eller, or rotor wash. Operation of jetrs, and helicopter rotor blades are



(d) Freezing Rain and Drizzle. Precipitation existingambient temperatures below 0°C where rain dropstate.

Icing Problems on the Ground3. Aircraft on the ground, during ground storage or groundof the conditions that can be encountered in flight in additiooperations. These include:

(a) Supercooled ground fog and ice clouds.

(b) Operation on ramps, taxiways, and runways cont

(c) Blown snow from snow drifts, other aircraft, build

(d) Snow blown by ambient winds, other aircraft or g

(e) Recirculated snow made airborne by engine, propengines in reverse thrust, reverse pitch propellecommon causes of snow recirculation.



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roduce frost accretions on aircraftrost point. Frost accumulations arefter landing where aircraft surface higher altitudes. This is a commonf fuel cells. Frost accretions can alsold fuel (with the wing tanks full of

y be achieved on the ground byPD) fluid. In the air, surfaces proneller blades, engine intakes) may be

ed on the aircraft surface. On theuid. In flight the susceptible surfacesedges (de-icing boots) may be fitted

detection devices which activate a ice detection system consists of anstream from the side of the fuselage.



(f) Conditions of high relative humidity that may psurfaces having a temperature at or below the fcommon during overnight ground storage and atemperatures remain cold following descent fromoccurrence on lower wing surfaces in the vicinity ooccur on upper wing surfaces in contact with cocold-soaked fuel).

Anti-icing and De-icing4. Anti-icing is the prevention of ice formation and marepeatedly coating the aircraft with a freezing point depressant (Fto ice formation (leading edges of wing and tail surfaces, propeheated electrically, by hot bleed air or by heated fluids.

5. De-icing is the removal of ice or frost which has formground this may be done by spraying the aircraft with hot FPD flmay be heated as described above. Alternatively, flexing leading on lower performance aircraft.

Ice Warning6. In-flight warning of ice formation is provided by ice warning indicator on the flight deck. The traditional automaticelectric motor driving a serrated rotor which projects into the airAdjacent to the rotor is a fixed cutting edge.



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the cutting edge, the greater the icerque reaction trips a microswitch,ally at Figure 5-1.

rofoil section tube mounted at righteading edge. During flight air entersthe trailing edge, hence there is a the points of an electrical relay. Inssure in the tube falls and the relay. Vibrating rod ice detectors consist0 KHz. Ice forming on the probe

warning circuit. Both devices are



7. As ice builds up on the serrated rotor it is scraped off bybuild-up the greater the torque loading on the motor. The toactivating a cockpit warning. The system is shown diagrammatic

8. Pressure operated ice detector heads consist of a short aeangles to the air flow and having four small holes drilled in the lthese holes and can partially escape through smaller holes in pressure build-up in the tube and this pneumatically holds openicing conditions the leading edge holes become blocked, the prepoints close to illuminate a warning indication on the flight deckof a probe in the airstream which vibrates at approximately 4reduces the vibration frequency and activates the ice detectorillustrated at Figure 5-2.



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FIGUIce DeDeviceRotor

d externally and easily visible fromly form on these, warning the flightese visual detectors are fitted withtion.



RE 5-1tection - Serrated

9. Visual ice detectors, in the form of small aerofoils situatethe flight deck are often fitted. In icing conditions, ice will readicrew that it will also be forming in more critical locations. Thinternal heaters, so that they can be cleared for further ice evalua



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FIGUIce DeDeviceOperaVibrat



RE 5-2tection s - Pressure ted / ing Rod



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atically-operated de-icing systems.ached to the leading edges of wings,llel to the span of the flying surface

craft's compressed air system.

d to break away. When not in use a leading edge. De-icer boots may be. A diagram of a pneumatic de-icing



De-Icing Systems

Pneumatic De-Icing Systems10. Some piston-engined and turbo-prop aircraft use pneumThese take the form of flat inflatable tubes, closed at the ends, atttailplanes and fin. The rubberised fabric tubes normally run paraand they are inflated and deflated cyclically with air from the air

11. Inflation of the tube, or boot, causes any ice layer formevacuum source is applied to the boots to hold them flush with theattached to the airframe by screw fasteners or by bonding cementsystem is shown at Figure 5-3.



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FIGUPneumicing S

mance aircraft, since they obviously



RE 5-3atic De-ystem

12. De-icing boots are only suitable for relatively low perfordisturb airflow over the wing to some extent.



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ise an automatic alternating timeessure and vacuum gauges or lightsdamage from oils, greases, knocks,

an electrically-driven pump. Thearried rearward over the aerofoil town at Figure 5-4.



13. Flight deck controls and indications typically comprsequence of pressure and vacuum to the de-icing boots, with prand an on/off switch. The rubberised fabric is susceptible to abrasions and exposure to strong sunshine.

Fluid De-Icing Systems14. De-icing fluid is distributed to porous metal panels byescaping fluid destroys the bond between ice and aircraft and is cprevent further ice build-up. A typical fluid de-icing system is sho



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FIGUTypicaicing S

aft with no aerofoil de-icing system. produces a surface upon which icein flight. It must be emphasised that flight into known or forecast icing



RE 5-4l Fluid De-ystem

De-Icing Paste15. A hand applied de-icing paste is sometimes used on aircrIt is spread onto the leading edges of wings and empennage andmay form, but to which it will not bond. The ice then blows off the use of de-icing paste does not provide reliable protection forconditions.



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sage where it is unavoidable pilots must available, be it thermal, fluid or pilot for initiation and duration ofedures for this eventuality. Since de-

ed, the decision for activating the, most systems incorporate a cyclicm to de-ice aerofoil section leading de-ice the sections using a timed

of associated systems (redundancy) of limits.

or operation and control dependentroach and landing phase.

the propeller becomes heavier and,ltimately will cause damage to the

t-up. The removal is carried out by and to remove it by centrifugal andurrent is switched off, the ice begins.



Initiation / Timing of De-icing System U16. Where possible the icing situation should be avoided, butmake a critical analysis of the situation. Whatever system iselectrical, it must be fully utilised, requiring a judgement by theuse. Type operations manuals will lay down recommended procicing systems are designed to remove ice, once it has been formsystem is vital to ensure adequate removal. Once switched ontiming device, which automatically controls the de-icing mediuedges or propellers. This automatic control is programmed tosequence selected alternately.

17. The system is capable of continued operation with loss and will have warnings of normal operation and any exceedance

18. The type operations manual will contain the procedures fon flight profile i.e. some aircraft may require system off for app

Propeller De-Icing19. If ice is allowed to build-up and remain upon a propeller,more importantly, unbalanced. Vibration is generated which uengine and the airframe.

20. De-icing systems are used to remove ice after it has builmeans of electrical heating elements in the blades, to melt the iceaerodynamic forces. After a short period of time, the electrical cto form again, the current is switched on, the ice melts, and so on



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ontrol switch on the control panel,ter which indicates the current beingers. Modern systems tend to use one

-iced at any one moment, a timer orent to the propeller blades of the

controls power relays located in the kept to the minimum.

hub will be found the slip ring fromssed via brushes attached to the rear found on some installations). Since of time, only a thin film of ice willeristics of the propeller blades. Theith the heated propeller blade melts,mic and centrifugal forces.

r DC or AC. Although only one seture that there is a null period of ating switched off and the de-icing ofapplied to a set of propellers over a

the propeller de-icing is part of the heating circuit. The spinner is also

peller de-icing system.



21. The pilot controls the propeller de-icing by means of a csee Figure 5-5. Also on the control panel will be found a load medrawn. On early systems, one load meter served all of the propellload meter for each set of propellers.

22. Since the propeller blades of only one engine are being decycling unit is required to arrange for the application of currappropriate engine at any particular moment. The timer in fact engine nacelles. In this way, the length of heavy current wiring is

23. Mounted on the engine front casing behind the propeller which the current to the heating elements on the propellers is paof the propeller hub (but the opposite arrangement may also bethe propeller has remained unheated for a relatively short periodhave built up, without interfering with the aerodynamic charactdeposited ice acts as a thermal insulator. As the ice in contact wthe main body of ice is removed under the influence of aerodyna

24. According to aircraft type, the current used may be eitheof propellers is being de-iced at any one time, the timer will ensleast one second between the de-icing of one set of propellers bethe next set of propellers being switched on. Heating current is typical period of 30 seconds.

25. In the case of propellers attached to a turbo-prop engine,engine de-icing system, and the timer is part of the engine intakede-iced on turbo-prop installations.

26. Figure 5-5 shows a schematic diagram of an electrical pro



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systems using exhaust gas heatustion heater, system is used in some



27. Some turbo-prop aircraft employ thermal de-icingexchangers to produce the hot air required. An alternative, combpiston-engined aircraft.



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FIGUPropelSystem



RE 5-5ler De-icing



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ti-iced by hot compressor bleed airf valve controls the air flow into ae of the wing. Telescopic ducts feedrface.

FIGUThermSystem



Anti-Icing

Thermal Anti-icing Systems28. The leading edges of wings, tailplanes and fins may be ansupplied by the aircraft pneumatic system. A motorised shut-oftitanium distribution duct routed through the fixed leading edgthe hot air into a perforated duct (piccolo tube) inside the slat su

RE 5-6al Anti-icing



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the remainder of the slat structure.penings and under wing drain holes.g edge section.

is thermostatically controlled by al valves. The major source of failurestribution ducting, which can causeperature sensing probes are usually



29. The hot air anti-ices the slat leading edge and flows intoThe hot air is then exhausted overboard through the slat track oIf the slats are retracted the exhausting air warms the wing leadin

30. The temperature of the air supply to the leading edgessystem of leading edge temperature sensing units and duct controin these systems is leakage of the very hot bleed air from the didamage to other systems, particulary electrical circuits. High tempositioned at strategic locations adjacent to the ducting.



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FIGUSlat AnDuct



RE 5-7ti-icing



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FIGUTypicaAnti-ic

eck windscreens is widely used inaterial (can be referred to as Goldeating process will provide a non-rovided with normal and failurehe vinyl interlayer is the ‘fail-safe’ inner panel should fail. The outer a hard scratch resistant surface. A



WindshieldRE 5-8l Electrical e System

31. The use of electrical heating elements built into flight dmodern transport aircraft. A layer of transparent conductive mFilm) is supplied from the aircraft A/C electrical system. The hshattering quality to the window and the flight crew are pindications. The (inner) glass panel is the load-bearing agent. Tload carrying member and prevents the window shattering if theglass panel has no structural significance, it provides rigidity and



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anel to permit electrical heating for due to arcing can lead to visibilityde-fogging can be gained from the on the outer panel also assists inows a typical electrical windshield

h if neglected can spread and causeation is laid down by manufactures.

esulting in local overheating causinge control and in the extreme, panel

ct from removing ice build-up. Anti-anti-icing system must therefore be enter icing conditions. When fluiding point, and it must also combinecommonly used, since it is readilyof isopropyl alcohol is its high



conductive film is applied to the inner surface of the outer glass panti-icing and de-fogging. However damage of the outer panelproblems. For an electrical supply failure a limited amount of windscreen warm air de-misting supply. A conductive coatingdissipating static electricity from the windscreen. Figure 5-8 shanti-ice system.

32. Specific problems associated with this type of system are:

Delamination Separation of the vinyl windscreen plies whicvisibility or electrical problems. Limitations of allowable delamin

Arcing Indicated by a breakdown in the conductive coating rfurther damage to the panel. This can lead to loss of temperaturfailure.

Propeller Anti-Icing33. Anti-ice systems are used to prevent ice build-up, as distinice systems are generally ineffective once ice has built-up. Any switched on before it is anticipated that the aircraft is about toanti-icing systems are used, the fluid used must have a low freezreadily with water. For this reason, isopropyl alcohol is most available and of relatively low cost. The major drawback flammability.



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typical propeller anti-icing system. a contents gauge will be found onsure availability of the fluid at alls to permit gravity feed to the fluid

o that rate of flow can be controlledd 10 psi, but sufficient to open theing of the anti-icing fluid when the

e delivery pipe mounted in the frontar of the propeller hub. The fluid is influence of centrifugal force.

of pump pressure, a relief valve isgulating the fluid flow.

, in order to assist in the efficientas anti-icing boots, extend out along



34. Figure 5-9 shows diagrammatically the main features of aThe reservoir is positioned remotely from the flight deck, and sothe flight deck. The tank must be vented to atmosphere, to encontents levels. The position of the fluid reservoir must be such apump(s), via a filter. The fluid pump is controlled by a rheostat, sto match existing conditions. Pump pressure is quite low, arouncheck valves which have been found necessary to prevent syphonsystem is not in use. On reaching the propeller, the fluid leaves thend of the engine, and enters the slinger ring mounted at the redistributed to the propeller blades from the slinger ring under the

35. On some installations, air pressure may be used in placeinstalled in the air supply line together with a control valve for re

36. Additionally, propeller blades may have overshoes fitteddistribution of the fluid. These overshoes, sometimes referred to one third of the propeller blade.

37. Normally one pump will supply two propellers.



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FIGUPropelicing



RE 5-9ler Anti-



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a gas turbine or turbo fan engine isill utilize the pneumatic air supply

nes. Engines which use a probe forr to prevent build up on the probe.

yclic system to anti-ice intakes.

r to entering icing conditions on thee penalty which will be annotated in

switching on the flight deck and is the engine spinner. The system canated icing conditions. Switching only controlled and pressure operated.he flight deck. Correction to enginetomatic upon activation of the anti-f anti-icing, the engine ignition ist Operating Manual will containtallations the engine RPM must beude, to give adequate protection. air anti-icing system.



Power Plant and Air Intakes38. Powerplant. The Engine air intake nacelle and spinner ofanti-iced using the aircraft pneumatic air supply. Some systems wto additionally anti-ice the initial stage LP compressor stator vasensing intake pressure (P1 PROBE) will use the engine ant-ice ai

39. Turbo Prop aircraft use an electrical system employing a c

40. Switching for anti-ice systems is initiated by the pilot prioground or when airborne. Some systems may incur a performancthe type operations manual.

41. Air Intakes. Engine compressor bleed air is controlled byducted to the powerplant intake cowl and, in some installations,be operated on the ground or during flight in known or anticipthe system activates engine cowl anti-ice valves that are electricalSystem normal operation and failure warning is annunciated on tmaximum EPR (engine pressure ratio) indication and limits is auicing system. As a protection against flame-out during use oselected ON either manually or automatically. The aircrafprocedures for normal and abnormal operation. On some insmaintained within certain parameters, dependent upon altitFigure 5-10 shows a schematic illustration of a gas turbine bleed



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FIGUAnti-ic



RE 5-10ing B757



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Stall Warning

FIGUAnti-icLocatio



Pitot Static Pressure Sensors and Devices

RE 5-11e Probe ns

Figure 5-11 shows a schematic illustration of a probe locations.



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by aircraft electrical system power.nd failure warnings are provided oned on the ground to protect againstivated by the undercarriage weight manner.

s also will be anti-iced by electrical

se the operation of the windshieldmember about rain repellant fluid is

ill obscure the view through the

een which forms on the surface andlass in between. The beads of water

e flight deck. Irrespective of switch/d.



42. Pitot probes contain an electrical heating element suppliedOnce switched on they are thermostatically controlled. Normal athe flight deck. Some provision is made for lower voltage appliburning out the element. This type of control is normally act(squat) switches. Stall warning detectors are anti iced in a similar

43. Where a TAT (Total Air Temperature) system is used thielements.

Rain Removal44. During flight in heavy rain it may be necessary to ceawipers with application of a rain repellant. The main thing to rethat, when inadvertently selected with a dry windscreen, it wwindscreen rather than improving it.

Rain Repellent45. During heavy rain a chemical is sprayed onto the windscrcauses the rain water to form beads, leaving large areas of dry gare easily removed by the natural airstream.

46. Activation of the system is by switch or push button on thbutton operation only a measured amount of fluid will be applie



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faces because it can cause smearingally deteriorate with further rainpendent on rain intensity, type of



47. The system should not be applied to dry windscreen surand restrict visibility. After application the film will graduimpingement. The length of time between applications is derepellent and windscreen wiper usage.





Fuel & Fuel Systems

Aircraft Fuels

Refuelling

De-fuelling

Definitions

Fuel Systems

Fuel Tanks

Fuel Feed Systems

Fuel Jettisoning

Fuel System Monitoring

Example of Fuel Management

Fuel Quantity Measurement Systems

Fuel Measuring Sticks



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erosene (gas turbines). It is vital to

he UK as AVGAS 100 LL (low lead).nd must comply with Directorate of85. 100 LL has an octane rating of a lean mixture is used and the 130

available. 100 AVGAS is coloured0/130 octane rating under lean/richL) than 100 LL and is thereforered red and has a lean/rich octane

Fuel & Fuel Systems


6Fuel & Fuel Systems

Aircraft Fuels1. Aircraft engines use either gasoline (piston engines) or ksafety that the correct fuel is used.

Aviation Gasoline (AVGAS)2. Aviation gasoline (AVGAS) is only generally available in tIt is coloured blue, has a specific gravity of approximately 0.72, aEngineering Research and Development (DERD) specification 24100/130, the 100 corresponding to the anti-knock qualities whento the anti-knock qualities when a rich mixture is used.

3. There are two other categories of AVGAS which may begreen and has identical anti-knock characteristics to 100 LL (10conditions) but obviously contains more tetra-ethyl-lead (TEconsidered to be environmentally unfriendly. Avgas 80 is colourating of 80/87.



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s use implies limitations. It is moren fuel lines and pumps, especially atknock qualities of AVGAS and mayes. Detergents used in MOGAS maynd in some grades of MOGAS, will the obvious hazard to flight safety.

ve the use of MOGAS, using it may in certain light aircraft/enginerce of supply, storage and operatingr lock, the temperature of fuel in the flown above 6,000 feet. Because of

arburettor hot air systems must beater than with AVGAS.

engined aircraft may be categorisedls. Kerosene type fuel has a higherene has an extremely low vapour

point and higher volatility thana wider range of temperature and

Fuel & Fuel Systems


Automobile Gasoline (MOGAS)4. Automobile gasoline (MOGAS) may be available, but itvolatile than AVGAS and thus more prone to cause vapour lock ihigh temperatures and/or altitudes. It does not possess the anti-lead to pre-ignition and detonation in some aircraft piston enginbe harmful to aircraft fuel system components and alcohol, foucause deterioration of the seals in most aircraft fuel systems, with

5. Most major aircraft manufacturers have refused to approinvalidate the engine manufacturer's guarantee. It is usedcombinations, but limited by conditions of fuel specification, souprecautions. In particular, in order to minimise the risk of vapouaircraft tanks must be less than 20°C and the aircraft must not bethe higher volatility of the fuel (and possible water content) cfunctional and operated, since the risk of carburettor icing is gre

Turbine Fuel (AVTUR)6. Turbine fuels (Avtur) for use in turbo-prop and turbo-jet into two types. These are Kerosene and wide-cut gasoline fuedensity but slightly lower heating value than gasoline. Kerospressure, resulting in neglible fuel boiling and vapour losses.

7. An advantage of wide-cut gasoline is its low freezingkerosene and its ability to produce flammable mixtures over altitude conditions.



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colourless or straw coloured. The

ound installations are marked withsupplies. The labelling and colour

It is also known as ‘Civil Aviationhas a specific gravity of 0.807 at of -40°C.

es. A kerosene fuel (again with no A. JET A-1 has a specific gravity ofg limit of -47°C and must comply with fungus suppressant and icingfication 2453.

asoline, known as a wide-cut fuel. low flash point. Primarily used by specification of 2486. Because of

er fire hazard than does JET A-1.n partly-filled fuel tanks and whichapour. JET B has a lower SG thanl control unit (FCU) is necessary.

Fuel & Fuel Systems


8. Turbine fuel (Avtur) is not colour coded and is naturallycommonly used types of AVTUR are:

Colour Coding and Labelling9. Aircraft fuelling points, re-fuelling vehicles, pipes and grdistinctive colour codes to distinguish gasoline and kerosene coding used is illustrated at Figure 6-1.

JET A: Primarily available for commercial use in the USA. Kerosene’ and contains no gasoline blend. Jet A +15.5°C and has a maximum specified freezing limit

JET A-1: The fuel most commonly used by international airlingasoline blend) with a lower freezing point than JET0.807 at +15.5°C, has a maximum specified freezinwith DERD specifications 2494. It is also availableinhibitor (FSII) additives, when it meets DERD speci

JET B: A blend of approximately 30% kerosene and 70% gIt has a very low freezing point of -60°C but also aUS (and some other) military aircraft. It has a DERDits greater volatility it presents a considerably greatThis applies particularly to the vapour which forms iis more flammable than either JET A-1 or AVGAS vJET A-1 and when used adjustment of the engine fue



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FIGUFuel LaColou

rawn from each of the aircraft tank considered unfit for aircraft use if

Fuel & Fuel Systems


RE 6-1belling and

r Coding

Fuel Sampling10. Before flight each day a small quantity of fuel should be ddrain valves and inspected in a glass container. Fuel should bevisual inspection shows:



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per or chemical detector.

ment of aircraft fuels:

d and can vary in appearance from

es of the sampling vessel, or as bulk

ear hazed or cloudy.

e fuel or settle on the bottom of the

ight. Clear refers to the absence ofarance of fuel which is free from

o obvious separation of fluids, testty the sample onto a smooth surfaceld be all water!

Fuel & Fuel Systems


(a) More than a trace of sediment.

(b) Globules of water.

(c) Cloudiness.

(d) A positive reaction to water finding paste, or a pa

11. The following should serve as a guide to the visual assess

Colour. AVGAS 100LL is dyed blue, AVTUR JET A-1 is undyecolourless to straw yellow.

Undissolved water (free water) will appear as droplets on the sidwater in the bottom.

Suspended water (water in suspension) will cause the fuel to app

Solid matter (rust, sand, dust, scale, etc) may be suspended in thsampling vessel.

Clear and Bright. The fuel sample should appear clear and brsediment or emulsion and bright refers to the sparkling appecloudiness or haze.

12. Because of its lower SG, fuel floats on water. If there is nthe sample by smell, colour and finally by evaporation test. Empand see how quickly it evaporates. The fluid in the container cou



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e regularly checked for fuel quality.k is used for direct checking. Freshlowed to rest on the bottom of the

d regularly. The size of fuel sample a full and conclusive check of theal Gallon (approximately 4.5 litres).lies by means of a thief pump. Theseess and general cleanliness.

sy the overwing method or pressurenozzle of the refuel hose into the

transferred under low pressure fromst jet transport aircraft Fuel underser via a coupling at a refuel panels refuel valve selection switches andmade at the refuel panel so that the tank becomes full, a float operated

Fuel & Fuel Systems


Stored Fuel Quality Control13. Refuelling vehicles and bulk fuel storage facilities must bWater-finding paste (or tip tape) applied to the end of a dipsticpaste (or tape) must be used for each check and the dipstick alcontainer for a short period, not exceeding 10 seconds.

14. More extensive quality control checks are also conductetaken from the storage facility should be sufficient to completestate of the fuel and should typically be approximately 1 ImperiFuel samples are drawn up from buried tanks and barrelled suppsamples are checked for colour, sediment, water globules, cloudin

Refuelling

Precautions During Fuelling Operation15. Refuelling. Refuelling of aircraft is carried out either brefuelling. The overwing method involves placing the refuel overwing refuel opening in the top of the tank. The fuel is then bowser to tank. The pressure refuelling method is used on mopressure and high flow rate is transferred from bowser/dispenusually located under one wing of the aircraft. The panel housefuel tank gauges. Pre-set contents or full tank selections can be operator can select which tanks are to be replenished. When ashut off valve is closed by the rising fuel level.



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be taken during aircraft fuelling

main aircraft engines should not beg.

nce with the appropriate guidelines.

hat:

ot obstructed.

rom the aircraft in an emergency.

portions of the aircraft in the event

ipment and the aircraft wing as fuel

arture.

s should not be driven or parked

ld be positioned so that the aircraftent.

or which has an exhaust efflux dis- fuelling operation it should not be

risk of igniting fuel vapours.

Fuel & Fuel Systems


16. The following general precautionary measures shouldoperations:

• The main aircraft engine(s) should not be operated. The used to power the aircraft electrical systems during fuellin

• Bonding, as appropriate, should be carried out in accorda

17. Fuelling vehicles and equipment should be positioned so t

• Access to aircraft for rescue and fire fighting vehicles is n

• A clear route is maintained to allow their rapid removal f

• They do not obstruct the evacuation routes from occupiedof fire including chute deployment areas.

• Sufficient clearance is maintained between the fuelling equis transferred.

• They are not positioned beneath the wing vents.

• There is no requirement for vehicles to reverse before dep

• All other vehicles performing aircraft servicing functionunder aircraft wings while fuelling is in progress.

• All ground equipment such as rostrums, steps etc., shousettling under the fuel load will not impinge on the equipm

• If an auxiliary power unit located within the fuelling zonecharging into the zone is stopped for any reason during arestarted until the flow of fuel has ceased and there is no



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or should battery chargers be con-

er generators and the use of batteryng process within the fuelling zonee intrinsically safe type, should be made prior to the start of fuelling

g has ceased.

ion should be carried out in the fuel-

issue guidance, depending on localnded due to the proximity of severe

e operated.

the exchange of units should be use or testing of such equipment

fuelling their route should avoid theficial. The use of personal hand helde ‘NO SMOKING’ rule should be

Fuel & Fuel Systems


• Aircraft batteries should not be installed or removed nnected, operated or disconnected.

• The practice of connecting and disconnecting ground powtrolleys to supply power to an aircraft during the fuellishould be prohibited. No aircraft switches, unless of thoperated during this time. However, connections may beand the circuit should then remain unbroken until fuellin

• No maintenance work which may create a source of ignitling zone.

• Oxygen systems should not be replenished.

• The Aerodrome Authority - Air Traffic Control should conditions, as to when fuelling operations should be suspeelectrical storms.

• Aircraft external lighting and strobe systems should not b

• Aircraft combustion heaters should not be used.

• Only checking and limited maintenance work such asallowed on radio, radar and electrical equipment. Anyshould be deferred until fuelling is completed.

• When passengers are embarking or disembarking during fuelling zone and be under the supervision of an airline oftelephones by passengers should not be permitted. Thstrictly enforced during such passenger movements.



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Boardine operators of fixed wing aircraftn board during fuelling operations

20 should not be permitted to be

nvolved and the fuel being supplied that passengers should disembarkisembark when AVGAS is involved.

should be warned that fuelling willtrical equipment or other potential

on together with sufficient interiorlighting should remain on until fuel-lts’ signs should be switched off and

ssenger doors, (or the main passen-oor is available), and preferably atpassengers in the event of an emer-ld be constantly manned by a cabin

uch as catering and cleaning shouldazard or obstruct exits.

Fuel & Fuel Systems


Fuelling Operations with Passengers on• To reduce turnround time and for security reasons, airl

may allow passengers to embark, disembark or remain oprovided the following safety procedures are followed.

• Fixed wing aircraft with a seating capacity of less thanfuelled with passengers on board.

• When wide cut turbine fuels (e.g. Jet B, JP4, Avtag) are idoes not contain an anti-static additive, it is advisablebefore fuelling. Passengers should always be required to d

• Cabin attendants, passengers and other responsible stafftake place and that they must not smoke, operate elecsources of ignition.

• The aircraft illuminated ‘NO SMOKING’ signs should belighting to enable emergency exists to be identified. Such ling operations have been completed. The ‘Fasten Seat Bepassengers should be briefed to unfasten their seat belts.

• Provision should be made, via at least two of the main pager door plus one emergency exit when only one main dopposing ends of the aircraft, for the safe evacuation of gency. Throughout the fuelling operation these doors shouattendant.

• Ground servicing activities and work within the aircraft, sbe conducted in such a manner that they do not create a h



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eas and exit access areas should be

ted to meet the above requirements,ent area should be kept clear of allccordingly.

deployed should also be kept clear.

ircraft Equipped

of aircraft steps need beoor should be mannedred for immediate use astable chute. Where slidet to the aircraft e.g. girtg process.

available for use one set of aircraftsenger door normally used for the

Fuel & Fuel Systems


• Inside the aircraft cabin the aisles, cross aisles, all exit arkept clear of all obstructions.

• Whenever an exit with an inflatable escape slide is designathe ground area beneath that exit and the side deploymexternal obstructions and the fuelling overseer informed a

• Access to and egress from areas where other slides may be

Wide Bodied Aircraft And All Other AWith Automatic Infalatable Chutes

NOTE:

When a loading bridge is in use no additional sets provided. However, either the left or right rear dconstantly by a cabin attendant and should be prepaan emergency escape route using the automatic inflaactuation requires the manual fitting of an attachmenbar, the slide should be engaged throughout the fuellin

18. As a precautionary measure when a loading bridge is NOTpassenger steps should be positioned at the opened main pasembarkation and/or disembarkation of passengers.



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c Inflatable Chutes

eps should be positionedly at the opposing end of

aircraft passenger stepsoors (i.e. preferably one

these are deployed, each

area on the ground in the vicinity of leading edge of the wing) should bem this door to be deployed.

nd to ensure aisles and emergency

Fuel & Fuel Systems


Aircraft Not Equipped With Automati

NOTE:

When a loading bridge is in use, one set of aircraft stat another opened main passenger door and preferabthe aircraft.When a loading bridge is NOT available for use, should be positioned at two of the main passenger dforward and one aft) which are to be open.Where aircraft are fitted with integral stairways and may count as one means of egress.

Concorde Aircraft19. Whether passenger steps or a loading bridge is in use, the the left hand centre passenger door (immediately forward of theunobstructed by ground equipment to allow the escape chute fro

Cabin Attendants20. Cabin attendants are required to supervise passengers adoors are unobstructed.



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ing aircraft fuelling with passengershe aircraft to secure the rapid safeg the minimum number of cabinthe number of passengers on boardcraft and the emergency exits and be on board:

f) on the aircraft, with at least one

in the aircraft interior or any otherat he/she has adequate means of until, in the opinion of the Fuelling

ocations throughout the

uel must not be returned to aircraftn from an aircraft's tanks may beto the fuel specification normally

Fuel & Fuel Systems


21. The aircraft operator should ensure that at all times duron board, that there are sufficient cabin attendants on board tevacuation of passengers if an incident occurs. In determininattendants required, aircraft operators are to take into account the aircraft, their location within the cabin, the size of the airescape facilities. The following minimum cabin attendants should

22. 1 cabin attendant for every 50 seats (or fraction thereocabin attendant for each separate passenger compartment.

23. If during fuelling, the presence of fuel vapour is detectedhazard arises, the Fuelling Overseer (who should ensure thcommunication) should be informed. Fuelling should be stoppedOverseer, it is safe to resume.

NOTE:

These precautions may not be observed at some lworld. Knowledge of local precautions is necessary.

De-fuelling24. When an aircraft is de-fuelled, in whole or in part, that ftanks unless satisfactory quality checks are obtained. Fuel takecontaminated with water or sediment and may not conform provided at that aerodrome.



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le or storage tank segregated fromhe aircraft operator and its disposal

el said to have a good anti-knockquality is good, compared with

conditions.

l produces sufficient vapour to be

fuel particles (crystals) disappear

r will ignite in air at atmosphericnition source is not present. Also.

ion, or resistance to motion. Thefluid will flow or pour.

hange from liquid to vapour). Anlatility high enough to permit easy readily form excessive vapour in

Fuel & Fuel Systems


25. Consequently, de-fuelling should be into an empty vehicother parts of the installation. The fuel remains the property of tis his responsibility.

Definitions

Anti-Knock Value A fuel's resistance to detonation. A fuvalue when its detonation resisting other fuels under the same operating

Flash Point The lowest temperature at which fueignited by a small flame or spark.

Freeze Point The temperature at which visible solidon warming.

Spontaneous IgnitionTemperature

The temperature at which fuel vapoupressure, even though an external igknown as ‘auto ignition temperature’

Viscosity A measure of a fluid's internal frictlower the viscosity, the more freely a

Volatility The tendency of a fuel to vaporise (cideal aircraft fuel is one which has voengine starting, but not so high as tothe fuel system.



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em are considered.

rigid. Integral tanks are formed byxible tanks are manufactured to fit

etimes used for extra fuel storage inally.

cture during manufacture. In largee front and rear spars and the upperel. The fuel-tight seal is achieved byating the whole of the tank interiortre section is used as an integral fuelheads (baffles) are frequently fittedurge in the tanks during aircraft

ic sheeting and they are attached bybay. Wing fuel tanks in light aircraft

t aircraft, mounted in the fuselage.

Fuel & Fuel Systems


Fuel Systems26. In this section the major components of a typical fuel syst

Fuel Tanks27. Fuel tanks are found in three types; integral, flexible andsealing part of the aircraft structure (usually wings), whereas fleinto an available space in wing or fuselage. Rigid tanks are somthe aircraft fuselage or, in military aircraft, may be carried extern

28. Integral tanks are incorporated in the aircraft wing struaircraft the whole of the wing torque box (the ‘box’ formed by thand lower wing skins) is sealed and compartmented to contain fuinserting a sealant compound between mating surfaces and by cowith a fuel-proof coating. In light aircraft often only the wing centank, many do not incorporate integral tanks at all. Semi bulkinternally within large fuel tanks in order to prevent fuel smanoeuvres.

29. Flexible tanks are made from re-inforced rubber or plastmeans of cords or buttons to the structure surrounding the tank are often of the flexible type.

30. Rigid tanks are sometimes used for fuel storage in lighThey are generally constructed of aluminium alloy.



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eed, fuel tank is often adequate. Assential system requirements. Low-

nk to engine. The tank, or tanks, arehe centre of gravity as possible.

ntained in the wing structure (theented, typically on the basis of one

return valve) in order to maintainrtial vacuum forming in the tank asthe lowest point in the tank, in the

rporate vent surge tanks. They aree fuel tank venting system. Surge

flow and prevent spillage, especially fitted with a pump to transfer the

ves to prevent any excessive positivee systems incorporate positive andnegative pressure in the vent system

craft is shown at Figure 6-2.

Fuel & Fuel Systems


31. In single-engined high winged aircraft a single, gravity-ffilter and shut-off valve between tank and engine are the only ewing monoplanes will require a fuel pump to supply fuel from tausually mounted at the aircraft's longitudinal centreline as near t

32. Multi-engined aircraft normally have the fuel tanks cofuselage being required for payload). The tanks will be compartmtank per engine plus a centreline tank.

33. Fuel tanks must be vented to atmosphere (via a non-ambient atmospheric pressure inside the tank. This prevents a pafuel is used. Provision is made for draining off fuel/water from form of a sump drain.

34. The fuel tank system on large jet transport aircraft incoconstructed the same as integral fuel tanks and form part of thtanks are normally empty and are designed to contain fuel overduring refuelling to full tank loads. Some vent surge tanks areaccumulated fuel back into a main tank.

35. Fuel tank vent systems also contain overpressure relief valpressure build up in the vents during ram air venting. Somnegative relief valves to prevent wing damage due to positive or at all aircraft attitudes.

Fuel Feed Systems36. A simple fuel system of the type found in modern light air



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FIGUSimpleSystem

ly to the engine-driven fuel pumpch and landing. The booster pumprsed at the bottom of the fuel tank. a booster pump.

ent a partial vacuum forming in thew from the tanks would eventuallyhich are to be parked in the open forg in the tanks, but could lead to fuel.

Fuel & Fuel Systems


RE 6-2 Fuel

37. The booster pump serves to provide positive fuel suppduring critical flight phases such as take-off and climb, approamay be fitted in the fuel feed line, as shown, or it may be immeSome high-winged light aircraft rely upon gravity feed instead of

38. The fuel tanks are vented to atmosphere in order to prevtanks as fuel is drawn off. Were this allowed to occur, fuel flocease. It is good practice to completely fill the tanks of aircraft wany length of time. This reduces the risk of condensation forminventing overboard when ambient temperatures subsequently rise



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systems to accommodate re-fuelling some cases, the ability to jettisonel system for a four-engined aircraft

h can supply the needs of any one low pressure (LP) fuel pump willay have to be reduced in order toctrically driven, usually by an ACually operated by a DC motor.

Fuel & Fuel Systems


39. Multi-engined aircraft require more complex fuel supply and tank venting in addition to provision for cross-feeding. In(dump) fuel in flight may be necessary. The starboard half of a fuis shown at Figure 6-3.

40. Each tank contains two booster pumps, either of whicengine. In the event of failure of both pumps the engine-drivencontinue to supply the engine, but aircraft operating altitude mprevent cavitation in the feed line. The booster pumps are eleinduction motor, whereas in light aircraft the booster pump is us



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FIGUFuel CSystem

ss-feed and inter-engine valves, then tank (cross-feed and inter-engines from starboard tanks (cross-feeds (cross-feed and inter-engine valves

f the tank, to avoid water residueof fuel at the bottom of the tanksl.

Fuel & Fuel Systems


RE 6-3ross-feed

41. From Figure 6-4 it can be seen that, by use of the crosystem can be operated so that each engine draws from its owvalves closed), port engines from port tanks, starboard enginevalve closed, inter-engine valves open) or all engines from all tankopen).

42. The booster pumps draw fuel from above the bottom oreaching the engines. Consequently there is always a quantity which cannot be pumped and this forms part of the unusable fue



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uel in a flight emergency, in order to provision for this. Typically, the re-p, protected by a jettison valve. One, and it is this pump which is usedty of fuel remains in the tanks. Suchining replaces the use of the booster

pipe in each wing, the pipe being short manifold is fitted between thefrom each tank into the manifold.fer valves, the transfer pumps being. When the jettison pipe is in the master jettison valve. The circuitse is locked in the extended position.

nditions under which fuel may beitions it is only safe to jettison fuelexpressed authorization of ATC. Asd vaporises, hence it presents a firer electrical storm activity. In aircraftted to prevent any possibility of fuel

Fuel & Fuel Systems


Fuel Jettisoning43. In large transport aircraft it may be necessary to jettison freduce all-up weight for landing. Some aircraft are equipped withfuelling gallery is extended to an open-ended pipe near the wingtiof the booster pumps in each tank draws fuel from a standpipewhen fuel jettison is selected. This ensures that a specified quantia system is illustrated at Figure 6-3. In some aircraft gravity-drapumps during fuel jettisoning.

44. In another type of system, fuel is jettisoned through alowered into the airstream by an electrically-operated actuator. Amain tanks in each wing, and a jettison valve controls flow Auxiliary tanks are fed into the main tanks by the normal transinter-connected with the circuits operating the jettison valvesretracted position it forms a seal at the manifold, and acts as acontrolling the jettison valves are not armed until the jettison pip

45. National legislation stipulates the aircraft operating cojettisoned. Below certain altitudes and in certain climatic condover the sea and fuel dumping may only be carried out with the jettisoned fuel escapes from the aircraft it quickly breaks up anhazard. To reduce this, jettisoning should never be attempted neawith underwing engines, dumping is carried out with flaps retracvapour being directed into the exhaust gas stream.



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ttisoning the fuel in the tanks belownd thereafter allowing 45 minutes

ne unless it can be shown that the weight less the fuel necessary for a at the airport of departure with thethat used in meeting the applicable

order to maintain supply of fuel toentre of gravity or undue stresses onth two tanks in each wing (inboard

trating weight in the outer wings to, however, the centre of gravity willks is burned off, the landing weightuel to meet the various criteria is a

Fuel & Fuel Systems


46. A means must be provided in the fuel system to prevent jea level that would permit a climb from sea-level to 10,000ft acruise at a speed for maximum range.

47. A fuel jettison system must be installed on each aeroplaaeroplane meets a climb gradient of 3.2% at maximum take-off15 minute flight comprised of a take-off, go-around and landingaeroplane configuration, speed, power and thrust the same as take-off, approach and landing climb performance requirements.

Fuel System Monitoring48. It is important that the fuel system is properly managed inthe engines, whilst avoiding excessive movement of the aircraft cthe wings. Consider, for example, a large swept-wing aircraft wiand outboard) and a fifth, centre-section tank.

49. At high all-up weights there is structural benefit in concenreduce the bending moment of high wing loads. As fuel is usedshift because of wing sweep. If all the usable fuel in the wing tandistribution will be wrong. Consequently, management of the fcomplex business.



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dications are grouped together on abe associated with the fuel system atmps, engine LP cocks and the valvesrate are indicated on gauges, withlure). Fuel tank contents are usuallyvolume. This is convenient since theal to its weight (mass).

is with the cross-feed closed and fuelank (CWT) in addition to the maine-off and empty it before switching lateral imbalance occurs, or to feed configuration, especially in a twin-

Fuel & Fuel Systems


50. To facilitate fuel system management the controls and incontrol panel. Figure 6-4 illustrates a control panel which would Figure 6-3. It will be seen that this provides control of booster puin the cross-feed system. Fuel temperature, pressure and flow warning lamp indication of low fuel pressure (booster pump faimeasured by a capacitive system in terms of weight, rather than energy available in the fuel (calorific value) is directly proportion

51. Normal operation of a multi-engined aircraft fuel system feed from tank to engine. In aircraft fitted with a centre wing twing tanks it is normal practice to switch to this tank after takback to the main tanks. The cross-feed is used to transfer fuel if aan engine from other than its ‘own’ fuel tank. The ‘engine out’engined aircraft, would be with cross-feed open.



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FIGUFuel SyContro

illary) tank(s) will require the centrebe consumed first.

Fuel & Fuel Systems


RE 6-4stem l Panel

Example of Fuel Management52. A four engined aircraft with wing and centre section (auxsection tank contents (except for any fuel required as ballast) to



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eed. When established in the climb,ank, backed up by booster pumps in remain off.

utput pressure than the wing tanknter-engine valves selected open, all

r pumps will continue to supply all pump low pressure warning light

same as the outboard tanks, thene valves closed. The fuel system isflight, unless uneven fuel burn or an

l tanks, one booster pump forward,sed. During descent with a low fuele forward booster pumps should bet booster pumps should be used, as pump inlets.

Fuel & Fuel Systems


53. For take-off it is normal to select direct tank-to-engine fthe fuel system will be configured to supply fuel from the centre tthe inboard wing tanks. The outboard wing tank booster pumps

54. By arranging the centre tank pumps to have a higher obooster pumps, fuel will flow from the centre tank. With the iengines will be fed from the centre tank.

55. When the centre tank is empty, the inboard tank boosteengines, and the centre tank pumps can be turned off. A tankilluminates for each pump when the tank is empty.

56. When the quantity of fuel in the inboard tanks is theoutboard tank booster pumps are turned on and the inter enginow configured for tank-to-engine feed for the remainder of the engine failure results in fuel load asymmetry.

57. The booster pumps are located at the low point of the fueone aft. During take-off and climb, the aft booster pumps are uquantity, the rear booster pump inlets may be uncovered and thused. In the event of a go-around at a low fuel quantity, the afaircraft acceleration and climb may uncover the forward booster



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e it passes through the low pressurenot enter the filters as ice and causeles begin to form in kerosene (Jet A)ing fuel filters and again heating the

l/fuel heat exchangers in which heatalternative, and in some instancesthe heating medium is compressorgh manual control of air/fuel heat

arning lamp, activated by a pressurelter is restricted by ice or wax thet to operate the icing warning lamp.r bleed air to heat the fuel until the

tically-controlled systems are fullyve 0°C.

Fuel & Fuel Systems


Fuel Heating58. It is usual in large aircraft for the fuel to be heated beforfilters. This is to ensure that any droplets of water in the fuel do blockage of the fuel supply to the engines. Solid, wax-like particat temperatures below -40°C. These particles are capable of cloggfuel prevents this.

59. In many cases the fuel is heated by passing it through oifrom the engine lubricating oil is transferred to the fuel. An additional, method employs air/fuel heat exchangers in which bleed air. Both are usually thermostatically controlled, althouexchangers is sometimes employed.

60. Manually controlled fuel heating systems incorporate a wdifferential switch in the fuel filter. If fuel flow through the fidifferential pressure across the filter increases until it is sufficienThe heat exchanger valve is then opened to admit hot compressofuel temperature gauge reaches a limiting value. Thermostaautomatic, maintaining fuel temperature within pre-set limits abo



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ms

AC Electricity section. A capacitored by a material resistive to currentity supplied to it, and will discharge is removed. The magnitude of thee dielectric strength of the material

pacitive probe which is immersed inth one tube inside the other. These form the conducting ‘plates’ of aitor alternately stores an electricalde of the discharge is determined bys the fuel in the tank or, if the tank ishe tank rises or falls the dielectricmore than twice that of air so thein proportion to the fuel level in the produces a signal whose strength is

ge circuit to provide an amplifiediple of a bridge circuit is explained

Fuel & Fuel Systems


Fuel Quantity Measurement Syste

Capacitive61. The principle of electrical capacitance is described in theconsists of two conducting plates in an electrical circuit, separatflow called the dielectric. The device will store a charge of electricthat stored charge back into the circuit if the charging currentstored charge and discharge (the capacitance), depends upon thwhich separates the conducting plates of the capacitor.

62. Capacitive fuel quantity measurement systems utilize a cathe fuel tank. The probe consists of two conducting tubes, wiextend from the top of the fuel tank to the bottom and theycapacitor. When supplied with alternating current, the capaccharge and then discharges it back into the circuit. The magnituthe dielectric strength of the separator. In this case the dielectric iempty, the air in the tank. Consequently, as the fuel level in tstrength of the capacitor varies. Fuel has a dielectric strength capacitance, or charge-storing capability, of the capacitor varies tank. The magnitude of the capacitor discharge is measured andproportional to tank contents.

63. The output from the capacitive probe is used in a bridsignal, which actuates the tank fuel contents gauge. The princshortly in the description of resistive fuel quantity measurement.



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, rather than volume. The probes correction for fuel temperature andtype (Jet A-1 or Jet B). This isn the weight of fuel, not its volumence depends upon the mass of that

e connected in parallel in each fuelto aircraft attitude change and wingrallel circuit carrying a continuous amplified tank probe signal. In thellel circuit drives the tank contents

n the indicator circuit, again causing functioning correctly. Alternatively,ead a given value, as specified in the

s are of the resistive type. Thesected in a wheatstone bridge circuitindicator gauge. The principle is

Fuel & Fuel Systems


64. Capacitance systems measure fuel quantity by weightmeasure the volume of fuel in the tank, a compensator applies aa densitometer adjusts for different fuel specific gravity or advantageous because the fuel energy available is dependent upo(the number of molecules present in a given quantity of a substaquantity, regardless of the volume occupied by it).

65. In practice a number of capacitive probes, or sensors, artank in order to compensate for fuel movement in the tank due flexure. Each fuel tank indicating system is provided with a pa‘empty’ signal. In normal circumstances this is suppressed by theevent of failure of the indicating signal, the current in the paragauge to the ‘empty’ position.

66. Operation of the TEST circuit simulates an empty signal ithe tank contents gauge to read empty if the indicating circuit ispressing the test button may cause each tank contents gauge to ramplified checklist and the manufacturer's system description.

Resistive67. Many light aircraft fuel tank quantity indicating systememploy a float-operated variable resistance in the tank, connecontaining reference fixed resistances and the fuel quantity illustrated at Figure 6-5.



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FIGUWheatBridgeQuant

Fuel & Fuel Systems


RE 6-5stone Fuel ity Indicator



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f the same resistive value (say 100ine by Ohm's law that the current

DC) will be the same. Under theser and it will read zero (tank empty).perated variable resistance increasesits resistance has risen to 200 ohms.s the variable resistance) will be lesse between points B and D such thatsing Ohm's Law, to show that theurrent to flow from B to D, throughding under these circumstances.

ctly indicate tank contents with theon provided is obviously inaccurate.the fuel due to temperature changesfuel weight rather than volume.

ted or to obtain the contents in the be used. These may be dipsticks,

in the top of the tank and when

Fuel & Fuel Systems


68. The three fixed resistances (R1, R2 and R3) are each oohms). If the variable resistance is also 100 ohms, we can determflow through each side of the bridge circuit (sides ABC and Acircumstances there will be no current flow through the indicatoAs the level of fuel in the tank rises and lifts the float, the float-oits resistive value. Suppose, for example, at maximum fuel level The current flow through side ABC of the bridge (which containthan that through side ADC. This will create a voltage differencthe voltage at B will be higher than that at D (it is possible, uvoltage difference in this instance would be 4v). This will cause cthe indicator, which would be calibrated to show a ‘tank full’ rea

69. Resistive fuel quantity indicators are calibrated to correaircraft in a level attitude. At other aircraft attitudes the indicatiSince they indicate tank fuel level, expansion or contraction of will result in erroneous readings if the indicator is calibrated for

Fuel Measuring Sticks70. In order to confirm that a quantity of fuel has been uplifevent of indication system failure, fuel measuring sticks maydripsticks or magnetic level indicators.

71. The dipstick is simply a calibrated stick which is fittedremoved will indicate by wetness, the level of fuel in the tank.



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ottom of the tank and when pulled the tube and exits at the drip hole.

FIGUFuel LeDripst

Fuel & Fuel Systems


72. The dripstick is a calibrated tube which is fitted in the bdown, will indicate fuel level in the tank when fuel flows through(Figure 6-6).

RE 6-6vel

ick



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g inside the tank surrounded by aagnet at the upper end. When the

ic coupling occurs between the rod calibrations on the indicator rod.

, the stick readings are converted tor fuel specific gravity. The aircraft/nose down, pitch and roll attitude.

Fuel & Fuel Systems


73. The magnetic level indicator system consists of a housinfloat containing a magnet. The calibrated indicator rod has a mindicator rod is lowered out of the bottom of the tank, magnetand float. The level of fuel in the tank may be read off from(Figure 6-7).

74. Since all the fuel measuring sticks indicate volume of fuelfuel weight using tables or charts and applying a connection foattitude is also required at the time of stick readings i.e. nose upAn aircraft spirit level indicator is provided for this purpose.



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FIGUMagneIndicat

Fuel & Fuel Systems


RE 6-7tic Level or



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n engine aircraft.

r source.

Fuel & Fuel Systems



QUESTIONS:QUESTION 1.

What are the main components in a pneumatic system for a pisto

QUESTION 2.

Identify the aircraft systems which use compressed air as a powe

QUESTION 3.

Identify the main components of an air conditioning system.

QUESTION 4.

What heating sources are used in air conditioning systems.

QUESTION 5.

How is cabin temperature controlled.

QUESTION 6.

Identify the main components of a pressurisation system.

QUESTION 7.

What is cabin altitude.



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erating altitude.

ntrol system.

es in pressurisation system.

Fuel & Fuel Systems


QUESTION 8.

What is differential pressure.

QUESTION 9.

What is the maximum cabin altitude under normal conditions.

QUESTION 10.

How does the maximum differential pressure limit maximum op

QUESTION 11.

What areas of the fuselage are pressurised.

QUESTION 12.

What is the principle of operation of a cabin pressure altitude co

QUESTION 13.

Identify and explain the principle of operation of the safety devic

QUESTION 14.

Where on the aircraft are pneumatic de-icing systems located.

QUESTION 15.

Describe the working principle of inflatable rubber boots.



Stacey Garland

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ntrolled.

Fuel & Fuel Systems


QUESTION 16.

In a pneumatic de-icing system, how is inflation and deflation co

QUESTION 17.

When should the pneumatic de-icing system be operated.

QUESTION 18.

List the power sources which drive pneumatic systems.

QUESTION 19.

What is the purpose of a pneumatic system.

QUESTION 20.

In a pneumatic system what is the function of an isolation valve.

QUESTION 21.

What is the function of the pressure regulating valve.

QUESTION 22.

What is the function of the system bleed valve.

QUESTION 23.

State how the pneumatic system is controlled and monitored.



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system.

Fuel & Fuel Systems


QUESTION 24.

What are two possible pneumatic duct failures that can occur.

QUESTION 25.

What is the purpose of an air conditioning system.

QUESTION 26.

Explain how air temperature is controlled.

QUESTION 27.

What is the function of the ram air valve in the air conditioning

QUESTION 28.

What is the function of the re-circulation fan.

QUESTION 29.

Describe the use of hot trim air.

QUESTION 30.

Describe the construction of the 3 types of fuel tanks.

QUESTION 31.

What is the function of the fuel tank baffles.



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ially filled.

lve.

to the engine.

Fuel & Fuel Systems


QUESTION 32.

What is the function of an overpressure relief valve.

QUESTION 33.

What is the function of the surge vent tank.

QUESTION 34.

What is the refuelling sequence for tanks that are only to be part

QUESTION 35.

What is meant by unusable fuel.

QUESTION 36.

Describe the various methods of refuelling.

QUESTION 37.

What is the operating principle and function of a fuel screen.

QUESTION 38.

What is the operating principle and function of the cross feed va

QUESTION 39.

What method is used to control the temperature of the fuel feed



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Fuel & Fuel Systems


QUESTION 40.

Describe the typical controls of a fuel system.

ANSWERS:ANSWER 1.

Engine driven compressor.

Pressure regulator.

Main storage bottle.

Pressure reducing valve.

Oil and water trap.

Indication and warning.

ANSWER 2.

Landing gear, main and alternate.

Wheel Brakes.

Radiator shutters.

Aerofoil de-icing.



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ld air in response to the temperature

Fuel & Fuel Systems


ANSWER 3.

Ram air inlet.

Heat exchanger.

Refrigeration unit.

Water separator.

ANSWER 4.

Exhaust heater muff-Ram air.

Combustion heater-Ram air.

Engine driven blower.

ANSWER 5.

Cabin air temperature is controlled by mixing the hot air with cocontroller.



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to the prevailing cabin air pressure.

he cabin pressure compared to thepressure is higher than ambient and

imum value and because maximumaltitude will be limited.

Fuel & Fuel Systems


ANSWER 6.

Pressurisation controller.

Outflow (Discharge) valves.

Safety valves.

Inward relief valves.

Dump valve.

Ditching Control.

ANSWER 7.

Cabin altitude is the altitude in the pressure cabin corresponding

ANSWER 8.

Differential pressure is the difference in air pressure between taircraft ambient pressure. It is a positive value when the cabin negative value when the ambient is higher than cabin pressure.

ANSWER 9.

10,00ft

ANSWER 10.

The pressurisation controller will limit the differential to a maxcabin altitude is also limited it means that the aircraft operating



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automatic pressurisation controllerflow valve(s).

ximum design structural differential

bin should ambient pressure exceed

air and vacuum respectively which

s on a cyclic time sequence allowing

Fuel & Fuel Systems


ANSWER 11.

Passenger Cabin and Flight Deck

Cargo Holds

ANSWER 12.

Air entering the pressure hull is mass flow controlled and an governs the amount of air leaving the hull by positioning the out

ANSWER 13.

Safety valves. Relieve pressure from the pressure cabin when mapressure is reached. Inward relief valves. Equalise pressure between ambient and cacabin pressure.

ANSWER 14.

On the leading edges of the mainplane, fin and tailplane.

ANSWER 15.

Rubberised fabric tubes are inflated and deflated cyclically withwill break up the ice layers above them.

ANSWER 16.

Inflation and deflation is controlled by electrically operated valvepressure or vacuum into the inflatable tubes.



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s formed on the aerofoil surfaces.

.

Fuel & Fuel Systems


ANSWER 17.

It should be switched on when the recommended depth of ice ha

ANSWER 18.

Engine compressor bleed air

Engine driven blower

APU bleed air

Ground air supply

ANSWER 19.

Air conditioning and pressurisation.

Engine starting.

Thermal anti-icing.

Pressurising hydraulic reservoir and potable water supply.

Leading edge flaps.

Emergency lowering of landing gear.

ANSWER 20.

To shut off the engine bleed air supply from compressor to system



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umatic system.

ive once the appropriate bleed airitored by reference to duct pressurecating pneumatic duct leaks.

to vibration or overflexing causing

pressurisation at a comfortable

mixing hot and cold air in variablebed limits.

Fuel & Fuel Systems


ANSWER 21.

To control the designed pneumatic system pressure.

ANSWER 22.

To admit air from the appropriate compressor stage into the pne

ANSWER 23.

It is controlled by the pressure regulating valve which is actselection has been made on the flight deck. The system is mongauges on the pneumatic system display with warning lights indi

ANSWER 24.

OVER PRESSURE. This could cause duct leakage of hot air.MECHANICAL. Duct joints can become loose or detached duehot air to leak from the duct.

ANSWER 25.

To maintain an adequate supply of air for ventilation andtemperature and humidity.

ANSWER 26.

The temperature of the air supply to the cabin is controlled by proportions to maintain the cabin air temperature within prescri



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event that there is total loss of air

air back to the mix manifold, thus

hot trim air to the affected zone.

onstructed from re-inforced rubberhe aircraft structure. Can be in the

is fitted to relieve excess pressure. at all aircraft attitudes.

Fuel & Fuel Systems


ANSWER 27.

To provide a means of admitting ram air for ventilation in theconditioning.

ANSWER 28.

To reduce the air conditioning load, by recirculating some cabinsupplementing incoming fresh air.

ANSWER 29.

Individual zone temperature requirements are satisfied by adding

ANSWER 30.

Rigid-Generally constructed of Aluminium Alloy. B. Flexible-Cor plastic sheeting in a non rigid form. C. Integral-Built into tlower centre fuselage or commonly in the wing.

ANSWER 31.

To prevent fuel surging in the tank during aircraft manoeuvres.

ANSWER 32.

Where the area above a fuel tank is pressurised a relief valve Located in the vent system it will relieve any positive or negative



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back into the main tank system.

board tanks to the required levelng tanks.

oved by pumping. It is determined

ernal source by placing a fuel nozzlepressure from an external source is

er contamination of the fuel system.e pump.

the ability to maintain fuel balancegines.

Fuel & Fuel Systems


ANSWER 33.

To collect any fuel which flows into the vent system and pump it

ANSWER 34.

The sequence of refuelling is to commence refuelling the inaccording to requirements and continue outboard for all remaini

ANSWER 35.

A quantity of fuel at the bottom of the tanks which cannot be remby the shape or position of the tank.

ANSWER 36.

OVERWING. A method whereby a tank is refuelled from an extdirectly into the tank filler opening. PRESSURE. Fuel under pumped into the fuel tank via a connection at the refuel panel.

ANSWER 37.

A fuel screen is a filter which is required to prevent foreign mattEg. A fuel filter is fitted between the fuel tank and the inlet to th

ANSWER 38.

A cross feed valve will give flexibility of fuel management and between tanks. It allows any tank to supply fuel to any or all en



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e fuel is maintained at a minimum through a heat exchanger using a

ster Pump switching. Fuel heater

Fuel & Fuel Systems


ANSWER 39.

Prior to entering the main fuel filter and engine fuel control, thtemperature above zero degrees centigrade. The fuel is passedmedium of hot bleed air or oil to warm the fuel.

ANSWER 40.

Typical controls consist of; Fuel system shutoff valves. Booselection. Fuel tank transfer valves switching.





Electrics-DC

Basic Principles of DC Electricity

Magnetism and Electricity

Electro-Magnets

Relays

Solenoids

Electro-Magnetic Induction

Control and Protection

Monitoring Devices

Wheatstone Bridge

DC Motors

Electrical Consumers

Bonding and Screening



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vely charged nucleus and negativelyumber of protons and neutrons, therge. An oxygen atom, for example, eight neutrons (no charge). Hencee of the electrons is balanced by the

through a conductor. When matterains electrons it becomes negatively

re of electrical current flow was notily assumed to be from positive togard to current flow has persisted,lves the transference of negatively

which has a deficiency of electrons).ion of current flowing from positivemit the rate of current flow.

Electrics-DC


7Electrics-DC

Basic Principles of DC Electricity1. All matter is made up of atoms, which consist of a positicharged orbiting electrons. The nucleus is formed by an equal nprotons being positively charged and the neutrons having no chacomprises eight electrons (negative), eight protons (positive) andsuch an atom is electrically neutral, since the total negative chargtotal positive charge of the protons.

2. An electrical current is quite simply a flow of electrons gives up electrons it becomes positively charged, and when it gcharged.

3. During the early experiments with electricity the true natufully understood and the direction of current flow was arbitrarnegative, as illustrated at Figure 7-1. This convention with realthough it was subsequently proved that a current flow invocharged electrons to a positively charged atom (that is, an atom For the sake of conformity, these Notes will maintain the conventto negative, preferably through a load or resistance, which will li



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FIGUDirectCurren

electrons is called a conductor, andd Gold are both particularly goodare reluctant to release electrons isflow of electrical current. Insulatorsnductors. For example, electrical

they are to be located where other, might otherwise come into contact good insulators. Polyvinyl chloridehose resistance is midway between

only used semiconductor materialsd circuits.

Electrics-DC


RE 7-1ion of t Flow

4. A substance consisting of atoms that readily give up theirwill offer little resistance to a flow of electricity. Copper anconductors. Conversely, a substance consisting of atoms that called an insulator, because it offers a very high resistance to the are used to protect conductors, insulating them from other cocurrent-carrying cables are covered with insulating material if conductors, such as the metal case of a device or a human beingwith them. Natural rubber and most thermoplastic materials are(PVC) is a well-known example. A semiconductor is a material wthat of a good conductor and that of a good insulator. Comminclude silicon and germanium. (diodes, transistors and integrate



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duced into a body, which then storesdy. A good example of this occursrth for all of the aircraft electricalt is made to discharge this potentialround it is discharged to the tarmacsewheels).

nd an aircraft and a fuel bowser to the tank filler pipe. Were this notiously disastrous consequences. Fors of the aircraft structure to providects of a lightning strike.

ifferences, as in thermocouples; by cells; by pressure, as in carbon as in generators; and by chemical

of electrons through a conductingmeter. The lower the resistance toum current will flow when a short

and this causes an overload. Nos broken, for example by opening aite resistance.

Electrics-DC


5. Static electricity occurs when an excess of electrons are inthis electrical charge until able to discharge itself to another bowith an aircraft in flight. Since the airframe is used as the easystems it develops its own static electrical potential. An attempduring flight to atmosphere via static wick dischargers. On the gvia a conducting bead in the tread of the tyres (usually on the no

6. It is because of static electricity that it is necessary to boeach other during refuelling, and to bond the fuel pipe nozzle todone, static electricity could build up and cause a spark with obvthe same reasons it is necessary to bond together the various parta low resistance path for static discharge and to dissipate the effe

7. Electricity may be produced by heat or temperature dfriction, as in static electricity; by light, as in photo-electricmicrophones; by rotating mechanical forces in magnetic fields,action, as in batteries.

Definitions8. As already stated, an electric current is simply a flow element and it is measured in amperes (amps) by means of an amcurrent flow, the greater the current flow and vice versa. Maximcircuit exists (a direct connection between supply and return),current will flow when an open circuit exists (when the circuit iswitch). The break in the circuit has created a condition of infin



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nvenient in that they are either toorefixes are added. For example;

de of the surplus of electrons at onecy of electrons at another point is pressure, and is measured in volts

s to flow, and is effectively the sameolt. One volt is defined as the emfgh a resistance of one ohm.

omponent or electrical machine tos using an ohmmeter. The usual

generation of heat.

mples

rovolt, microamp

livolt, milliamp

watt, kilovolt

gohm, megawatt

Electrics-DC


9. The basic units of electrical measurement are often incosmall or too large for practical circumstances and consequently p

10. Potential difference (PD) is the way in which the magnitupoint in an electrical circuit, when compared with the deficienexpressed. Potential difference is often referred to as electricalusing a voltmeter.

11. Electromotive force (emf) is the force that causes electronas potential difference. The unit of measurement of emf is the vrequired to cause current to flow at the rate of one ampere throu

12. Resistance is the reluctance of any material, electrical cpermit the flow of electricity. Resistance is measured in ohmconsequence of a circuit offering resistance to current flow is the

Prefix Meaning Exa

Micro One millionth (10-6) Mic

Milli One thousandth (10-3) Mil

Kilo One thousand (103) Kilo

Mega One million (106) Me



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emperature and to conductor lengthonductor. As the temperature of a Any electron flow in the circuit willt is, increase in temperature causesf the conductor as the temperaturelled the temperature co-efficient of a conductor increases with increaseegative co-efficient if the resistance

e co-efficient (PTC) and negativems for measurement of temperature.ve resistor with either a positive or

cuit, for it to follow and there must The amount of current flow willment of current flow is the ampere,f current flow of one coulomb per

quantity. Current, or rate of flow of

rent flow (I) and resistance (R). Itortional to the voltage (emf) and

Electrics-DC


13. The resistance of a conductor is directly proportional to tand is inversely proportional to the cross-sectional area of the cconductor rises the atoms gain energy and they become ‘excited’.now experience difficulty in moving through the conductor. Thaan increase in conductor resistance. The change of resistance ochanges expressed as a fraction of its original resistance is caresistance referred to the original temperature. If the resistance ofin temperature it has a positive temperature co-efficient and a ndecreases with increase in temperature. Positive temperaturtemperature co-efficient (NTC) resistors are used in aircraft systeAn example of this is a thermistor which is a thermally sensitinegative temperature co-efficient.

14. For electrical current to flow there must be a path, or cirbe a potential difference (voltage/pressure) to cause it to flow.depend upon the resistance of the circuit. The unit of measureusually abbreviated to amp. One ampere is defined as a rate osecond, the coulomb being the unit of measurement of electrical electricity, is symbolised by the letter I.

Ohm's Law15. Ohm's Law gives the relationship between emf (V), curstates that the current in an electrical circuit is directly propinversely proportional to the resistance.



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mp to flow through a circuit having cause a current of 10 amps to flow

FIGUOhms

viz:

Electrics-DC


16. In other words, an emf of 1 volt will cause a current of 1 aa total resistance of 1 ohm. Equally, an emf of 25,000 volts willthrough a circuit having a total resistance of 2,500 ohms.

The equation for Ohm's Law is: I =

This is usefully given diagrammatically as shown at Figure 7-2.

RE 7-2Law

17. This simplifies transposition of the Ohm's Law equation

VR----

V I R×=

IVR----=

RVI----=



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other way, the rate of dissipation of amperage is the measure of rate ofe two. The unit of measurement of one volt moves a quantity of one

Law equation it is possible to derive

hen electrical energy is expended. expended and is measured in unitser in 1 second

) I× I2R=

---�V

2

R------=

Electrics-DC


18. Power is defined as the rate of doing work or, to put it anenergy. Since voltage is the measure of electromotive force andcurrent flow it follows that electrical power is the product of thpower is the watt and is defined as the power expended whencoulomb per second through a conductor.

1 Watt = 1 Volt x 1 Amp

19. By combining the power equation above with the Ohm's alternative power equations.

and

20. Electrical power is the rate at which work is done wElectrical work is a measure of the amount of electrical energycalled Joules. 1 Joule represents the work done by 1 watt of pow

Power (W) V I×=

Since V I R it follows that W× I R×(= =

Since IVR---- it follows that W V

VR-��×= =



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FIGUElectriWatts

istance of 6 ohms, thanks to an emfumed in the circuit is therefore:

Electrics-DC


RE 7-3cal Power =

21. At Figure 7-3 a current of 2 amps is flowing through a resproduced by the battery of 12 volts (V = I x R). The power cons

or

or

V I× 12 2× 24 Watts= =

I2R 4 6× 24 Watts= =

V2

R------

1446

--------- 24 Watts= =



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FIGUShort

its terminals by a conductor of lowcuit will be:

load)

Electrics-DC


RE 7-4Circuit

22. At Figure 7-4 the battery has been short-circuited acrossresistance (0.1 ohms). The current flowing through the short cir

IVR----

120.1------- 120 amps (an over= = =



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FIGUOpen

a switch, introducing an infinite

n only one load, or resistance, foreparate loads, such as a number of loads is to place them in series with

Electrics-DC


RE 7-5Circuit

23. At Figure 7-5 the circuit has been open-circuited byresistance. Current flow will now be:

Series Loads24. Up to this point the simple circuit diagrams have showsimplicity. Practical electrical circuits usually contain several slight bulbs, for example. One method of connecting a number ofone another, as shown at Figure 7-6.

IVR----

12∞------ 0 amps= = =



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FIGUSeries

must flow through each load ore sum of all the resistances:

a whole must equal the sum of the

olts

Electrics-DC


RE 7-6Loads

25. When loads are connected in series, since the currentresistance sequentially the total resistance (RT) of the circuit is th

26. Applying Ohm's Law to the circuit:

27. Taking this further, the voltage drop across the circuit asvoltage drops across each resistance. Viz:

RT R1 R2 R3+ +=

IVR----

126

------ 2 amps= = =

R1 1ohm( )V IR 2 1× 2 v= = =



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few practical applications since it is entire circuit. Similarly, if one loadntly such circuits are rarely used in

n which the loads are connected inircuit each load may be individuallyaining loads.

FIGUParalle

his circuit it is necessary to use the

olts

olts

Electrics-DC


28. Circuits in which the loads are connected in series have impossible to switch off supply to one load without affecting thefails causing an open circuit, the entire circuit fails. Consequeaircraft systems.

Parallel Loads29. Figure 7-7 shows a more conventional electrical circuit iparallel with each other and with the source of emf. In such a cdisconnected without disrupting the flow of electricity to the rem

RE 7-7l Loads

30. In order to determine the equivalent total resistance of tformula:

R2 2ohm( )V IR 2 2× 4 v= = =

R3 3ohm( )V IR 2 3× 6 v= = =

Total 12 volts



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the smallest individual resistance in

Ohm's Law:

gh each load in the normal way, andts flowing in each load.

n

s

Electrics-DC


31. In a parallel circuit the total resistance is always less thanthe circuit.

32. In this case the total resistance is:

33. The total current flow in a parallel circuit is found using

34. Ohm’s law may be used to find the current flowing throuthe current drawn from the battery will be the sum of the curren

1R total----------------

1R1------

1R2------

1R3------ and so o+ +=

1R total----------------

11---

12---

12---+ +=

1R total----------------

11--- 0.5 0.5+ + 2= =

R total12--- 0.5 ohm= =

ITVRT-------

120.5------- 24 amp= = =

In R1 I;VR1------

121------ 12 amp= = =



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lel circuit is always the same as the using Ohm's Law.

lternative formula, derived from thebe used. This is given as:

ds are connected in series with theh a circuit is shown at Figure 7-8.

each

volts

lts across each

Electrics-DC


35. The voltage drop across each load, or resistor in a paralsupply emf (in this case 12 volts). This too can be demonstrated

36. When only two resistances are connected in parallel an ageneral formula for the total resistance of a parallel circuit, may

Series/Parallel Loads37. A series/parallel circuit is one in which some of the loasource of supply and some in parallel with it. An example of suc

In R2 and R3 I;122

------ 6 amps in = =

IT 12 6 6+ + 24 amp= =

R1 1ohm( )V IR 12 1× 12= = =

R2 and R3 2ohms( )V IR 6 2× 12 vo= = =

RT

R1 R2×R1 R2+-------------------=



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FIGUSeries Loads

y resolving it into the equivalent of it is first necessary to find the totalhe result to the sum of the series

3 ohms

83 ohms

Electrics-DC


RE 7-8/ Parallel

38. The total resistance (RT) of this circuit may be found bthree resistances in series and adding them together. To do this,resistance of R3 and R4, the two paralleled loads, and add tresistances.

The total resistance of the circuit (RT) is given by:

R(3,4)

R3 R4×R3 R4+-------------------

4 2×4 2+------------

86--- 1.3= = = =

R1 R2 R(3,4)+ + 0.5 2 1.33+ + 3.= =



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ed against current flows that are in Apart from damage to the circuitause a very high current flow whichd. The standard safety devices take

protect the distribution cables of ad. The fuse is placed in series withted value of the fuse flows through

. The fuse wire is normally made of

e selected on the basis of the lowesthe known thermal characteristics of

peration of the aircraft, fuses are ofWhen replacing a blown fuse it iso circumstances should a fuse of ad at a value slightly in excess of theloads of initial switching on.

Electrics-DC


Circuit Protection39. It is essential that any electrical circuit should be protectexcess of the maximum rated value of the circuit in question. components, a short circuit to the airframe (for example) would cwould generate a great deal of heat and pose a serious fire hazarthe form of fuses and circuit breakers.

Fuses40. Fuses are thermal devices whose primary function is to circuit against excess current flow due to short-circuit or overloathe load (component) it protects. If a current that exceeds the rait, the fuse wire overheats and melts, resulting in an open circuita zinc alloy, which has the desired low melting point.

41. All fuses are rated in amperes (amps). In general fuses arrating which will ensure reliable operation of the system, given tthe cables, but which will not be prone to ‘nuisance failure’.

42. In emergency circuits, failure of which may affect safe othe highest rating possible consistent with cable protection. important that the new fuse be of the correct rating. Under nhigher value be used, since the rating has already been calculatenormal maximum load of the circuit, to accommodate the surge



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eramic tube with the fusible elementing caps. The caps fit into the inletaft cartridge type fuses are shown at

FIGUCartri

Electrics-DC


43. Aircraft fuses are of the cartridge variety, consisting of a c(wire) passing through it and connected at either end to conductand outlet terminals of the fuse holder. Examples of typical aircrFigure 7-9.

RE 7-9dge Fuses



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t power distribution points. Theseents that rupture successively underat absorbs the explosive effects ofg lugs and bolts, as opposed to thet the fuse holders incorporate an

er, heavy-duty, circuits. They have arrent before rupturing. This ensures, or transient, overload. The fusiblere to the dimension required for thegs. The central, waisted, portion ofwindow. Examples are shown at

Electrics-DC


44. Heavy Duty (High Rupturing) fuses are often installed aare of the cartridge type, but contain a number of fusible elemoverload conditions. The elements are packed in a medium thrupture. These fuses are fixed in position by means of mountinspring-clip attachment of lighter-duty fuses. In some aircrafindicator lamp which illuminates when the fuse ruptures.

45. Current Limiting Fuses are also used to protect high powhigh melting point and will thus carry a considerable overload cuthat power to the whole circuit is not lost in the event of a surgeelement is typically a strip of tinned copper, "waisted" in its centfusing area. The ends of the strip are formed into attachment luthe fuse is enclosed in a ceramic housing with an inspection Figure 7-10.



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FIGUCurrenFuses

Electrics-DC


RE 7-10t Limiting



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y. Repeated replacement of the fuse

uestion should be switched off. Theuching the live terminals of the fuse

f a lower rating may be tried, but of

aft is stipulated by Joint Aviation

n be replaced in flight, consisting of

ircuit by means of a mechanical tripoccurs. The advantage of a circuitce the overload situation has beenend by an increasing amount as the to the rated current flow for that

he circuit.

ip mechanism can be adjusted atistics. Thus, the circuit breaker canfor a given current. The ratings fors.

Electrics-DC


46. After a fuse has first blown it should be replaced once onlcould eventually result in overheating of the circuit.

47. When replacing fuses, ideally the power to the circuit in qperson replacing the fuse does not then have to worry about toholder.

48. If a fuse of the correct rating is not available, then a fuse ocourse it is likely to blow again.

49. The number of spare fuses to be carried in an aircrRequirements.

50. Spare fuses for all electrical circuits, the fuses of which ca50% of the number of each rating.

Circuit Breakers51. A circuit breaker or thermal trip is designed to isolate a cthat opens a switch whenever a surge of current, or overload, breaker over a fuse is that a circuit breaker can be reset onremedied. Circuit breakers make use of bimetallic strips, which btemperature of the strip increases. At the temperature matchedparticular circuit the bimetallic strip bends sufficiently to break t

52. The linkage between the bimetal element and the trmanufacture to achieve very close tolerance ‘trip-time’ characterbe matched not just to current, but to a specific maximum time circuit breakers are established in much the same way as for fuse



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d the non trip-free. In the trip-freenor will it reset the circuit breaker,l safety device that ensures that no

ke the circuit by holding the button, since the circuit breaker cannot beers are equipped with manual triptching device.

in some essential service circuits tod conditions, despite the fire hazard

, who require that it must not bead conditions. Figure 7-11 shows aration.

Electrics-DC


53. There are two types of circuit breaker, the trip-free ansystem depressing the reset button will not remake the circuit, until the overload condition has been cleared. It has an internaharm will be caused if the reset button is held in.

54. With a non trip-free circuit breaker it is possible to remain. Once the button is released the circuit will be broken againreset until the overload has been cleared. Some circuit breakbuttons so that they can be also used as a manually operated swi

55. In the past, non trip-free circuit breakers were installed permit emergency manual reconnection of supply under overloainvolved. Nowadays this practice is not permitted by the CAApossible for circuit breaker contacts to be re-made under overloschematic diagram of a non trip-free thermal circuit breaker ope



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FIGUThermBreake

can be connected together by a linkg link is commonly referred to as ahrow or double pole, double throw.ctrical systems namely push, toggle,witches. Most of these switches are

Electrics-DC


RE 7-11al Circuit r

Switches56. A basic switch consists of two contacting surfaces which or isolated by the reverse action (open the link). The connectinpole and there are various combinations i.e. single pole, single tThere are a number of different type switches used in aircraft elethermal, pressure, mercury, time, rotary, micro and proximity sexplained in the following text apart from the following :



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rded as a general purpose switch. A prevent accidental operation of the

before movement of the switch is

sed on the effects of differences ofVAR and steel. The different heat

ure control is required, typically thetems.

t in general it is based on an electriccts microswitches or operates springammed sequence of operations at

circuit, behave in some ways like a is applied in one direction. A fully original charging supply, when the

ial, separated by a non-conductingallel with a 6v DC source. If the 3-

across the capacitor is the same as

Electrics-DC


(a) Toggle Switch. A tumbler type switch which is regaguard cap can be incorporated which is a device toswitch. The guard has to be physically lifted possible.

(b) Thermal Switch. The priciple of operation is baexpansion between dissimilar metals typically INsensitivity can be used where automatic temperatoperation of control valves in thermal de-icing sys

(c) Time Switch. The principle of operation varies, bumotor driving a cam assembly which in turn contadriven mechanisms. This will produce a progrlimited intervals.

The Capacitor57. Capacitors are electrical storage devices which, in a DC battery. They will accept and store an electrical charge if voltagecharged capacitor will discharge, in the opposite direction to thecharging supply ceases and a suitable circuit is provided.

58. A capacitor comprises two plates of conducting matermaterial called a dielectric. Figure 7-12 shows a capacitor in parposition switch is placed in position A, the potential differencethat across the battery (6v).



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FIGUCapacOpera

has an excess of electrons whilst therons. In other words, one plate is of this is to put the dielectric underComparison can be made with and strength. However the electricallarities, whereas a magnetic field is

rent through a conductor.

will store its electrical charge, therecapacitor to store an electric chargehe capacitance of a given capacitore thickness of the dielectric and the of a material are measured againstnity (1).

Electrics-DC


RE 7-12itor tion

59. This is because the plate connected to the negative sourceplate connected to the positive source has a deficiency of electnegatively charged and the other is positively charged. The result‘strain’ establishing an electric field between the two plates. magnetic field which has lines of force with definite direction afield is achieved as a result of opposing positive and negative poachieved naturally from a permanent magnet or by passing a cur

60. If the switch is now moved to position B, the capacitor being no circuit through which it can discharge. The ability of a is called capacitance and is measured in units known as farads. Tis determined by three factors, the area of the capacitor plates, thmaterial of which the dielectric consists. The dielectric qualitiesthe dielectric property of air, which is given a constant value of u



Chapte

bility of air in a capacitor is said toa, which has a dielectric constant of

d to position C, the capacitor willacitors are discussed in depth in the

e is given by the formula:

s given by the formula:

A single cell, or more commonly abattery cells produce direct current,

harged is known as a primary cell.ll.

so on

so on

Electrics-DC


61. A material that has five times the electric charge storing ahave a dielectric constant of 5. A material commonly used is mic5.5.

62. When the switch in the circuit at Figure 7-12 is movedischarge its stored charge through the circuit that is made. Capchapter on AC Electrics.

63. When cacitors are connected in series the total capacitanc

When capacitors are connected in parallel the total capacitance i

Batteries64. Chemical action produces electricity in an electric cell. number of cells connected together, is known as a battery. All because they are of constant electrical polarity.

65. A cell that cannot be recharged once it has become discConversely, a cell that is rechargeable is known as a secondary ce

1CTOTAL--------------------

1C1------

1C2------

1C3------

1C4------ and+ + +=

CTOTAL C1 C2 C3 C4 and+ + +=



Chapte

ists of two plates made of differenttrolyte is a solution of water and ae of conducting current because itns.

milar plates causes an electron flow positively charged (a deficiency oftrons). There is thus a potential

ectrons to flow from the negatively will eventually balance the chargerrent flow ceases – the cell is said touired to restore the differential (or

sitive core is made of carbon, the chloride in paste form, hence the

milar low voltage portable devices.igure 7-13.

Electrics-DC


66. The principle of any electric cell is simple. The cell consmetals and placed in a solution termed the electrolyte. An elecchemical compound that will conduct electricity. It is capablcontains atoms having a positive or negative charge, known as io

67. The chemical action of the electrolyte acting on the dissifrom one plate to the other. Consequently, one plate becomeselectrons) and the other negatively charged (a surplus of elecdifference (voltage) between the two plates.

68. Connecting the two plates via a conductor will allow elcharged plate to the positive – an electrical current flow. Thisbetween the plates, so that there is no potential difference and cube discharged and a reversal of the chemical action will be reqelectrical pressure).

Primary Cells69. In the case of the conventional primary dry cell the ponegative case is made of zinc, and the electrolyte is ammoniumname ‘dry’ cell. Dry cells are mainly used in flashlights and siDiagrams of a simple primary and secondary cell are shown at F



Chapte

FIGUPrimarSecond

rcraft electrical system is to providehe engine-driven generators, is notcraft is shown at Figure 7-14.

Electrics-DC


RE 7-13y & ary Cell

Secondary Cells70. The principal function of the battery or batteries in an aielectrical power when the primary source of electrical power, tavailable. A basic electrical supply system for a single engine air



Chapte

FIGUBasic AElectriSystem

be used as the emergency source of It must be possible to recharge thehe generators) so that it is availableries are currently in common use in

Electrics-DC


RE 7-14ircraft

cal Supply

71. Secondary cells are clearly necessary when a battery is toelectrical power, as is the case in all aircraft electrical systems. battery after use, from the aircraft’s primary electrical source (twhen subsequently required. Two types of secondary cell batteaircraft, the lead-acid battery and the Nickel-Cadmium battery.



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attery of cells, known as a storageve plates made of lead peroxide andgative plates, comprising a cell, islectrolyte of 30% sulphuric acid and

oth plates to lead sulphate, and theecific gravity of the electrolyte also original compositions, and the acid specific gravity.

ell, the plates of dissimilar metalsigure 7-15, to present the greatestrs, made of insulating material, are

l short-circuiting.

Electrics-DC


Lead-Acid Batteries72. Secondary cells are usually grouped together to form a bbattery. In the case of the lead-acid battery this consists of positinegative plates made of lead. Each pair of positive and neconnected in series with the next and the whole immersed in an e70% water.

73. As the battery is discharged the chemical action changes bstrength of the electrolyte is reduced. As a consequence the spreduces. As the battery accepts a charge the plates revert to theirof the electrolyte strengthens. This results in a higher electrolyte

74. In order to optimise the performance of a lead-acid cdescribed previously are sandwiched alternately, as shown at Fpossible surface area for a given volume of electrolyte. Separatofitted between the positive and negative plates to prevent interna



Chapte

FIGUSecond(Lead Const

when in a fully charged state. It isroduce a battery that has a nominallts when connected to a substantial2v (6 cells in series) or 24v (12 cells

ty. Battery capacity is measured iny the battery was discharged to zerois checked every 3 months, if it fallsraft service.

Electrics-DC


RE 7-15ary Cell

Acid) ruction

75. A lead-acid cell will produce a maximum of 2.1 volts therefore necessary to connect 6 such cells together in series to pcharge of 12 volts, since cell voltage falls to approximately 2 voload. Aircraft batteries of the lead acid type are usually rated at 1in series).

76. Batteries are rated according to their voltage and capaciterms of ampere-hours at a five-hour discharge rate; that is to savoltage in five hours to determine its capacity. Battery capacity below 80% of its rated capacity the battery is removed from airc



Chapte

urs (Ah) is capable of supplying 30n alternative capacity rating is oftenharge rate. This is based upon theat an initial temperature of 26.7°Cul indication of a lead-acid battery's

it is installed by placing a rated, oreter. The voltmeter should show the seconds. If it fails to do so this is ariorated and it should not be placed

olution tends to break up into itsf in gaseous form from vents in the electrolyte solution diminishes andquires topping up with water from

-acid cell by using a hydrometer torged the SG should be between 1.25n fully discharged the SG is likely toigh state of charge; 1.24 to 1.275ge state.

Electrics-DC


77. A 12 volt battery with a capacity rating of 30 ampere-hoamps for 1 hour, 3 amps for 10 hours or any multiple thereof. Aapplied to aircraft storage batteries, known as the 5-minute discmaximum current a battery will supply over a 5-minute period and a final average cell voltage of 1.2 volts. This provides a usefengine starting performance.

78. The health of a fully charged battery is checked before known, load across the battery terminals in parallel with a voltmrated voltage of the battery, and should continue to do so for 15good indication that the internal condition of the battery has detein service.

79. During charging some of the water in the electrolyte sconstituent elements, hydrogen and oxygen, which are given oftop of the battery casing. Hence the proportion of water in thethe specific gravity of the solution increases. Hence a battery retime to time.

80. It is possible to determine the state of charge of a leadcheck the specific gravity (SG) of the electrolyte. When fully chaand 1.30, depending upon the age and condition of the cell. Whehave fallen to approximately 1.17. 1.275 to 1.3 indicates a hindicates a medium charge state; 1.2 to 1.24 indicates a low char



Chapte

d battery, because of the risk of acidf checking the state of charge of ane (OCV) and closed-circuit voltage on load.

y state of charge until the battery isf charge diminishes.

f direct current of greater emf thanarge must be kept reasonably low,ill result. There are a number ofther than during the trickle chargeelow:

area, to disperse vented gas.

disconnecting the charging leads; to

rrectly.

arging as the electrolyte tends to.

onnect the negative lead first; whenhelps avoid accidental short-circuitsthe battery.

Electrics-DC


81. A hydrometer check is not usually practical on an installespillage and resultant airframe corrosion. An alternative way oinstalled battery is by checking the battery open-circuit voltag(CCV), in other words checking the battery voltage off-load and

82. Off-load voltage (OCV) is virtually unaffected by batteralmost fully discharged, but CCV will fall significantly as state o

83. Batteries are recharged by connecting them to a source othe battery itself. The rate at which the battery accepts the chotherwise overheating and subsequent buckling of the plates wprecautions to be observed when charging lead-acid batteries, oreceived from the aircraft generators, and these are summarised b

(a) Charging must be carried out in a well-ventilated

(b) The battery charger should be switched off beforeavoid sparks at the battery terminals.

(c) Ensure the battery vents are clean and working co

(d) Remove the battery from the aircraft before chvaporise during charging and it is highly corrosive

(e) When removing the battery from the aircraft discreplacing it reconnect the negative lead last. This between the airframe and the positive terminal of



Chapte

explosive mixture of oxygen andan aircraft is vented to atmosphere.y significant period. If they are, the not subsequently accept a charge.

craft structure and to human tissue.nt for dealing with deposits in lead-t Aid treated with a copious flow of

te of soda.

ndary cell in use. Another commontery. In this the positive plates are

hydroxide. The electrolyte is aDuring charging the negative platestes pick up oxygen to form nickel

id cells, they have a longer life, they They also have a greater power-to-ly capable of producing 1.2 to 1.25ted voltage of 24 volts may have 19m total voltage required.

Electrics-DC


84. As a lead-acid battery is being charged it gives off anhydrogen. It is for this reason that the battery compartment in Lead-acid batteries should not be left in a discharged state for anplates will become coated with lead sulphate and the battery will

85. Battery fluid spillage will cause problems, both to the airBicarbonate of soda (baking soda) is an effective neutralising ageacid battery compartments. Acid burns to the skin should be Firswater, followed by treatment with a dilute solution of bicarbona

Nickel-Cadmium Batteries86. The lead-acid battery is by no means the only type of secotype of cell in aircraft use is the Nickel-Cadmium (NiCad) batmade of nickel hydroxide and the negative plates of cadmiumsolution of 70% distilled water and 30% potassium hydroxide. give up oxygen and become cadmium, whilst the positive plaoxides. During discharge the process is reversed.

87. NiCad batteries are lighter and more robust than lead-acare easier to store and they do not give off gases whilst charging.weight ratio. However, they are more expensive and they are onvolts per cell. Consequently, a nickel-cadmium battery with a racells or 20 cells connected in series, depending upon the maximu



Chapte

f the state of charge of this type ofeck against rated load, since voltageod. The only check that a NiCadn charge’. As the state of charge

ypically about 28 volts in a nominale battery voltage falls to about 25

r to that of Lead-Acid batteries, ind negative plates immersed in thetive terminals. A continuous wovenomprising the battery are assembledxy resin. A diagram showing the

Electrics-DC


88. The specific gravity of the electrolyte is no indication obattery, nor can the state of charge be determined by a voltage chremains substantially constant over most of the discharge peribattery is fully charged is the value of battery voltage when ‘oreaches full charge the battery voltage rises to a maximum level, t24-volt battery. As soon as the charging current is removed thvolts.

89. Construction of NiCad batteries is fundamentally similathat the cells are made up of interleaved alternate positive anelectrolyte and joined at their upper ends to the positive and neganylon separator insulates the plates from each other. The cells cwithin a container made of fibreglass or steel coated with epoconstruction of a typical nickel-cadmium cell is at Figure 7-16.



Chapte

FIGU

the formation of white crystals on reacting with the carbon dioxide in

Electrics-DC


RE 7-16

90. A Nickel-Cadmium battery may indicate overcharging bythe top of the battery. This is due to expelled electrolyte vapourthe atmosphere to produce potassium carbonate.



Chapte

r than that of a lead-acid battery ofnd for a longer period. The batteryhe discharge cycle, falling markedlyteristic makes the NiCad batteryong start cycle requires protracted power to recharge the battery.

total plate area within the cells and The Ah rating is always determined

emperature conditions and chargingndition known as thermal runaway

g and eventual melting of the plates

hich Nickel-Cadmium batteries areing, oxygen is formed at the positivet will re-combine with the cadmiumtinue the battery may be seriously

afe limits and by monitoring batterymperature sensor that activates anit.

Electrics-DC


91. The capacity of a nickel-cadmium battery is much greatesimilar size and weight and thus it will deliver far more power avoltage remains essentially constant over almost the whole of tonly as the battery becomes fully discharged. This characparticularly suitable for gas turbine engine starting, where a lbattery discharge before the engine-driven generators can supply

92. The capacity of a NiCad battery is a direct function of themay be up to 80-ampere hours (Ah) in a typical 24-volt battery. at a 5-hour discharge rate unless otherwise specified.

Thermal Runaway93. Batteries will perform to their rated capacities so long as trates are kept within the specified limits. If either is exceeded a comay occur, which causes boiling of the electrolyte, violent gassinand battery casing.

94. Thermal runaway, or vicious cycling, is a condition to wparticularly susceptible at high charging rates. During overchargplates of the battery. If this oxygen reaches the negative plates iand generate heat as a result. If this process is allowed to condamaged, or even explode.

95. The condition is avoided by keeping charge rates within stemperature. Some aircraft NiCad batteries incorporate a teoverheat-warning indicator, or a temperature gauge, in the cockp



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tery must be disconnected from the

d in series or in parallel as shown atof all the battery voltages, but the

y. When connected in parallel there-hour capacity is the sum of the

FIGUBatterConneSeries

Electrics-DC


96. If the temperature rises above a prescribed level the batbus bar by opening the battery master switch.

Connecting Batteries97. Two or more batteries of the same rating may be connecteFigure 7-17 and Figure 7-18. In series the voltage is the sum ampere-hour capacity remains the same as for a single battervoltage remains the same as for a single battery, but the ampecapacities of all the batteries.

RE 7-17ies cted in



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FIGUBatterConneParalle

FIGUCompLead ANicad

d

volts per cell

remains constant during entire harge cycle

can present a big problem if ery is overcharged internal cell tance is low and a high charge ent can cause overheating

be stored for long periods of time charged or discharged state

Electrics-DC


RE 7-18ies cted in l

RE 7-19arison of cid and Batteries

Lead Acid Nica

Voltage 2.1 volts per cell 1.2

Load Behaviour Close circuit voltage (CCV) falls gradually from full charge to end of discharge cycle

CCVdisc

Thermal Runway Not normally affected Thisbattresiscurr

Storage Life Best stored for short periods Canin a



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ater power-to-weight ratio than the charge rates and can discharge atd-acid batteries. It also has a lower nicad battery is that it develops a

lectrical systems. They provide usesback-up supply to some navigationnt that all power generating sources

electrical current flows through a conductor. The greater the currentetic field surrounding it. This is the

field, providing that there is relativehe lines of magnetic force, electricalgnetic induction and is the principle

ss the basic principles of magnetism.

Electrics-DC


98. An advantage of the nicad battery is that it contains a grelead acid and it also has a higher capacity. It can accept highequally high rates without the voltage drop associated with leasusceptability to very low temperatures. A disadvantage of thememory which must be periodically erased.

99. Either lead-acid or nicad batteries are used in aircraft eranging from main engine or APU starting, emergency lighting equipment (INS) through to an emergency DC supply in the evefail.

Magnetism and Electricity100. Magnetism and electricity are inseparable. When an conductor a magnetic field is created around the current-carryingflow through the conductor, the greater the strength of the magnprinciple upon which electro-magnets work.

101. When an electrical conductor is placed within a magneticmovement between the two such that the conductor cuts across tenergy is induced in the conductor. This is known as electro-maupon which a generator works.

102. Before considering these principles it is necessary to discu



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ial that from early times was foundls. The early explorers used thisreely, one end always points in a

ms or molecules appear to combine domain. Each domain has a northon these domains will be randomly

FIGUDomaMagneIron

oft iron the magnetic domains beginepeated application of the externalming aligned until, ultimately, they

Electrics-DC


Magnetism103. The name is derived from magnetite, an iron-oxide materto have the property of attracting iron and similar materia"lodestone" as a primitive compass since, when suspended fNortherly direction.

104. The composition of soft iron is such that groups of its atoto produce small permanent magnets, each known as a magnetic(N) pole and a south (S) pole. In a demagnetised piece of soft iraligned as at Figure 7-20.

RE 7-20ins - De tised Soft

105. When an external magnetic field is applied to a piece of sto align themselves with the polarity of the external field. Rmagnetic field will result in more and more of the domains becoare all aligned in the same direction.



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its length in one direction with oneo-magnetic field will have the samemagnetic field strength and it is saidre 7-21.

FIGUSoft IrMagneSatura

of the molecules of the soft iron willr magnetised, or at best only a fewal magnetism. This is because softagnetised. It becomes magnetised

s, but for the same reason it equally

ity. The internal friction of theirolecules have become aligned, or

such a material is removed from thent magnet.

Electrics-DC


106. A bar of soft iron can be magnetised by stroking it alongpole of a permanent magnet. Repeated application of an electreffect. At this point the soft iron bar is producing its maximum to be magnetically saturated. This situation is illustrated at Figu

RE 7-21on - tically ted

107. However, once the source of magnetism is removed most revert to their random polarisation and the material is no longeremain aligned and the material retains a small amount of residuiron has high permeability, the ability of a material to become measily because there is little internal friction between its moleculeeasily loses its magnetism.

108. Hard iron and some steel alloys have low permeabilmolecules makes them difficult to magnetise, but once the mpolarised, the same internal friction keeps them aligned. When magnetising source it retains its magnetism to become a permane



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et is shown at Figure 7-22.

FIGUPerma- Flux

nt to each other flux patterns arensidered as travelling from north torovides a useful reference by whichtic materials that can be used for

Electrics-DC


109. The magnetic field, or flux, pattern of a permanent magn

RE 7-22nent Magnet Pattern

110. When unlike or like magnetic poles are placed adjaceproduced as shown at Figure 7-23. Magnetic force, or flux, is cosouth in invisible lines. Whilst this is not literally the case it pcalculations can be made and effects considered. Ferromagnepermanent magnets are:

(a) Hard steel

(b) Nickel

(c) Cobalt

(d) Alnico (alloy of iron, nickel and cobalt)

(e) Remalloy

(f) Permandur



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FIGUPolar F

arrying Coresgnetic field is produced around thate direction of the magnetic flux ismined by application of the ‘screw

Electrics-DC


RE 7-23lux Patterns

Magnetic Fields Surrounding Current C111. When an electric current flows through a conductor a maconductor due to the movement of electrons through it. Thdependent upon the direction of current flow and can be deterrule’. This is shown at Figure 7-24.



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FIGUScew R

tion of current flow the direction ofust be rotated in order to propel the

op, the combination of two circularas illustrated at Figure 7-25.

Electrics-DC


RE 7-24ule

112. If you visualise a simple wood screw pointing in the directhe magnetic flux is given by the direction in which the screw mscrew in the direction of the current flow.

113. When the current-carrying conductor is formed into a loflux patterns produces a field having a north and south polarity,



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FIGULoop F

Electrics-DC


RE 7-25lux Patterns



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p and the direction of current flow.current flowing ‘into’ A and ‘out of’d as shown. The strength of this

other way of indicating the directionesents the tail, or flight of an arrowof the paper. The arrow indicateslockwise around the cross and anti-ic field, from north to south.

loop can be determined by ‘looking’5.

agnetic field, by using a number ofuced by a coil of wire is shown at

Electrics-DC


114. Figure 7-25 shows a plan view of the current-carrying looFigure 7-25 shows a section through the loop at A and B. With B the two circular fields combine to produce a polarised fielmagnetic field will depend upon the strength of current flow.

115. The ‘dot’ and ‘cross’ convention used in Figure 7-25 is anof current flow and the associated flux direction. The cross reprentering the paper, the dot the point of the arrow coming out conventional current direction (+ to -) and flux flow is always cclockwise around the dot. Flux flow is always out of the magnet

116. The magnetic polarity at either end of a current-carrying at the end of the loop and using the method shown at Figure 7-2

117. The effect can be concentrated, to produce a stronger mloops in the form of a coil of wire. The magnetic field prodFigure 7-26.



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FIGUMagnefrom CCarryi

by a conducting coil is directly the coil and inversely proportional

coil is wound around a cylindricalf the core.

Electrics-DC


RE 7-26tic Field urrent

ng Coil

118. The magnetic field intensity, or strength, produced proportional to the current strength and the number of ‘turns’ into the length of the magnetic circuit. If it is imagined that thecore, the length of the magnetic circuit is, effectively, the length o



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rrying coil shown at Figure 7-27 theign with the flux direction so that itc field. Soft iron is used because ofnetised. That is to say the magnetic, but will take up random alignment

, the electro-magnet can be switchede strength of the electro-magnet canreasing the current flow in the coil.e core determine the strength of angure 7-27.

Electrics-DC


Electro-Magnets119. If a soft iron rod, or core, is placed within the current-camagnetic flux will traverse the iron core and its domains will albecomes a magnet with polarity the same as the electro-magnetiits high permeability, or ability to become magnetised or demagdomains will readily align with the external (coil-produced) fieldwhen the external field is removed.

120. Since the external field requires current flow to produce iton or off by switching the coil current on or off. Furthermore, thbe increased (up to saturation level in the soft iron core) by incThus, the number of turns in the coil and the permeability of thelectro-magnet. An example of an electro-magnet is shown at Fi



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FIGUElectro

an electro-magnet may be used to operate a piece of equipment, suchtically operated mechanical linkage

low-current switching devices. The

Electrics-DC


RE 7-27 Magnet

Relays121. The flux field of the magnetism induced in the core ofcreate mechanical movement, which subsequently can be made toas a switch. An electro-magnet having a fixed core and a magneis called a relay and is often used for the remote operation of principle is illustrated at Figure 7-28.



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FIGURelay

Electrics-DC


RE 7-28



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rmature made from a material thatthe field of the electro-magnet. The some cases several sets, of contactagnet closes the relay contact pointsm the contact circuit by means off a relay will determine its type. Argised. A normally closed relay willoy a changeover contact which willcoil is energised.

which case the device is known as aagnetic field produced attracts the

ing. When current to the coil of thes and the spring forces the core outate a variety of mechanical devices

on-magnetic sleeve fixed within theeeve. A solenoid is illustrated at

Electrics-DC


122. Attached to the mechanical linkage of the relay is an aresponds to a magnetic field and which is therefore attracted by movement of the armature is used to open or close a set, or inpoints. In the example at Figure 7-28 the action of the electro-mto complete an electric circuit. The coil circuit is insulated froinsulated arms or insulated stops. The fuctional requirement onormally open relay will not activate a circuit until its coil is enede-activate a circuit when its coil is energised. A relay can emplfunction to activate one circuit and de-activate another when its

Solenoids123. In some electro-magnets the soft iron core is movable, in solenoid. When the coil of the electro-magnet is energised the msoft iron core and draws it into the coil against the action of a sprelectro-magnet is switched off the electro-magnetic field collapseof the coil. The linear motion of the core may be used to opersuch as electrical contact points, valves and circuit breakers.

124. Usually the core of a solenoid consists of two parts, a ncoil windings and a soft iron core that slides within the slFigure 7-29.



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FIGUSoleno

emote operation. They can be sitede position using low power switches

ectrical energy without the aid of of a magnetic field it is known as

Electrics-DC


RE 7-29id

125. The major advantage of solenoids is their suitability for ralmost anywhere within the aircraft and controlled from a remotor control units and circuitry.

Electro-Magnetic Induction126. Induction is the name given to the transference of elconductors. When this transfer of energy is achieved by meanselectro-magnetic induction.



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ovement between a conductor ands of magnetic flux. If the conductor induction. The relative movementfield or a stationary conductor in aically moving a magnet, as in someo-magnet, as in a transformer. Thetion of AC Electrics.

netic induction produces power, is

FIGUGener

ld an electromotive force (emf), ornd thus power to be produced, if the The direction in which the inducedtrated at Figure 7-31.

Electrics-DC


127. Electro-magnetic induction occurs when there is relative ma magnetic field such that the conductor is cutting across the linemoves parallel to the lines of flux, there is no electro-magneticmay be achieved either by a moving conductor in a stationary moving field. The field flux may be made to move by mechangenerators, or by varying the strength of the current in an electrsubject of transformers is discussed in the Power Distribution sec

128. The principle of generator action, in which electro-magillustrated at Figure 7-30.

RE 7-30ator Action

129. As the conductor is moved across the magnetic flux fievoltage, is induced in it. The voltage will cause current to flow, aconductor forms a complete electrical circuit, as at Figure 7-30. emf acts is determined by Fleming's right hand rule. This is illus



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FIGUFleminHand R

ose laid down by Faraday and Lenz.

the magnetic field linked with the

the rate of change of the magnetic

Electrics-DC


RE 7-31gs’ Right ule

The Laws of Electro-Magnetic Induction

The important laws concerning electro-magnetic induction are th

Faraday's Laws state that:

1st Law. An induced emf is established in a circuit whenevercircuit changes. (in intensity or polarity).

2nd Law. The magnitude of the induced emf is proportional toflux linked with the circuit.

Lenz's Law states that:



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poses the change in flux that causes

tor within a magnetic field results inontinuous generation of emf in the conductor and magnetic field. Thistween the poles of a magnet.

ducting loop, or armature, placed in an external mechanical drive (in an

tes, the sides of the loop cut throughop and field induces an emf in the

Electrics-DC


The induced emf acts to circulate a current in a direction that opthe emf.

Generators130. It was shown in section 2.127 that movement of a conducthe induction of an emf in the conductor. It follows that, for cconductor, there must be continuous relative movement betweenis achieved by rotating a conducting loop within the flux field be

131. Figure 7-32 shows a simple generator consisting of a conthe field between the poles of a permanent magnet and rotated byaircraft this drive is from the engine). As the armature loop rotathe flux field and this relative movement between conducting loloop.



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FIGUSimple

Electrics-DC


RE 7-32 Generator



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ring, which is in contact with ad to an electrical circuit in which isrrent to flow through the complete

onary magnetic field is shown in

llel to the flux field. Hence, there isfore no induced emf. This is also

es of the loop are moving at rightnduced emf is maximum. This is

ept that the armature has rotated

ept that the armature has rotatedare moving in the opposite directionnow of opposite polarity. Since,

trical polarity is reversed the currentput curve at Figure 7-32.

Electrics-DC


132. Attached to each end of the loop is a conducting slipstationary brush made of carbon. The two brushes are connecteplaced a load. The emf induced in the armature loop causes cucircuit formed by the armature and load circuit.

133. The progressive rotation of the armature in the statiFigure 7-32.

134. At point 1 the sides of the armature loop are moving parano relative movement between conductor and field and thereindicated on the output curve at Figure 7-32.

135. At point 2 the loop has rotated through 90° and the sidangles to the field; relative movement is at its greatest and iindicated on the output curve.

136. At point 3 the situation is the same as at point 1, excthrough 180°.

137. At point 4 the situation is the same as at point 2, excthrough 270° in the stationary field, so the sides of the armature relative to the field. Induced emf is at a maximum, but is conventionally, current flows from positive to negative if the elecflow will be reversed. This is indicated by the reversal of the out



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and the situation is the same as att of this simple generator reverses invery half cycle, or half revolution ofterminals of the brushes will reverserent is known as alternating current or windings.

lternator. Most modern aircraft use the output of the alternator cannot Consequently, the output of theeet such requirements. Virtually alldriven alternators whose output isas DC. Larger commercial aircraft is distributed as AC and only a few

or is to produce direct current at the are heavier than alternators, andraft is limited almost entirely to ator) found in some small turbine

r to that of the simple AC generator

Electrics-DC


138. At point 5 the armature has completed a full revolutionpoint 1. From the foregoing it will be seen that the voltage outpupolarity and fluctuates in magnitude from zero to a maximum ethe armature. Any current flowing in a circuit connected to the in direction and fluctuate in magnitude every half cycle. Such cur(AC). All rotary generators produce AC in their armature loops,

139. The device illustrated above is a simple AC generator, or aalternators as their primary source of electrical power, but clearlybe used where DC is required, such as for battery charging.alternator must be converted to DC, by means of a rectifier to mmodern, light piston-engine aircraft are equipped with engine-rectified and distributed to all the major electrical components are equipped with AC Generators (big alternators) whose outputservices use rectified AC (i.e. DC).

DC Generator140. An alternative to rectification of the output of an alternatoutset, by means of a DC generator. Because DC generatorsgenerally require more maintenance, their use in modern airccombined engine starter motor and generator (starter generapowered aircraft

141. The construction of a simple DC generator is very similadescribed previously.



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a pulsating DC can be obtained bylip ring’, or split ring, known as aalf of the commutator split ring andf mica.

Electrics-DC


142. Figure 7-33 shows a simple DC generator from which replacing the slip rings of the AC generator with a ‘two-part scommutator. Each end of the armature loop is attached to one hthe two halves are insulated from each other, usually by a strip o



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FIGUSimpleGener

Electrics-DC


RE 7-33 DC ator



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that, as the commutator rotates, thee current flow in the armature hashe opposite direction. This ensuresgments is minimal at this point ofure that the electrical polarity of thethe load circuit remains constant.

DC produced by such a generator isng devices, which require constante number of armature loops and the which has only a slight pulsation,

duced by a single armature loop to appreciated that, in the latter, the

Electrics-DC


143. The stationary brushes are placed opposite each other so brushes pass from one commutator segment to the next as thceased flowing in one direction and is about to start flowing in tthat the potential (emf difference) between the commutator sechangeover, or switching. The effect of the commutator is to ensbrushes remains constant and so the direction of current flow in

144. As can be seen from the output curve at Figure 7-33 the pulsating and this is unsatisfactory for many current-consumivoltage supply. This can be very nearly achieved by increasing thnumber of magnetic poles, producing a voltage and current flowknown as commutator ripple.

145. Figure 7-34 illustrates the change from pulsating DC prothe ripple DC achieved by four armature loops. It should becommutator would consist of eight segments instead of two.



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FIGUDC froArmat

ting loop of wire would be minute.n core to form the armature coil, org the magnetic field into the desired

Electrics-DC


RE 7-34m Four

ure Loops

146. Clearly the output of any generator using only one rotaConsequently many thousands of loops are wound onto a soft irowinding. The core serves additionally as a means of concentratinarea.



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system is that the voltage deliveredons of electrical load. The aircraftnd therefore its speed of rotation is

at the emf induced in the armaturetion of the armature. However, it isd control as a means of controlling

tro-magnetic induction is dependent

of magnetic flux (in this case speed

number of turns in the armature

).

first factor as a means of generatorcture. This leaves only one optionwhich the armature rotates.

Electrics-DC


Field Excitation147. A fundamental requirement of any electrical distributionto the bus bar should be maintained constant under all conditigenerator, whether AC or DC, is driven by the aircraft engine avariable, especially in the case of the piston engine aircraft.

148. Referring back to Faraday’s second law it is apparent thwindings of a generator will vary directly with the speed of rotaclearly not a practical proposition to use generator variable speethe emf induced and therefore the output voltage.

149. The magnitude of the emf induced in a conductor by elecupon the following factors:

(a) The rate at which the conductor is cut by the linesof armature rotation).

(b) The length of the conductor (determined by thewinding).

(c) The flux density (the strength of the magnetic field

150. We have already seen that it is not practical to use the output voltage control, and the second is fixed during manufaremaining, variation of the strength of the magnetic field within



Chapte

Figure 7-33 the generator field isariable. However, as was shown ind varying the current strength canwound around a soft iron core thewill be an electro-magnet, the fluxs is known as field excitation.

, and therefore the generator outputctrical load by varying the currentwinding is known as the generator current. Figure 7-35 illustrates the

Electrics-DC


151. In the simple generators illustrated at Figure 7-32 andproduced by a permanent magnet and is therefore clearly not vthe previous chapter, passing a current through a conductor anproduce a variable strength magnetic field. If the conductor is magnetic field will be concentrated in the core and the device density of which can be varied from zero to saturation level. Thi

152. By this means the emf induced in the generator armaturevoltage, can be controlled regardless of generator speed or elesupplied to the core winding of the electro-magnet. This core field winding and the current supplied to it as the generator fieldbasic concept.



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FIGUField E

d from a battery and the strength ofe. Decreasing the resistance at thecreases the flux density in the core

ure loop is turning. Consequently,tput voltage rises. Increasing the

Electrics-DC


RE 7-35xcitation

153. In the diagram at Figure 7-35a the field current is suppliethe field current is controlled by means of a variable resistanccontroller allows more current to flow to the field coil, which inand thus the strength of the magnetic field in which the armatgreater emf is induced in the armature and the generator ouresistance at the controller has the opposite effect.



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tic form. When the generator field this case, this is known as external, maintain constant polarity of the excited generator is normally the

ken from the generator output, andmatically at Figure 7-36. From the the field coil through a controlling

FIGU

Electrics-DC


154. Figure 7-35b shows the field excitation circuit in schemacurrent is supplied from an external source of direct current, as inor separate excitation. Field current must be DC in order tomagnetic field so the source of field current for an externallyaircraft battery.

155. In DC generators the current to the field coil is usually tathis is known as self-excitation. This arrangement is shown scheoutput (positive) terminal of the generator current is supplied tovariable resistance to the field winding.

RE 7-36



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e generator armature, this is knownThe current flow through the field the distribution system loads and isnt output voltage.

5 a single electro-magnet is shown,nly around the armature.

nt of magnetism, known as residualagnetism is sufficient to induce an

, which initiates a current flow fromcreasing the magnetic field of then.

ersed, due to excess heat, shock orestored by briefly passing a current field.

tained constant and load variations that the field windings are in seriesnerator in which some of the field

n as a compound-wound machine ise depends on the number of turns innd the supplied load.

Electrics-DC


156. It will be noted that the field winding is in parallel with thas a shunt-wound generator and is the usual configuration. winding is quite separate from that flowing from the generator totherefore relatively simple to control, in order to maintain consta

157. In the purely diagrammatic representation at Figure 7-3but in practice there are a number of electro-magnets spaced eve

158. The soft iron of the electro-magnets retains a small amoumagnetism, even when there is no field current. This residual memf in the armature of the generator when it first starts to rotatethe generator. This current flow supplies the field coils, inelectromagnets, increasing induced emf in the armature, and so o

159. Residual magnetism may be lost, or its polarisation revreversal of field current flow. The residual magnetism can be rthrough the field. This is known as field flashing, or flashing the

160. In circumstances where the generator speed can be mainare relatively small it possible to configure the DC generator suchwith the armature (and therefore the load). Also, a type of gewindings are in series, and some in parallel with the load, knowsometimes used. In summary it can be said that generator voltagthe armature, the strength of the field, the rpm of the armature a



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urrent power distribution suppliedAC output being rectified to DC at suffer from the problems of arcingn.

mutator there is no need for theent to be transferred to the output

inding is in the stationary casing oflectro-magnets are on the rotor.ed through brushes and slip rings to

field cuts through the stationarye winding is connected to the outputpplied to the distribution bus bars.

machine illustrated would produceC is explained in paragraphs 306-

Electrics-DC


Alternators161. Many light single and twin-engined aircraft use direct cfrom alternating current generators (alternators), the alternator source. Alternators are lighter than DC generators and do not(and the consequent radio interference) produced by commutatio

162. Because the alternator does not require a rotating comarmature to be mounted on the rotor and the large load currterminals through slip rings and brushes. Instead, the armature wthe machine and the generator field windings and their eConsequently, only the relatively small field current need be passthe rotating field windings.

163. Thus, in aircraft alternators, the rotating magnetic conductors of the armature winding, inducing emf. The armaturterminals of the alternator, from which the load current is suthrough a rectification system that converts the AC output to DC

164. Figure 7-37 shows a simple alternator arrangement. Thesingle-phase alternating current. Single-phase and three-phase A321.



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FIGUSimple

Electrics-DC


RE 7-37 Alternator



Chapte

by regulating the current suppliedariable resistance in series with thehematic alternator circuit for a light

FIGUAltern

pplied from the aircraft battery.

Electrics-DC


165. The alternator output voltage is automatically controlledto the field windings in the rotor. This is done by means of a vfield windings called a voltage regulator. Figure 7-38 shows a scaircraft 28-volt DC distribution system.

RE 7-38ator Circuit

166. The field current for the separately excited alternator is su



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rrent flow in one direction only and

ator supplying DC to the electrical a current limiter. In the majority oftomatically opens if load becomes

the alternator output voltage and Typically the voltage regulator is ator field coil when alternator outputtage rises above a set value (say 28.5e is repeated about 2000 times per

contact voltage regulator. In lightregulation is often by means of al, which regulates generator voltaget Figure 7-39.

Electrics-DC


167. A rectifier is a static semiconductor device that permits cuthereby converts bi-directional AC into unidirectional DC.

Control and Protection168. The control and protection devices needed for an alterndistribution system of a light aircraft are a voltage regulator andsystems the current limiter is simply a circuit breaker that auexcessive.

169. Voltage regulators work on the principle of sensing adjusting field current to maintain voltage at a constant value. transistorised unit that allows a set current to flow to the alternavoltage falls below a set value (say 27.5 volts). When output volvolts) it cuts off the current supply to the field coil. This cyclsecond, maintaining alternator output voltage at about 28 volts.

170. An older version of the same principle is the vibrating aircraft, using comparatively low output generators, voltage vibrating contact regulator. The regulator contains a voltage coiand current. A vibrating contact regulator system is illustrated a



Chapte

FIGUVibratiRegula

in or out of the field circuit by anng contact). The contact points arethe voltage coil creates sufficientn the points against the force of the

he voltage coil is insufficient to openows through the points, creating ae.

Electrics-DC


RE 7-39ng Contact tor

171. The regulator contains a fixed resistance that is switchedelectro-magnetically-controlled spring-loaded switch (the vibratiheld closed by the spring until the current flow through electromagnetism to attract the ferritic contact armature and opeadjustable spring.

172. When generator voltage is low the current flow through tthe contact points. Field current to the shunt field winding flstrong generator magnetic field and causing output voltage to ris



Chapte

rrent flow through the voltage coilth to open the contact points. Field resistance. The low resultant fieldf the voltage coil, until the spring

r second, maintaining an essentiallyted by adjusting the tension of theh generators where the field currentct points would become overheated

uired for an alternator-supplied DCw current flow and a voltmeter towo instruments, the ammeter is theith voltmeters, only with a warning

truction. An ammeter is connected A voltmeter is connected in parallele usually of the permanent magnet

Electrics-DC


173. When output voltage reaches a predetermined level, cubecomes sufficient to create an electromagnet of sufficient strengcurrent can now only reach the shunt field winding via a fixedcurrent reduces output voltage, and electromagnetic strength ocloses the contact points once more.

174. This process is repeated at between 50 and 200 cycles pesteady voltage. The generator output ‘control’ voltage is adjusspring. Vibrating contact regulators are only suitable for use witis low (less than 8 amps). Above this level the vibrating contaand possibly fuse together.

Monitoring Devices175. There are basically only two monitoring instruments reqdistribution system. These instruments are an ammeter to shoindicate proper functioning of the voltage regulator. Of these tmore important and most single engine aircraft are not fitted windication if voltage strays outside preset limits.

176. Ammeters and voltmeters are remarkably similar in consin series with each generator to show the output current (load). with circuit loads to show circuit voltage. Both instruments ar‘moving coil’ type, shown at Figure 7-40.



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the instrument itself. Increasing the additional resistances connected in. Increasing the measuring range of parallel with the instrument. This is

FIGUMovinInstrum

iron core. The core is carried on an the poles of a permanent magnet.nd of which is attached to the core

Electrics-DC


177. Voltmeters usually have the required resistance built intomeasuring range of the voltmeter can be achieved by the use ofseries with the instrument. The resistances are called multipliersan ammeter is achieved by the addition of a resistor connected incalled a shunt resistance.

RE 7-40g Coil

ent

178. A coil is wound around a former mounted upon a soft spindle supported by a bearing, so that it is free to rotate betweeRotation of the core is limited, however, by a hairspring, one espindle and the other to the casing of the instrument.



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which interacts with the permanentd field, against the hairspring. Thehed to the rotating core is a pointerquired.

ection of the rotary movement andsed in conjunction with batteries toil indicators employ two hairspringsirection.

FIGUCentreAmme

pe, illustrated at Figure 7-42, whichased current flow through the coil.

Electrics-DC


179. Current flow through the coil sets up a magnetic field, field and causes the coil and core to rotate towards the weakenegreater the current flow, the greater the extent of rotation. Attacthat moves against a fixed scale, calibrated in amps or volts as re

180. Reversal of current flow through the coil reverses the dirthis is employed in the ‘centre-zero’ arc scale type of ammeter ushow charge or discharge, see Figure 7-41. Reversible moving cowith opposite winding, to restrict rotation of the core in either d

RE 7-41-Zero ter

181. A generator ammeter, or loadmeter, is of the ‘left-zero’ tyshows an increased deflection in a clockwise direction with incre



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FIGULeft-ZAmme

FIGUVoltme

are designed to offer a very lownd are designed to offer a very highload and not through the voltmeter.xed resistance, so the scale of the. See Figure 7-43.

Electrics-DC


RE 7-42ero ter

RE 7-43ter Circuit

182. Ammeters are connected in series with the load and resistance. Voltmeters are connected in parallel with the load aresistance, so that the majority of the current flows through the Current flow is proportional to the voltage in a circuit of fiinstrument can be calibrated in volts for a range of current flows



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ridge circuit is in balance if the fourconsider the variable resistance RVe bridge has a total resistance of 200 ohms. Applying ohms law (V = I xd R3 is .12 amps (24 = I x 200) and

Using these values we can calculateop across R1 is 8V (V = 0.8 x 100).

2 = 12V (V = .12 x 100). Thereforeen points B and D, current will flowent can be represented on a suitably

Electrics-DC


Wheatstone Bridge183. A typical bridge circuit is illustrated at Figure 7-44. The bresistances R1, R2, R3 and RV are equal i.e. 100 ohms. Now changes to a value of 200 ohms. It can be seen that one side of thohms R2 + R3 and the other side R1+RV a total resistance of 300R) with the supply voltage of 24V the current flow through R2 anthe current flow through R1 and RV is .08 amps (24 = I x 300). the voltage drop R2/R3 and R1+RV Points B and D. Voltage drTherefore potential at B = 24 - 8 = 16V. Voltage drop across Rpotential at D = 24 - 12 = 12V. If an instrument is placed betwefrom Point B (16V) to Point D (12V). Measurement of this currcalibrated scale i.e. fuel contents gauge.



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FIGU

f various systems in an aircraft canning lights are red, caution lights -rate magnetic indicators (MIs) as aor accompanied by aural warnings

Electrics-DC


RE 7-44

Annunciators184. Display of normal, abnormal, malfunctioning or failure obe indicated by using various coloured lights for instance, waramber and indicating lights - blue/green. Some aircraft incorpoform of annunciation. Some annunciations can be pre empted (bells and chimes).



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r an electric motor to operate fromr. An obvious example of this is theerating from a direct current supply.oducing greater starting torque andpplications. Examples of these are

s and that is free to rotate within aupplied with direct current and thegnet or it may be electro-magnetic. carbon brushes that are in contact

commutator. Figure 7-45 illustratesa single conducting loop and theotors the armature comprises manygnetically.

Electrics-DC


DC Motors185. There are a number of instances where it is necessary fothe aircraft battery, rather than from the engine-driven alternatoengine starter motor. Clearly such a motor must be capable of opCompared with its AC equivalent, the DC motor is capable of prso is preferable, even when AC power is available, in certain awing flap motors and landing gear.

186. A DC motor comprises an armature mounted in bearingstationary magnetic field. The armature is a conducting loop sstationary magnetic field may be supplied by a permanent maDirect current is supplied to the rotating armature via stationarywith the two halves of a longitudinally split cylinder known as aa very simple DC motor in which the armature consists of stationary field is supplied by a permanent magnet. In practical mconducting loops and the stationary field is controlled electro-ma



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FIGUSimple

and Rule. To demonstrate this theuch that they are mutually at rightn of the magnetic field (N to S) andithin the conductor, the thumb will

ustrated at Figure 7-46.

Electrics-DC


RE 7-45 DC Motor

187. Electric motors operate according to Fleming’s Left Hthumb, first and second fingers of the left hand are arranged sangles to each other. With the first finger pointing in the directiothe second finger pointing in the direction of conventional flow wpoint in the direction the conductor will tend to move. This is ill



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FIGUFleminHand R

Electrics-DC


RE 7-46g’s Left ule



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in a magnetic field is due to thehe conductor by the current flowing5 with the direction of current flowuent flux field produced. The crossdicates current flowing toward thetract the armature in the diagram is

FIGUDC MPrincip

Electrics-DC


DC Motor Principle of Operation188. The movement of a current carrying conductor withinteraction between the static field and the flux created around tthrough it. Figure 7-47 shows the simple DC motor of Figure 7-4in the armature indicated by a cross and a dot, and the conseqindicates current flowing away from the viewer and the dot inviewer. Since, in magnetism, like poles repel and unlike poles atcompelled to rotate in a clockwise direction.

RE 7-47otor le (1)



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through the left hand half of the current flow through the armaturelectro-magnetic field as described inon produces a torque force to rotate

FIGUDC MPrincip

ched commutator will have reachedommutator there is now no currentlux. However, inertia will keep the

Electrics-DC


189. With refernce to Figure 7-47, current is being suppliedcommutator and is returning through the right hand half. Thisloop, indicated by the cross and dot symbols, has produced an ethe preceeding paragraph. The consequent repulsion and attractithe armature in a clockwise direction.

RE 7-48otor le (2)

190. After a quarter of a revolution the armature and its attathe position shown at Figure 7-48. Because of the gaps in the cflow through the armature and therefore no electro-magnetic farmature rotating.



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FIGUDC MPrincip

es beyond the quarter turn position,the same direction as before. Thetant, maintaining the direction of

Electrics-DC


RE 7-49otor le (3)

191. It will be seen from Figure 7-49 that as the armature passcurrent once again flows through the commutator halves in polarity of the electro-magnetic field is thus maintained constorque.



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d relative to a magnetic field an emf through the stationary field as theature. This emf produces a currenterefore reduces the total armature

f is known as the net emf, and it isn order to ensure that the net emf iss possible.

egins to rotate, is determined by thew the current flow will be very high. current flow through the armature. starting and then quickly falls to ators have a resistance built in to the increases.

ersticsduced electro-magnetically, so that field strength. DC motors fall inton respect to the armature windings.nd wound, They are illustrated

Electrics-DC


Back emf and Net emf192. It has already been shown that when a conductor is moveis induced in the conductor. The loop of an armature is movingarmature rotates and this inevitably induces an emf in the armflow that opposes the applied current from the battery and thcurrent flow. The induced voltage is known as back emf.

193. The difference between the applied emf and the back emthis that determines the torque produced in the armature shaft. Isufficient the resistance of the armature windings is kept as low a

194. The initial current flow through the armature, before it bapplied voltage and the armature resistance. If the resistance is loAs the motor gains speed the back emf increases and reduces theThis explains why the current to a DC motor shows a surge onmuch lower value. To avoid excess starting current, some DC moarmature windings, which automatically cuts out as motor speed

Types of DC Motors and their Charact195. The stationary field of a DC motor is invariably provarying the current flow through the field windings will vary thethree types, according to the arrangement of the field windings iThese types are series wound, shunt wound and compoudiagrammatically at Figure 7-50, Figure 7-51 and Figure 7-52.



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series with the armature. Thus, theced in both armature and field coils

is particularly true at starting, when series wound motor is high startingrequired to start against a high loadrangement for a series wound motor

sed are engine starter motors, flap

operate without a mechanical load destruction. The reason for this isuces the net emf and therefore theuced current flow through the fieldld, preventing the back emf froms a net emf to continue accelerating

Electrics-DC


Series Wound Motor196. In the series wound motor the field coils are connected incurrent flow through both is the same and the magnetic flux induwill be strong, producing high torque in the armature shaft. Thisthe current flow is very high. Consequently a characteristic of thetorque. This is useful in circumstances where the motor will be and where the running load is also high. The schematic wiring aris shown at Figure 7-50.

197. Examples of instances where series wound motors are uoperating motors and landing gear operating motors.

198. Series wound motors should never never be allowed to applied. This is because they are liable to overspeed, possibly tothat, as the armature speed increases the induced back emf redcurrent flow through both armature and field windings. The redwindings reduces the strength of the stationary magnetic fiereaching a value where the net emf is zero. Hence there is alwaythe unloaded motor.



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FIGUSeries

FIGUShunt

Electrics-DC


RE 7-50Wound

RE 7-51Wound



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FIGUCompWoun

parallel) wound DC motor is shownare connected in parallel with the set to limit the field current to that

than the armature resistance. This isire for the field windings.

h, because of its low resistance. Thetheir relatively high resistance, andristic of the shunt wound DC motor

Electrics-DC


RE 7-52ound d

Shunt Wound Motors199. This schematic arrangement of the wiring for a shunt (or an Figure 7-51. From this it will be seen that the field coils armature windings. The resistance of the field coils is deliberatelyrequired for normal operation of the motor, and is much higher usually achieved by using many thousands of turns of very fine w

200. On start up the current flow through the armature is higcurrent flow through the field coils is low however, because of field strength is correspondingly weak. Consequently, a characteis low starting torque.



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will cause the armature current tohrough the field coils, since they aret to the armature current increases

eases until back emf almost equalsed. This speed is virtually constant

w and increases with motor speed.g load conditions is a requirement.

pound wound DC motor. This hasature and the other in parallel. Theigher resistance shunt windings ineristics of the series wound and theut will not overspeed under lightd under varying conditions of load.

here loads may vary from zero toy are often used to drive hydraulic

Electrics-DC


201. As the armature speed increases, increasing back emf decrease. Because the back emf does not affect the current flow tin parallel to the armature, the relative value of the field currenwith increased motor speed. The result is that the torque incrapplied emf and the motor settles at its normal operating spewithin the design mechanical load range of the motor.

202. Shunt wound motors are used when starting torque is loThey are particularly useful where constant speed under varyinTypical applications in aircraft are fuel pumps and fans.

Compound Wound Motors203. Figure 7-52 shows the schematic arrangement of a comtwo sets of field windings, one connected in series with the armlow resistance series windings are shown in heavy lining, the hlighter lining. The compund wound motor combines the charactshunt wound motor. It is capable of high starting torque, bmechanical loading and will maintain a reasonably constant spee

204. The compound wound motor is suited to applications wmaximum and where starting loads may be high. In aircraft thepumps.



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aps and landing gear, it is necessary means of a switching arrangementr the armature (but not both). Thiserse the direction of rotation of the

s more usual to employ a spli-fieldtion.

either wound in opposite directionsde of the motor casing. The motor is be placed in one of three positions. motor is broken. When thrown inates in (say) the clockwise direction. are energised, the field polarity issplit-field schematic circuit is shown

Electrics-DC


Reversible DC Motors205. In certain aircraft applications, such as motor-operated flfor the motor to be reversible. Clearly this could be achieved bythat reversed the polarity of the DC supply to either the field owould reverse the magnetic attraction and repulsion and thus revarmature. However, such switching would be complex and it imotor where it is necessary for the motor to rotate in either direc

206. In a split-field motor there are two sets of field windings,on a common pole (or core) or on alternate poles around the insicontrolled by a single-pole double throw (SPDT) switch that canWhen the switch is placed in the mid-position the supply to theone direction one set of field coils is energised and the motor rotWhen thrown in the opposite direction the alternate field coilsreversed and the motor rotates in the anti-clockwise direction. A at Figure 7-53.



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FIGUSchemDiagraSplit-F

r also uses series wound motors, thee wound with thick, low resistanceor so employ a remotely operated

field coil and in these the directionply. Since only the armature supplystems often use permanent magnet

Electrics-DC


RE 7-53atic Circuit m for a ield Motor

207. Since the type of applications requiring a reversible motofield coils are in series with the armature and must therefore bwire. Reversible motors with a rating higher than 20 amps switching relay in the supply to the field coils.

208. Some low powered DC motors use a magnet instead of aof rotation is reversed by reversing the polarity of the DC supneed be reversed the switching is simple. Light aircraft flap syreversible motors.



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t aircraft electric motors have a highrably higher rotary speeds than inconsequently a tendency for aircraftally necessary to dissipate the heature diameters are deliberately kept

onditions a cooling period is oftenfor intermittent duty only and theseoperated continuously. Continuouse type of duty for which a motor is, on a rating plate attached to the

e the armature, the yoke (or casing)

the motor shaft, with the armaturented on the armature shaft are the connected. The armature shaft isthe ends of the motor casing.

and form two poles fitting closely

Electrics-DC


Aircraft DC Motors209. In order to keep weight to a minimum it is important thapower-to-weight ratio. To achieve this they operate at consideindustrial applications and at higher armature currents. There is motors to operate at high temperatures and air-cooling is usugenerated. To keep centrifugal forces as low as possible armatsmall.

210. When DC motors are operated under emergency load cnecessary before they can be used again. Some motors are rated will overheat and the internal insulation may burn if they are duty motors are necessarily of a lower power-to-weight ratio. Thrated is listed in the manufacturer’s specifications and, possiblymotor casing.

DC Motor Construction211. The major components of a typical aircraft DC motor arand the field coils.

212. The armature is typically a soft iron drum mounted on conductors set axially into the surface of the drum. Also moucommutator segments, to which the armature conductors aremounted in ball bearings at each end, the bearings being held in

213. The field windings are attached to the inside of the yokearound the armature with a running clearance of about 2.5mm.



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tor is at Figure 7-54.

FIGUSectioDC M(Schem

Electrics-DC


214. A schematic drawing showing a section through a DC mo

RE 7-54n Through a otor atic)



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doubles as the engine starter motor.ine start, when the starter generator. This permits a high current flowing the engine.

switched out and a shunt windingd DC generator supplying current to

arious items of electrical equipment

FIGUSingle DistribSystem

Electrics-DC


Starter Generator215. Some small turbine-powered aircraft use a generator thatTypically the machine has two sets of field windings. During engis acting as a motor, a low-resistance series field winding is usedthrough the field winding to give the high torque needed for turn

216. Once the engine is operating the series field winding is energised in its place, so that the machine becomes a shunt-wounthe aircraft’s electrical system at 28 volts and up to 300 amps.

DC Power Distribution217. In any aircraft, electrical power will be distributed to the vby one of two methods.

RE 7-55Pole ution



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d on aircraft of metal construction.ower supply to the equipment, and aircraft structure itself, as shown atthat are not constructed entirely ofer supply to the equipment, and thea second, earth wire, as shown at

FIGUTwo PoDistribSystem

ich is simply a copper strip acting asowever, distribution is complicated

of a power source or faults in the

nd the battery to be connected to aare drawn, through fuses or circuitematic is illustrated by Figure 7-57.engine starter circuit, a bus bar withection and protection for the radioattery and master solenoid, starter

Electrics-DC


218. The single pole (unit pole) or ‘earth return’ system is useIn this system one wire (the single pole) connects the electrical pthe return path from the equipment to the power source is via theFigure 7-55. The dipole or two-pole system is used on aircraft conductive materials. Here one wire connects the electrical powreturn path from the equipment to the power source is via Figure 7-56.

RE 7-56le (Dipole) ution

219. All aircraft electrical systems use at least one bus bar, wha junction for the generator(s), battery and the various loads. Hby the need to provide for abnormal conditions such as loss distribution system.

220. In single-engine aircraft it is normal for the alternator asingle bus bar, from which the relatively few consumer services breakers of suitable rating. A simple DC power distribution schThe system consists of a battery circuit, an alternator circuit, an circuit breakers and lighting circuits. Not shown are the conncircuits. High current carrying cables are connected between b



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bus bar. All the normal loads are alternator output is indicated by a by individual circuit breakers (CB).s practical and well insulated. Noteternator is running and connected tolied to the consumers. This systemn.

Electrics-DC


solenoid and starter motor and between alternator and main supplied by the alternator once on and running. Failure of thewarning light. The loads are fed from the main bus and protectedAny wires that are not protected by CB’s should be as short athat this schematic is a negative earth return system. When the althe bus, the ammeter will indicate the electrical load being suppshows the option to incorporate an external DC power connectio



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FIGU

Electrics-DC


RE 7-57



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raft, the only remaining source oftrical loading to a minimum and totaneously a landing should be madefor thirty minutes flying following

ne it is most likely to be an electricalg various circuits in turn by use of to turn off the generator or batteryble.

ss the total power demand exceeds

have the least possible effect on

t the power supply to the remainder

generators. The second and third be isolated from the distributionm disruption to services is ensuredritisation of consumer services into

Electrics-DC


221. If the generator fails in flight on a single engined aircelectrical power is the battery. It will be necessary to reduce elecland as soon as practical. If the generator and battery fail simulas soon as possible. A battery should provide enough power failure of a generator.

222. Should the presence of fire be detected other than an engiproblem. The source could be identified by a process of isolatincircuit breakers. On a single engined aircraft it may be necessary(or both). Once again a landing should be made as soon as possi

223. Typical requirements of an aircraft electrical system are:

(a) Supply to all equipment must be maintained unleavailable supply.

(b) System faults (earths, short-circuits, etc) shouldoverall system functioning.

(c) A fault on one piece of equipment should not affecof the system.

224. These requirements may be met by parallel operation ofrequirements are achieved by arranging for faulty systems tonetwork by means of fuses and/or circuit breakers. The minimuby providing each power source with its own bus bar, and priothree categories; vital, essential and non-essential.



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ency when all main power sources,, which include emergency lightingected directly to the aircraft battery

e flight in an emergency situation.her from a generator or the aircraft

e safely disconnected during an in-ied from the generator bus bars.

DC distribution system designed to

Electrics-DC


225. Vital consumers are those services required in an emergnormally the engine-driven generators, are lost. These servicesand fire detection/protection, are provided from a bus bar connsupply.

226. Essential consumers are those services necessary for safThey are connected to a bus bar that can always be supplied eitbatteries.

227. Non-Essential Consumers are those services that can bflight emergency, for purposes of load shedding. These are suppl

228. Figure 7-58 shows a basic twin-engined aircraft bus bar conform to the above categories.



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FIGUDC BuDistrib

Electrics-DC


RE 7-58s Bar ution



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DC to AC, which is necessary for (avionics in modern jargon). Theh a combining bus bar. In the eventer, and the remaining generator can

power (at 24v) is always available.by either generator and/or Batterieses are supplied from the Generatorrtant and makes load shedding easy at a frequency of 400 Hz.

y are used to convert direct currentter, DC is used to drive a DC motoror) to provide alternating current at

parts and achieve the same result. The circuitry of the static inverteritors and transformers, all of whichm an oscillator circuit that convertsic inverters are usually designed to

ter with switching to select loade of suitable rating protects each ofe alternators are typically connected

Electrics-DC


229. Note that the system includes inverters. These convertmany of the flight instruments, radios, and navigation systemssystem is designed for parallel operation of the generators througof failure a generator can be isolated, by opening its circuit breaksupply all services.

230. Vital services are taken from the Battery Bus Bar, whereEssential services are supplied from the Combining Bus Bar, fed and therefore not subject to interruption. Non-Essential servicBus Bars and could be subject to interruption, but this is unimpowhen necessary. The inverters would, typically, supply 115v AC

231. Inverters may be either rotary or static. In either case the(DC) to alternating current (AC). In the case of the rotary inverat constant speed. This in turn drives an alternator (AC generatconstant frequency (usually 115 volt, 3 phase AC at 400 Hz).

232. Static inverters, as their name suggests, have no movingelectronically. They are much more common in modern aircraftcontains such electronic components as diodes, transistors, capacare explained in later sections. These solid-state components forDC input into a 400 Hz constant frequency AC output. Statproduce single phase AC.

233. System monitoring typically comprises a single ammemonitoring of one alternator at a time. A circuit breaker or fusthe consumer services at their connections to the bus bars and thto their bus bars through current limiting fuses.



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istribution system thus allowing ally or generating systems. The sourcen alternative at some airports is toal. It is important that the unit is ofg is not exceeded. When connecting off first. An aircraft should not beying the aircraft systems.

tch, situated in the cockpit, operatese starter switch operates a remotelyre widely used in aircraft electrical heavily loaded circuits, but also for

systems in twin-engined aircraft theway, but the generator buses must

the generators are not equipped toportant that supply to all essential

witch called a bus tie breaker (BTB)s-connect them when it is closed.

pen.

Electrics-DC


234. A ground power source can be connected to the bus bar delectrical systems to be powered independently of aircraft battercan be either a motorised generating unit or a battery unit. Aconnect up to a cable from a ground supply routed to the disperscorrect voltage and polarity and that maximum amperage loadinor withdrawing the connection it is advisable to switch the unitleft unattended if a ground power supply is connected and suppl

Elementary Switching Circuits235. In Figure 7-57 it will be noted that the battery master swia remotely located solenoid-operated switch. Similarly, the enginlocated relay. Relay- and solenoid-operated switching circuits adistribution systems, not only for remote operation of switches insequential switching functions.

236. For example, in non-paralleled (split bus bar) distributiongenerators are connected to their own bus bars in the normal never be cross-connected with both generators operating, sinceoperate in parallel. However, should either generator fail it is imservices be maintained without interruption.

237. A simple split bus bar system is shown at Figure 7-59. A sis situated between the two generator buses, which will crosHowever, the BTB can only be closed if one or both GCB’s are o



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FIGUSplit BSystem

uit breakers (GCB) are closed, but ited, keeping both bus bars energisede this is illustrated at Figure 7-60.

Electrics-DC


RE 7-59us Bar

238. The bus tie breaker is held open when both generator circwill automatically close if either generator circuit breaker is openand maintaining electrical services. A switching circuit to achiev



Chapte

FIGUSwitchSplit BSystem

Electrics-DC


RE 7-60ing Circuit - us Bar



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spective GCB solenoids. When aropriate GCB solenoid, which closeste double-pole switches, which also

rator switch is opened its associatedfrom its bus bar and connecting thes-connecting the generator bus bars.

ids will be de-energised and bothbe energised, cross-connecting thengines shut down and the aircraft

their bus bars the BTB solenoid isolated from each other.

re as follows:

ghting. Lights can be powered fromr from aircraft or separate batteries., anticollision lights, landing lights,

ts and emergency lights.

r de-icing/ anti-icing of airframe,tall and warning devices. Some airlement heating.

Electrics-DC


239. The generator switches in the cockpit operate their regenerator switch is closed a 28-volt DC supply energises the appto connect the generator to its bus bar. The GCB solenoids operacontrol a 28-volt DC supply to the BTB solenoid. If either geneGCB solenoid will be de-energised, disconnecting the generator 28-volt DC supply to the BTB solenoid, closing the BTB and cros

240. If both generator switches are open, both GCB solenogenerators isolated from their buses. The BTB solenoid will generator bus bars. This would be the situation with both econnected to an external power supply.

241. When both generators are operating and connected to isolated from its 28-volt supply and the generator bus bars are is

Electrical Consumers242. Some typical electrical consumers (loads) and their uses a

(a) Lighting. Most aircraft will have some form of lithe aircraft main generating system (DC or AC) oAmongst these will be position lights (Nav lights)instrument panel lights, warning lights, cabin ligh

(b) Heating. Electrical heating circuits are used fopropellor, engine, windscreen, pitot probe and sconditioning systems use electrical supply to supp



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operate various mechanisms usinghes. Some indication systems are

encompass a variety of electronicmmunication and navigation radio,ight management systems.

not be possible without the use ofich are electrically or electronicallyture, altitude, velocity and rates ofhave a commonality.

ter or lesser extent, acquire electro-ent sections of the airframe acquire, and sparking (arcing) across smallio interference and at worst it couldhe airframe are electrically bondednce path to discharge points on the

ch are copper strips extending fromying control surfaces.

Electrics-DC


(c) Magnetic Devices. Electrical supply is used to electro magnets i.e. solenoids, relays and switcmagnetically operated i.e. MIs (dolls eye).

(d) Avionic Systems. Aviation electronics (avionics) systems. Avionic systems in aircraft can include, coautopilot, weather radar, inertial navigation and fl

(e) Instruments. Operation of modern aircraft wouldinstruments. Large aircraft have instruments whoperated. Instruments measure pressure, temperaflow. Navigation instruments and auto flight will

Bonding and Screening243. An aircraft flying through the atmosphere will, to a greastatic charges in the metallic structure of the airframe. If differdifferent electrical potentials then current will flow between themgaps in the structure is liable to occur. At best this will cause radlead to fires. In order to prevent this the individual parts of ttogether, using woven copper wire strips to provide a low resistastructure.

244. In flight these points are the static wick dischargers, whipoints of static concentration such as trailing edges of primary fl



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ugh conducting tyres which have ah any accumulated static electricity

e only fitted to the smaller wheels,

roduce electro-static charges. Thisling when the atmosphere is dry,00 gallons per minute or more. It is

e bowser and aircraft are connectedto earth. In overwing re-fuelling aded to the aircraft tank filler pipe,ves the nozzle.

terference as a result of sparking in by fitting suppressors in the cablesheath around the cables. Ignition equipment making and breaking ather sheathed or suppressed.

of capacitors across the source ofthe stray voltages induced by thent flow.

Electrics-DC


245. On the ground the aircraft is earthed to the tarmac throhigh proportion of carbon in the tyre material and through whicis discharged. Since these tyres are relatively expensive they arnormally the nose wheels.

246. Friction due to the flow of liquid in a pipe can also paccumulation of static is particularly marked during re-fuelespecially in pressure re-fuelling when the rate of flow may be 1,0therefore essential that before pressure re-fuelling commences thelectrically (bonded) to each other, and that both are bonded further requirement is that the re-fuelling nozzle should be bonsince the static electricity may be induced in the fuel flow as it lea

247. Screening is incorporated on aircraft to prevent radio inelectrical components. The radio interference can be suppressedattached to any source of sparking, and by installing a metal systems, DC generator and motor commutators, and indeed anycircuit (especially at frequencies in excess of 10Hz), need to be ei

248. Suppression is usually achieved by connecting a numberinterference, so that they provide a low resistance path for fluctuating magnetic fields associated with interruptions of curre



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FIGU

Electrics-DC


RE 7-61




System (VSCF)


Electrics-AC

Frequency

Sine Wave Format

RMS Equivalent

Inductance

Inductive Reactance

Capacitance

Capacitive Reactance

Impedance

Single and Three Phase Supplies

Remote CSDU Disconnect

Variable Speed Constant Frequency

Integrated Drive Generator (IDG)

Real Load Sharing





Reactive Load Sharing

Monitoring Equipment

Transformer Function

Types of Transformer

Transformer Rectifier Units (TRUs)

Types of AC Motor

The Induction Motor



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rnating current power distributionsult in very large current flows if the. Because it is very easy to change high voltage (usually 115v AC) andary. Avionic systems incorporatingcitors require AC power for their

wered. Even in light aircraft somently, the study of alternating current

iodically changes in direction andthe chapter covering generation ofnges in electrical polarity every half

gative, this means that the currentThe magnitude of the voltage (emf) flow in a supplied circuit, varies as°.

Electrics-AC


8Electrics-AC

Alternating Current (AC)1. Virtually all large modern commercial aircraft use altesystems. The large power requirements of such aircraft would relow voltage (usually 28v) of DC distribution systems were usedvoltages with AC it can be generated and distributed at relativelyreduced to lower voltage for conversion to DC where necesscomponents such as transistors, transformers, diodes and capaoperation. A high percentage of aircraft lighting is also AC poalternating current is used, produced by inverting DC. Consequeis an essential requirement for pilots.

Alternating Current2. Alternating current is defined as current flow that percontinuously changes in magnitude. As has been shown in electricity the emf produced in the armature of a generator charevolution of the armature.

3. Since current (conventionally) flows from positive to neflow reverses in direction every half revolution, or half cycle. induced in the armature, and the resulting magnitude of currentthe sine of the degree of rotation of the armature, from 0° to 360



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graphically follows a sine curve, or

FIGUProducSine W

Electrics-AC


4. Thus, the magnitude of voltage and current, when shownsine wave. This is illustrated at Figure 8-1.

RE 8-1tion of AC ave



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cle, of the generator armature. Onend/or current increases from zero to repeats the process with oppositeak voltage/current rise is known as

e repeated, or cycles per unit time, is

measured in units called hertz (Hz). cycles per second. Since each cycle (see Figure 8-1), it follows that thency. For a ‘two-pole’ generator asrly need to rotate at 400 revolutions

lutions per minute (rpm) and, sinceeed to be:

Electrics-AC


5. Each complete sine curve represents one revolution, or cycycle of alternating current covers a period in which the voltage amaximum value in one direction, falls back to zero and thenpolarity and direction of current flow. The magnitude of the pethe amplitude of the cycle. The rate at which the cycles of AC arknown as the frequency of the AC

Frequency 6. AC frequency is the number of cycles per second and is In aircraft AC systems the standard frequency is 400 Hz, or 400of AC is the result of one revolution of the generator armaturegreater the speed of rotation, the higher the output AC frequeshown (one North Pole, one South Pole) the armature would cleaper second to produce a frequency of 400 Hz.

7. Machine rotary speed is conventionally measured in revothere are 60 seconds per minute the rpm of the armature would n

In other words:

Therefore, by transposition:

400 60× 24 000 rpm,=

FN60------ for a 2-pole generator=



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c poles, since the more the ‘pairs’ ofature necessary for a given output poles and two south poles, spacedotary speed required for an outputa for calculating AC frequency must

obviously twice the number of pole

t Figure 8-1 and from this it can beed by a rotating armature, varies

Electrics-AC


8. Most practical AC generators use more than two magnetipoles (North and South) the lower the rotary speed of the armfrequency. If the generator shown at Figure 8-1 had two northevenly around the armature in the sequence N, S, N, S, the rfrequency of 400 Hz would be halved. Consequently, the formulinclude the number of pairs of poles (P) and is given as:

9. In some cases the number of poles is given. Since this is pairs the formula then becomes:

or

Sine Wave Format10. The geometric production of a sine curve is illustrated aseen how it is that the output of an AC generator, producsinusoidally with time.

FN P×

60-------------=

FN P×60 2×---------------= F

NP120---------=



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FIGUGeomProducCurve

ors are constructed with a rotatinghe casing (stator).

primary supply sources are thatr brush assembly, requiring less

over the operating speed range.

lex additional exciter mechanism iseparate excitation from the aircraft

er direct current when it comes toers how its voltage can be readilyload current can be decreased if

on significantly reduced as a result.eight.

Electrics-AC


RE 8-2etric tion of Sine

11. As has been shown already, most practical AC generatmagnetic field (the rotor) and a stationary armature winding in t

12. The advantages of alternators over DC generators asalternators (AC generators) are lighter and have a simplemaintenance. Also, their output voltage remains more constant

13. The main disadvantage lies in the fact that, unless a compincluded in the AC generator construction, it is reliant upon sbattery.

14. Alternating current has some significant advantages ovdistribution. It will be explained in the text covering transformtransformed either up or down. Thus, for a given power, distribution voltage is increased and power losses in transmissiConducting cable diameter can therefore be reduced, saving in w



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rrents, radio interference and extra to the weight saving advantages,

al supply it is necessary to know theard since both are easily measured.nging sinusoidally and it becomes

of the square of the current gives arrent. This value is 0.707 (1/√2) of

hich it is determined it is known ascan be determined in the same way.

ms equivalent value. Given that:

ak voltage = 115 x 1.414 = 163v.

1×--------

2 1.414( )

Electrics-AC


15. The induced voltages caused by AC can lead to stray cugenerator loads, but these disadvantages are small comparedespecially in large aircraft with high power requirements.

RMS Equivalent16. In order to determine the power available from an electriccurrent and voltage of the supply. With DC this is straightforwWith AC, however, the value of current is continuously chanecessary to determine a mean effective current.

17. It has been found that the square root of the mean valuevalue that has the same heating effect as an equivalent direct cuthe peak value of alternating current. Because of the method by wthe root mean square (rms) equivalent value. The mean voltage

18. The current and voltage quoted for an AC system is the r

19. Thus, for an AC circuit with an RMS voltage of 115v, pe

RMS Equiv. ValuePeak Value

2----------------------------=

(Peak Value RMS Equiv. Value)= =



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ses the current flow in a DC circuit.t only by resistance, but also by

agnetic field is set up around theetic effect can be used to magnetise

inciple that is applied to solenoids,h of DC it is possible to control therol the output of a generator or the

rrent in a direction that opposes the alternating current, the current isetic flux continuously varying in

ate of movement of which depends

g circuit and that emf circulates aords, opposing the normal currentave the property of self-inductanceduced in the circuit. All AC circuits of self-inductance is the Henry (H)

Electrics-AC


Opposition to Current Flow20. Only the resistance of the circuit and its components oppoIn an AC circuit, however, current flow may be opposed noinductance and capacitance.

Inductance21. Whenever there is a flow of DC or AC current a mconductor. If the conductor is wound into a coil the electromagn(DC), or de-magnetise (AC) magnetic materials. This is the prrelays and so on. These devices use DC. By varying the strengtstrength of a magnetic field. This same principle is used to conttorque of a DC motor.

22. Lenz's Law states that induced emf acts to circulate a cuchange of flux which causes the emf. In a conductor carryingcontinuously varying in magnitude, which produces a magnstrength. In effect, a moving magnetic field is produced; the rupon the frequency of the alternating current.

23. This moving field induces an emf in the current-carryincurrent in a direction opposing the change of flux - in other wflow. This is known as self-inductance and a circuit is said to hwhen a change in the current in that circuit causes an emf to be inare self-inductive. The symbol for self-inductance is L. The unitwhich is defined as follows:



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1 volt is induced in the circuit by a

the original, or main, current flowut this is usually negligible. As wasctor is formed into a coil a much have an inductance that produces a

rrent change is greatest. With AC, (changing direction) and least whenen current is zero and least whenich it will be seen that the effect of

FIGUEffect Induct

Electrics-AC


24. A circuit has a self-inductance of one Henry if an emf ofrate of change of current of one ampere per second.

25. Because this emf produces a current flow in opposition toit is known as back emf. A straight conductor has inductance, bshown in the chapter covering electro-magnets, when a condustronger magnetic field is produced. Coils of wire in AC circuitssignificant back emf, opposing the applied current.

26. According to Lenz's Law, induced emf is greatest when curate of current change is greatest when it is passing through zeroat its peak value. Hence induced emf (voltage) is greatest whcurrent is maximum. This is illustrated at Figure 8-2, from whinductance in an AC circuit is to cause current to lag voltage.

RE 8-3of ance



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90° as shown and this is known ase there is always resistance as well,

nductive reactance and is given theby inductors, whereas inductance is Because it impedes current flownal to the inductance of the circuit

ductive reactance is:

ductance (H)

iceable when the current is switchedle problem, especially if frequency isue to variation in frequency, causessensitive. Frequencies above, andductive components, because of the

Electrics-AC


27. In a purely inductive circuit, current would lag voltage byphase quadrature. Pure inductive circuits are non-existent, sincand the phase lag between current and voltage is less than 90°.

Inductive Reactance28. The effect of inductance in an AC circuit is known as isymbol XL. It is the actual opposition to current flow created the ability of an inductor to oppose changes in current flow.inductive reactance is measured in ohms. It is directly proportioand the frequency of the alternating current. The formula for in

where: F = Frequency (Hz) L = In

29. In DC circuits the effect of inductive reactance is only noton or off. In AC circuits inductive reactance can be a consideraballowed to vary. The variation in opposition to current flow, dvariations in voltage to which avionic equipment is highly especially below, design frequency can lead to overheating in inincrease in current flow caused.

The equation for Ohm's Law states that:

XL 2πFL ohms=

IVR----=



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ber) it is apparent that a reductionn increase in current has a heating

ill cause an increase in current flow,ltage regulator ensures it does).

constant value, a decrease in ACverheating in inductive devices such

the section dealing with the Basic

osite effect to that of an inductor, in8-3.

Electrics-AC


30. If we substitute XL for R (both measured in ohms, rememin XL at constant voltage will cause an increase in I (current). Aeffect upon any conductor.

31. Since XL = 2πF.L, it follows that a decrease in frequency wprovided that voltage remains constant (which, of course, the vo

32. Clearly, since distribution voltage is maintained at a frequency will cause an increase in current flow. This can cause oas transformers, motors and generators.

Capacitance33. The basic principle of the capacitor was described in Principles of DC Electricity.

34. When a capacitor is placed in an AC circuit it has the oppthat it causes the current to lead the voltage, as shown at Figure



Chapte

FIGUEffect Capac

electric stress set up in the capacitorltage has reached its peak value theltage begins to fall current flows outr stored voltage is higher than the

is maximum because there is now circuit, current would lead voltage to some extent so this theoretical

s voltage in a capacitive circuit andul, where:

Electrics-AC


RE 8-4of itance

35. As voltage rises the capacitor becomes charged and the diopposes current flow, so current flow decreases. By the time vocapacitor is fully charged and current flow is zero. As supply voof the capacitor in the opposite direction because the capacitosupply voltage.

36. By the time supply voltage has fallen to zero, current flowno opposition to current flow. If there were no resistance in theby 90° (phase quadrature). Of course, all circuits are resistivecondition is never reached.

37. To help the student remember the fact that current leadlags voltage in an inductive circuit, the mnemonic CIVIL is helpf

(a) C = Capacitive

(b) I = Current



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an extremely large unit capacitorsfarads. One farad is defined as thetrical energy in the capacitor.

itive reactance. Because it opposesr capacitive reactance is XC, whichitors in an AC circuit. The formula

upply will affect the current flow in

Electrics-AC


(c) V = Voltage

(d) L = Inductive

38. The unit of capacitance is the farad (F). Since this is usually have values expressed in microfarads, nanofarads or picocapacitance present when one volt will store one coulomb of elec

Capacitive Reactance39. The effect of capacitance in an AC circuit is called capaccurrent flow in a circuit it is measured in ohms. The symbol forepresents the actual opposition to current flow caused by capacfor capacitive reactance is:

where: F = Frequency (Hz). C = Capacitance (Farads)

As with inductive reactance, a variation in frequency of the AC sthe circuit.

From Ohm’s Law:

therefore, in a capacitive circuit:

XC1

2πFC-------------- ohm=

IVR----=



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t constant voltage and capacitance,

ance XC is equal to the inductive

actance, capacitive reactance andver, impedance is not the sum of thectance have opposite effects, as has

between current and voltage so itsce. Consequently, impedance must

lustrated at Figure 8-4.

Electrics-AC


40. From this it will be seen that an increase in frequency, awill cause an increase in current flow.

Resonance41. A resonant circuit is one in which the capacitive reactreactance XL.

Impedance42. In an AC circuit the combined effect of inductive reresistance is known as impedance and has the symbol Z. Howethree. In the first place, inductive reactance and capacitive reabeen shown. Total reactance is the algebraic sum of the two:

43. Secondly, resistance does not cause a phase difference effect will be 90° ahead of inductance and 90° behind capacitanbe the vector sum of resistance and reactance. A vector sum is il

IV

XC------- V 2πFC×= =

XT XL X– C( )+=



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FIGUVectorImped

ircuits that include inductive and/orwith voltage (where there is a backalso contain some purely resistivel opposition to current flow, caused

e.

ce represents useful or real power, work done in overcoming reactancedone (power generated) will have tosystem load and the reactive (idle orf effective power to apparent power

Electrics-AC


RE 8-5 of ance

Power in AC Circuits44. Reactance and reactive load are the terms applied to AC ccapacitive components, causing current flow to be out of phase emf, or reactance, opposing current flow). All circuits will components and these, too, oppose the flow of current. The totaby inductance/capacitance and resistance, is known as impedanc

45. The work done by the generator in overcoming resistansince resistance is due to the load of the circuit components. Theis wasted effort and is known as reactive power. The total work be that necessary to overcome both the real (effective or active) wattless) load. This is known as the apparent power. The ratio ogives the power factor (PF) of the circuit being supplied.



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apacitive reactance in a circuit willt of phase with) current. If voltageseful work will be maximum, since

er the reactive load and the less thef circuit voltage and current thatle (φ) between voltage and current is

termining the power rating of thesystem to be supplied. In order toy the entire AC circuit, from real ormperes (kVA). AC generators are

of real power. The reactive load,peres reactive (kVAR).

flow due to the resistance of the AC

of work or heat and is measured in

r--

Electrics-AC


46. As we have already seen, the amount of inductive and cdetermine the extent by which circuit voltage lags or leads (is ouand current are in phase with each other, power available for upower (watts) = volts (V) x amps (I).

47. The more voltage and current are out-of-phase the greatuseful power available. Thus, it is the phase relationship odetermines the circuit power factor. The cosine of the phase angequal to the power factor:

48. Knowledge of the power factor is necessary when degenerator(s) needed to meet the real and reactive loads of the distinguish apparent power, which is the total power consumed breactive power, apparent power is quoted in units of kilovolt arated in kVA rather than kilowatts (kW), which is the unit sometimes displayed in larger aircraft, is measured in kilovolt-am

49. Real Power is the power resulting purely from the currentcircuit, and is given by the formula:

50. It is the power consumed usefully in the circuit in the formWatts (W) or kilowatts (kW).

PF Cosφ Real PowerApparent Powe-------------------------------------= =

Power (Watts) I2R=



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t flow due to the total reactance is given by the formula:

ps reactive (kVAR).

it that contains reactance as well as the impedance of the circuit.

ly as the horizontal and verticalent power is the vector sum of the

L XC– )

Electrics-AC


51. Reactive Power is the power resulting from the curren(inductive reactance – capacitive reactance) of the AC circuit and

52. It is measured in volt amps reactive (VAR) or kilovolt am

53. Apparent Power is the power consumed by an AC circuresistance. It is the power resulting from the current flow due to

54. It is measured in volt amps (VA) or kilovolt amps (kVA).

55. When real and reactive power are plotted respectivecomponents of a vector diagram, as at Figure 8-5 below, appartwo.

Power (Kilovolt Amps Reactive) I2

X(=



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FIGUAC Po

ound, since:

er is known, or vice versa, because:

Theorem, or by triangulation.

Electrics-AC


RE 8-6wer Vector

56. If the phase angle (φ) is known, the Power Factor can be f

57. And from this Real Power can be found if Apparent Pow

58. Reactive power can be found either by using Pythagoras’

PF Cosφ=

PFReal Power

Apparent Power---------------------------------------=



Chapte

EXAM

stem is 30°. When the apparentactive power.

Power Apparent Power PF×=

Real Power)2

72 950.24,

0 kVAR

Sinφ×

Electrics-AC


PLE 8-1EXAMPLE

The phase angle (φ) between current and voltage in an AC sypower consumed by the system is 540 kVA, find the real and re

SOLUTION

Using Pythagoras:

Using triangulation:

PF 30°cos 0.866= = PFReal Power

Apparent power----------------------------------------- Real∴=

540kva= 0.866×

467.6 kw=

Reactive Power( )2 Apparent Power( )2= (–

Reactive Power( )2 5402

467.62

–=∴ =

Reactive Power 72 950.24, 27= =∴

Reactive Power Apparent Power=

540 0.5×=

270 kVAR=



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and reactive power is greater than

AC circuit (real power) the voltagetance (reactive power), the voltage the phase difference into account two.

apacitive effects are opposite theyrging capacitor can compensate forrcome the current lead caused byis achieved a state of resonance ishase with voltage and the circuit

hich are inductive. To cancel theirto completely achieve a state ofin additional current (load) being

Electrics-AC


The observant student will have noticed that the sum of real apparent power.

This is because, when considering the effect of resistance in an and current are in phase, but when considering the effect of reacand current are out of phase. Hence the total effect must takeand the total, or apparent power must be the vector sum of the

AC circuits are rarely purely resistive, but since inductive and ccan be used to compensate for each other's reactance. A dischathe current lag caused by an inductor, or an inductor can ovecapacitor discharge. When an exact balance between the two said to exist. Under these circumstances current will be in pappears to be purely resistive.

Most aircraft AC equipment employs electro-magnetic coils, wreactance, capacitors are introduced. It is never possible resonance and any inductive reactance remaining will result carried wastefully by the alternators and distribution cables

467.6 270+ 737.6=



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an aircraft, plus low-powered ACperation, powering DC motors andgle wire supply and earth-returnase operation.

ng magnetic field principle, whichny avionic components also operater phases, to be produced by a three-

dependent armature windings in theach winding operates like a single-lternatively, the outputs can be usedstem is shown diagrammatically at

Electrics-AC


Single and Three Phase Supplies59. AC electricity is used to power most of the lighting onmotors. It is also converted to DC for battery charging, relay oother DC loads. This type of equipment operates on a sindistribution system from the alternator and is known as single-ph

60. However, heavy duty AC motors operate on a rotatirequires three AC supplies at equally spaced time intervals. Maon similar principles, so it is normal for the three AC supplies, ophase, or polyphase, alternator.

61. Three phase alternators have three equally spaced and instator casing, to give the required outputs at 120° intervals. Ephase alternator, and can be used to supply single-phase loads. Atogether for three phase (or two phase) loads. A three-phase syFigure 8-6.



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FIGUThree System

out to individual pairs of terminalsve use of cable. Instead, the three-nection as shown at Figure 8-7 and

Electrics-AC


RE 8-7Phase

62. In practice each of the three phases is not usually broughtas at Figure 8-6, since this would involve unnecessarily excessiphases are interconnected by means of either a star or delta conFigure 8-8.



Chapte

FIGU3 PhasConne

hich is the norm for aircraft ACmon junction known as the star, orrn, or neutral line from the single-

ted to a distribution, or live line.

een from Figure 8-7 that the single-windings. The voltage across these live line and neutral. It is typically

Electrics-AC


RE 8-8e Star ction

63. Figure 8-7 illustrates the star or wye connection, wgenerators. One end of each phase winding is connected to a comneutral point. Also connected to this point is the common retuphase AC loads. The other end of each phase winding is connec

64. The live lines supply the single-phase loads, so it will be sphase loads are connected in parallel with the generator phase loads is known as the phase voltage and it is the voltage between115v (rms).



Chapte

indings connected, but these phasee voltage is equal to phase voltage xld be 115 x 1.73, or 200v. In a star

d phase currents are equal, becauseine connected to it.

hase loads, so they are connected attempts to ensure that, in normaled between the three-phases.

urrent flow in the return line, equalple, if phase A load was 100 amps,um of these current flows would be

each other so the currents flowing instant of time.

umed in phase A loads, the currentre 8-7) and will be from neutral tot can be shown mathematically as

a

a

a

Electrics-AC


65. Between any pair of lines there are two generator phase wvoltages are in opposition and out-of-phase, so line-to-line, or lin√3. Since √3 = 1.73, if phase voltage were 115v, line voltage wouconnected or star wound generator distribution system, line anthe current in any phase load is clearly flowing through the live l

66. The majority of the loads on the generator are single-pbetween a line and neutral. In an aircraft system the designeroperation, the total aircraft loads are approximately equally shar

67. When the phase loads are not equal there will be a small cto the phasor sum of the current flows in the live lines. For examphase B load 120 amps and phase C load 110 amps, the phasor sonly 15 amps. This is because the three phases are at 120° tothrough the each of phase loads are in different directions at any

68. If a current flow from line to neutral of 100 amps is assflows in phase B and C loads will be less than full load (see Figuline. The values and direction of these currents at this instanfollows:

Phase A 100a Cos× 0° +100=

Phase B 120a Cos× 120° 60–=

Phase C 110a Cos× 240° 55–=



Chapte

from the star point (at this instant).and hence the reason for using thed by geometry.

e current flow in the neutral returnh. In an aircraft distribution systempment are switched on and off, so a

phase of a star wound three-phaseere will be no current flow in that than before, hence there will be ano cause a voltage drop in the twomagnitude of the phase loads.

r between a phase and neutral (line-rop to near zero. The high current

e three-phases will always be equal,s an alternative form of three-phase

Electrics-AC


69. Phasor Sum = 100 + (-60) + (-55) = -15amps or 15amps Students will appreciate that this is a trigonometrical solution cosine of the angles. The problem could equally have been solve

70. If the three individual phase loads were always equal, thwould always be zero and the return line could be dispensed withowever, the loads in the three-phases will vary as items of equineutral return is necessary.

71. Should an internal break lead to open circuiting of one generator the voltage across that phase will fall to zero and thphase. The phasor sum of the remaining currents will be higherincrease of current flow in the neutral return. This is liable tremaining phases, the magnitude of which will depend upon the

72. Short-circuiting between generator phases (line-to-line) oto-earth) will cause very high current flow and the voltage will dflow causes overheating of the stator windings.

73. As previously stated, if it can be assured the loads on ththere is no need for a fourth, return, wire. In these circumstanceconnection called a delta, or mesh connection can be used.



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FIGU3 PhasConne

ich it can be seen that voltage acrossnected to the opposite ends of thatn such a connection, line current isected to the ends of two phases so, greater than the current flow in one

ystem) compared to a DC generator. produce maximum circuit voltageommutation the problem of brusht supply is available for consumer

types, depending upon the system toontrolled generators. In either casee-phase output voltage at 115v, and

Electrics-AC


RE 8-9e Delta ction

74. The delta connection is illustrated at Figure 8-8, from whone phase will be the same as the voltage between the lines conphase. In other words, line voltage is equal to phase voltage. Iequal to phase current x √3. This is because each line is conngiven a balanced load situation the current flow in a line must bephase (phase current).

AC Generators (Alternators)75. There are several advantages of using an alternator (AC sThe alternator has a much better power/weight ratio and canrequirement at low rpm. Because there is no requirement for csparking is eliminated and in addition a wider range of currendevices.

76. AC generators in aircraft conform generally to two basic be served. These two types are frequency-wild and frequency-cthe generator is normally a three-phase machine producing singlthree-phase voltage at 200v.



Chapte

d DC generators, output voltage is field excitation, by varying the field

s are not frequency sensitive. Thesetrolled and whose output frequency variable frequency are known as

to be at a constant value. In orderintained constant, regardless of theystem are known as constant speed,

ft in which the primary distributionrator is only used to power specificibution to the bus bars the output ofd rectified to DC.

xcitation current being fed throughe 6-pole rotor. Generator output is

80 Hz to 400 Hz. The generator istor is at Figure 8-9.

Electrics-AC


77. As we have already seen when discussing alternators ancontrolled by a voltage regulator that controls the strength of thecurrent.

78. Simple resistive circuits, such as electrical de-icing systemcan be supplied with AC from a generator whose speed is unconis therefore variable. The generators used to supply AC atfrequency-wild generators.

79. Most AC equipment requires the frequency of the supplyto achieve this the rotational speed of the generator must be maspeed of the engine driving it. The generators supplying such a sor frequency-controlled generators.

80. Frequency-wild generators are usually only used in aircrasystem is DC. The AC output from the frequency wild AC generesistive systems, such as engine and propeller de-icing. For distrthe generator is transformed to a lower voltage, typically 28v, an

81. Frequency-wild generators are separately excited, DC ebrushes and slip rings to the series-connected field windings on thtypically three-phase, 200v, 22 kVA over a frequency range of 2cooled by ram air. A diagram of a typical frequency-wild genera



Chapte

FIGUFrequeGener

t speed drive units (CSDU's) in theally hydro-mechanical variable ratiof generator or transmission failure.sociated engine shut down.

field, or fields within a stationaryator are used instead of field andpole machine. The magnetic field ist current for field excitation. Sincethe field current, from an externalgs via brushes and slip rings.

Electrics-AC


RE 8-10ncy Wild ator

82. Frequency-controlled generators require complex constantransmission train between engine and generator. These are usudrives that can be remotely disconnected in flight in the event oThey can usually only be re-connected on the ground with the as

83. Alternators (AC generators) usually consist of a rotatingarmature. In order to avoid confusion the term rotor and starmature. This type of arrangement can be termed an internal created by an electro-magnet, which must be supplied with directhe field winding is wound around a rotating core (the rotor) source such as a DC bus bar, must be supplied to the field windin



Chapte

h-power machines producing veryinevitably occur with slip rings andshless AC generator is illustrated at

FIGUBrushlGener

Electrics-AC


84. Most large, frequency-controlled AC generators are higlarge current flows at full load. To eliminate the losses, which brushes, they are brushless machines. The principle of the bruFigure 8-10.

RE 8-11ess AC ator



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se are (1) an exciter-generator that rectifier assembly that converts thetor, producing AC at constant (400

armature at one end and the mainwo, is the rotating rectifier assembly to DC and supplying it to the main

by the exciter stator field windings.. The main generator field rotatesconnected to the generator output

and output frequency is 400 Hz.

r would therefore have to maintainronous speed.

pletely self-contained and requiresammatic form at Figure 8-11.

of poles--------------------

20-------- 6000rpm=

Electrics-AC


85. The generator consists of three major components. Theproduces the field current for the main generator, (2) a rotatingoutput of the exciter-generator to DC, and (3) the main generaHz) frequency for supply to the generator bus bar.

86. The rotor of the generator carries the exciter generator generator field at the other. Mounted on the rotor, between the tsupplied with AC from the exciter armature and converting thisgenerator field.

87. The exciter armature rotates in a magnetic field created These are supplied with regulated DC from an external sourcewithin the three-phase stator (armature) windings, which are terminals.

88. The number of poles in such a generator is usually eight,

89. The constant speed drive unit (CSDU) for such a generatoa constant drive speed of 6000 rpm. This is also known as synch

90. An alternative type of brushless AC generator that is comno external DC supply for its exciter-generator is shown in diagr

Given that frequency (F)rpm No.×

120--------------------------=

by transposition rpmF 120×

No. of poles-----------------------------

400 1×8

----------------= =



Chapte

FIGUSelf CoBrushlGener

rst is a permanent magnet generator alternating current in a stationaryU) and the DC is supplied to thes alternating current in the rotating

d on the rotating shaft, where it ise main generator. Here it induces

ain generator. The output voltage iswhich regulates the field current to

Electrics-AC


RE 8-12ntained

ess AC ator

91. This comprises three generators on the same shaft. The fi(PMG), consisting of a rotating permanent magnet that inducesarmature. This is rectified in the generator control unit (GCstationary field winding of the exciter-generator, where it inducearmature of the exciter-generator.

92. The exciter-generator output is fed to a rectifier mounteconverted to DC for supply to the rotating field windings of thalternating current at 400 Hz in the stationary armature of the mregulated at 115v single-phase, 200v three-phase, by the GCU, the exciter-generator field winding.



Chapte

f AC generators, which is necessaryeration is not required, many AC constant frequency AC. Inductiveonic components will only function

ndent upon its speed of rotation; inintain constant speed of rotation ofx) drives the generator, and engineto introduce a constant speed drivenge from a pneumatically operatedtilising either the wobble pump or

mechanical constant speed drive. Itixed displacement hydraulic unit, a. Hydraulic oil for operation of their and oil cooler.

Electrics-AC


Constant Speed Drive System (CSDS)93. Constant frequency AC is essential to parallel operation oin three- and four-engined aircraft. Even where parallel opcomponents are frequency sensitive and must be supplied withcomponents will overheat if frequency is too low and most aviproperly within a narrow frequency range.

94. The output frequency of an AC generator is totally depeorder to maintain constant output frequency it is necessary to mathe generator. Since the engine (through the accessories gearbospeed is variable from idling to maximum rpm, it is necessary system between gearbox and generator. The types of CSDUs racombined air motor/starter type to the oil controlled system uvariable speed arrangement.

95. Figure 8-12 shows a schematic block diagram of a hydro-is a variable ratio drive comprising three basic components, a fvariable displacement hydraulic unit and a differential gear unitsystem is supplied by a CSDU charge pump with its own reservo



Chapte

FIGUHydroConstDrive

ps/motors similar to that describeddependent of each other, connected

one to the other.

shaft with a gear at each end, one displacement hydraulic unit drive.peed.

Electrics-AC


RE 8-13-Mechanical ant Speed

96. The variable and fixed hydraulic units are swash plate pumin the Hydraulics section of the course. They are mechanically inonly by a stationary ported plate that permits a flow of oil from

97. The differential gear unit consists of a through carrier driven by the engine and the other meshing with the variableThus the variable unit rotates at a speed proportional to engine s



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re in mesh with each other and also unit. This permits the two sectionsar. One is in mesh with the fixedput drive to the generator.

ngle swash plate controlled throughpending upon the swash plate angleixed unit.

the required output speed for theriable displacement unit swash platerom, the fixed displacement unit.t drive from the engine, through thehe generator drive.

an initial tendency for the generatore variable displacement unit swashnt unit. The fixed unit is forced toion of the differential unit rotates in rotation of the planet gears and thepeed relative to engine input speed,ondition.

Electrics-AC


98. Carried on the carrier shaft are two planet gears which amesh internally with the two separate sections of the differentialto rotate at different speeds. Each section has an external gedisplacement hydraulic unit drive, the other meshes with the out

99. The variable displacement hydraulic unit has a variable aa control cylinder by a governor sensitive to generator speed. Dethe variable unit will either pump oil to, or accept oil from, the f

100. When the input speed from the engine is the same as generator, the governor, via the control cylinder, positions the vasuch that it is neither pumping oil to, nor accepting oil fConsequently the fixed unit remains stationary and there is direcdifferential unit carrier shaft and the variable hydraulic unit, to t

101. When the input speed from the engine increases there is to overspeed. This is sensed by the governor, which adjusts thplate such that the unit will accept oil from the fixed displacemerotate in the same direction as the variable unit, so the input sectthe same direction as the carrier shaft. This reduces the speed ofoutput section of the differential unit, reducing generator drive sto maintain a constant speed. This is known as the underdrive c



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nerator to underspeed causes thee variable hydraulic unit pumps oilte in the opposite direction to theates in the opposite direction to theet gears and output section of theed, to maintain a constant required

ht deck has provision for indication. The oil temperature on the outlet (sometimes both) are indicated oned by the illumination of an amberde an indirect indication of a faultyDU rpm. The significance of the of anticipating or trouble shooting

speed will result in unacceptableal failure of the CSDU or generatorearbox and necessitate engine shute, a facility is provided to disconnectnnect mechanism, which is remotely

Electrics-AC


102. When engine speed decreases, the tendency of the gegovernor to alter the variable unit swash plate angle such that thto the fixed hydraulic unit. This causes the fixed unit to rotavariable unit and so the input section of the differential unit rotcarrier shaft. This increases the speed of rotation of the plandifferential unit, increasing generator speed relative to engine spespeed. This is known as the overdrive condition.

103. To provide monitoring of the health of the CSDU the fligof oil temperature, oil pressure as well as generator indicationsside of the CSDU or the oil temperature rise across the CSDUgauges on the flight deck. Low oil pressure is normally indicatlight. Reference to the generator frequency meter can also proviCSDU. Some installations provide for indication of the CSmonitoring instruments is to provide the operator with a meansCSDU problems.

Remote CSDU Disconnect104. Failure of the CSDS to maintain constant generator fluctuations or variations of the AC supply frequency. Mechaniccould well cause expensive damage to the engine accessories gdown. To prevent the damage that would be caused in either casthe drive to the CSDU. Figure 8-13 shows the mechanical discooperated from the flight deck.



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FIGUCSDUDiscon

gagement with the threaded CSDUisconnect’ switch on the flight deckement with the threaded shaft. Thee dog clutch un-mesh, disconnectingpon the CSDU ceases to rotate.

Electrics-AC


RE 8-14 Mechanical nect

105. In normal operation the threaded pawl is held out of eninput drive shaft by a solenoid-operated pin. Operation of the ‘dwithdraws the pin and the pawl spring forces the pawl into engagrotating drive shaft therefore screws forward until the teeth of ththe CSDU shaft from the accessories gearbox drive shaft, whereu



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unning above a specified minimumcess to the engine is necessary. Byd shaft and a reset spring re-meshes out of engagement with the CSDUrated.

ck a captioned warning indicationequipped aircraft, a caption appearspower meter will read zero and theion.

cy System (VSCF)o as VSCF systems, are replacing AC systems. Using state-of-the-art

and allows for more flexibility ofess AC generator which is drivenut will vary with engine speed. The

the VSCF converter, where it is thenrter circuitry where it is formed intothree-phase 400Hz AC. Because the the aircraft and since there is noengine nacelle design. A schematicand frequency output of a VSCF is

Electrics-AC


106. Disconnect action can be made provided the engine is rrpm. The dog clutch can only be reset on the ground, since acpulling the reset handle the pawl is disengaged from the threadethe dog clutch teeth. The solenoid pin will now hold the pawlshaft until such time as the remote disconnect switch is again ope

107. When disconnect selection is made from the flight detypically displays on the electrical control panel and, on EICAS on the primary display. Also, the associated generator load or generator circuit breaker indicator will display an OPEN indicat

Variable Speed Constant Frequen108. Variable speed constant frequency systems, referred thydromechanical constant-speed drives in modern aircraft designelectronic components, the VSCF systems improve reliability generator installation. The system employs a standard brushldirectly from the engine gearbox and therefore its frequency outpvariable three phase output is fed to the full-wave rectifier withinchanged into DC and filtered. The direct current is fed to the invesquare wave outputs that are separated and summed to produce VSCF control system can be mounted virtually anywhere onrequirement for a CSD, it allows for a much more compact integrated VSCF system is shown at Figure 8-15. The voltage identical to the output of a CSDU or IDG.



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FIGUVariabConstFrequeSystem

) (LRU) that combines the brushless, which is mounted on the enginetachment ring. The integrated driven and cooling system and typicallys a reduction in generator size for antinuous operation at 90 kVA.

m for the generator. Oil pressure,ivates warning annunciators on theill automatically activate the CSDU

Electrics-AC


RE 8-15le Speed ant ncy s

Integrated Drive Generator (IDG109. The integrated drive generator is a line replaceable unitgenerator and its constant speed drive unit in a single casingaccessories gearbox by means of a quick attach/detach (QAD) atassembly is completely self-contained, having its own lubricatiorotates at a constant 12,000 rpm. The high rotary speed enablegiven power output. Modern IDG units are typically rated for co

110. The lubricating oil for the IDG is also the cooling mediutemperature and (in some cases) quantity is monitored and actflight deck. In some systems low pressure or high temperature wdisconnect mechanism.



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use in AC powered aircraft. Twin-ngine-driven generators each supplythe aircraft is approximately equally machines, producing 400 Hz, 115/by the APU also produces 115/200

craft’s power requirements and any most non-essential services. Theus bars can be cross connected, or

ctrical distribution system in which The advantage of such a system ofentarily, in the event of a generator

Electrics-AC


AC Power Distribution111. There are two types of electrical distribution system in engined aircraft use the Split Bus Bar System, in which the two etheir own AC bus bar, between which the total electrical load of divided. The generators are constant speed, frequency-controlled200 volts single-/three- phase. A third generator drive directly Volt 3 400Hz supply.

112. Any two generators are capable of supplying all the airsingle generator is capable of maintaining vital, essential, andgenerators cannot be cross connected, or paralleled; generator btied, in the event of generator failure.

113. Three- and four-engined AC powered aircraft use an elethe engine-driven generators are normally operated in parallel. distribution is that power supply is not interrupted, even momfailure.

φ



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quency AC as the primary powers (TRU’s). The generators providebars, from which non-essential ACither from one of the generator busry bus bar via the Essential DC busing connected by the bus tie breaker no load sharing circuits. A typical

Electrics-AC


Split Bus Bar Distribution114. Many modern airliners use non-paralleled constant fresource, with DC services provided via transformer-rectifier unit115v single-phase (200v three-phase), 400 Hz AC to their bus consumers are supplied. The essential AC bus can be supplied ebars, or from a static inverter supplied with DC from the battebar. The generator bus bars are isolated from each other, only be(BTB) in the event of failure of one of the generators. There aresplit bus bar system is shown at Figure 8-15.



Chapte

FIGUSplit BSystem

Electrics-AC


RE 8-16us Bar



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d in the Boeing 737-400 series. Thisother aircraft systems will differ

lar. Please refer to Figure 8-16 when

Electrics-AC


115. A practical example of the split bus bar system is that useis illustrated at Figure 8-17. It must be appreciated that considerably in detail, although the general concept remains simireading the following system description.



Chapte

FIGUBoeingSplit BSystem

Electrics-AC


RE 8-17 737 - 400 us Bar



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one of three sources. These are thetor and (on the ground) an externalures correct sequencing (to avoidity over an existing power source,

and DC systems. The AC standby and the DC standby bus from theth standby busses are supplied from

via a static inverter. A fully chargedht instruments, communication and

power sources to alternate poweron. If DC Bus 1 or Transfer Bus 1tery Bus. Automatic power transferrough the Standby busses when theutomatic transfer. This can be by-sition.

:

wer is connected and the ground 1 and BTB 2 will be open and theAC bus to the ground servicing bus.abin lighting (for the cleaners), the lighting circuits (for the engineers)

Electrics-AC


116. AC and DC power is normally supplied to the system fromengine-driven generators, an auxiliary power unit (APU) generapower supply. An interlocking system between breakers ensparalleling) and a source of power selected always takes priorautomatically disconnecting the latter.

Standby power. The standby busses supply the essential ACbus is normally supplied from the number one AC transfer busnumber one DC bus. In the event of failure of all AC supplies, bothe aircraft battery via the battery bus, and the AC standby bus battery has sufficient capacity to provide power to essential flignavigation equipment for a minimum of 30 minutes.

117. In flight, automatic switching is provided from normalsources when the Standby Power Switch is in the AUTO positiloses power, both Standby busses automatically switch to the Batis an in-flight feature only. To prevent the battery discharging thaircraft is on the ground, the air/ground safety sensor inhibits apassed by placing the Standby Power Switch in the BATTERY po

118. Consider a typical sequence of events in normal operation

External power (ground servicing only). If external popower control switch (on the flight deck) is selected OFF, BTBground-servicing relay will be energised to connect the external The ground servicing bus will normally supply power to the cfreight holds (for the loaders), the equipment/undercarriage bayand the battery charger.



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r control switch selected ON, bothstem. DC power is provided by the

generator bus to the GS bus.

nnected to both generator bus bars,from the flight deck control panel.eld open. BTBs 1 and 2 can only been switched off, otherwise they will

generator (say number one) isd closes GCB 1. The number one number two system is still suppliedwo generator bus bar.

itching its control switch ON opensly the two halves of the system and

ils, its GCB will be tripped open byliary contacts in the GCB closes toRMAL to ALTERNATE.

Electrics-AC


External power (aircraft systems). With the ground poweBTBs close to connect 3-phase AC external power to the whole syTRU’s and the GS Relay is energised to connect the number one

APU generator. On the ground the APU generator can be coto power the whole system, by selecting the APU relays closed Under these circumstances, GCBs 1 and 2 will be automatically hclosed once external power to the external AC receptacle has bealso be automatically held open.

Engine driven generators. When the first engine drivenavailable, switching its control switch ON trips BTB 1 open angenerator, via its bus bar, now supplies number one system. Theby external power or APU generator via BTB 2 and the number t

119. When the second engine driven generator is available, swBTB 2 and closes GCB 2. Their respective generators now suppground power can be switched OFF and disconnected.

Loss of an engine-driven generator. If either generator fathat generator's under-voltage protection system. A set of auxienergise the associated transfer relay, causing it to move from NO



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n, disconnecting the generator from position, connecting generator bus All services are maintained except generator bus via Main Bus 2. If

o system switch is set to ON, APU 2us 2), and the number two transferumber two generator.

e generator bus bars, since the APUtransfer relays.

ttery is the only source of power,ital DC), the Switched Hot Battery

generator is connected through ahich a proportion of the electrical

a bus tie breaker (BTB) to a singlencept is illustrated at Figure 8-17.

Electrics-AC


120. Imagine that the generator 2 has failed. GCB 2 trips opeits bus bar. The number two transfer relay moves to the ALTnumber one, via transfer relay two, to transfer bus number two.the non-essential AC consumers supplied from the number twothese are required, the APU can be started. When its number twwill close (connecting the APU generator output to generator brelay will return to NORMAL. The APU has now replaced the n

121. During flight the APU can only be connected to one of thcircuit breakers do not contain auxiliary contacts to activate the

122. In the event of failure of all generators the aircraft basupplying the Battery Bus (Essential DC), the Hot Battery Bus (VBus and the Standby busses.

Parallel AC Electrical Systems123. In paralleled AC electrical distribution systems each generator circuit breaker (GCB) to its own AC bus bar, from wservices is drawn. Each generator bus bar is connected throughbus bar, known as the tie bus or synchronising bus. The basic co



Chapte

FIGUParalleDistribSystem

s operating and all GCBs and BTBsl the generators are producing AC at each other, they will equally shareill open, disconnecting it from the

n the four bus bars, as shown at

Electrics-AC


RE 8-18led AC ution

124. It will be seen from Figure 8-18 that, with all generatorclosed, the generators are connected in parallel. Provided that althe same frequency and voltage, and that they are in phase withthe total electrical load. If one generator should fail its GCB wsystem, and the remaining generators will share the loads oFigure 8-18.



Chapte

FIGUParalleDistribSystemFailed

permits either isolated or paralleled-19.

Electrics-AC


RE 8-19led AC ution . No3

Split Parallel System125. An alternative form of split bus bar system is one that operation of generators. An example of this is shown at Figure 8



Chapte

FIGUSplit PBar Sy

Electrics-AC


RE 8-20arallel Bus stem



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bus bar. Pairs of load bus bars maybar. On a four-engined aeroplanetors. A split parallel system offers ars, or parallel operation of pairs ofm breaker. It is used in several large

cted to its own single-phase bus barom which single phase supplies areple, are taken by connecting to allenerator, bus bar is usually shown.C is taken from a bus bar with an

by transformer rectifier unit (TRUs)DC bus bars, which in turn provideconnected to both TRU and batteryraft batteries.

mum, whilst another (connected ineen them. The generator giving thetance armature windings of the lowainst its drive shaft, which may well damage to the engine accessories

Electrics-AC


126. Each generator is individually connected to its own load be paralleled by means of a synchronising, or combining, bus closing the split system breaker (SSB) will parallel all four generanumber of alternatives; split bus bar operation of the generatogenerators, or fully-paralleled operation by closing the split systemulti-engine aircraft, of which the Boeing 747 is an example.

127. In 3-phase AC distribution each generator phase is conneso there are, in reality, three load bus bars for each generator frtaken. 3-phase supplies, for larger induction motors for examthree bus bars. For the sake of simplicity only one load, or gNon-essential AC is taken from the load bus bar. Essential Aautomatic alternate source of power, the synchronising bus bar.

128. With the generators operating, DC supplies are provided supplied with AC from the appropriate bus bars and supplying non-essential DC. Essential DC is provided from DC bus bars supplies. Vital DC is from bus bars connected directly to the airc

Parallel Operation of AC Generators129. If the voltage output of one AC generator is at its maxiparallel with it) is at its minimum a heavy current will flow betwhigher voltage will be practically short-circuited by the low resisoutput generator. The low output generator will be motored agshear since it is usually designed as a weak point to preventgearbox.



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at is to say, their speed of rotationd generators is at a lower outputus speed of the higher frequency (itIn other words, the high frequency‘driven’ generator is doing nothing

stem real and reactive load is sharedhat all paralleled AC generators areby generator speed control, throughn current.

rotational speeds must be identical.lleled generators must lock together

establish system frequency and thisd, tending to motor the remaining

Electrics-AC


130. Similarly, AC generators are synchronous machines. Thand AC frequency is interdependent. If one of two parallelefrequency than the other, it will try to speed up to the synchronocan't of course, because it is coupled to the engine gearbox). machine attempts to drive the lower frequency one. Thus, the and the ‘driving’ generator takes the entire system load.

131. In order to prevent these imbalances and to ensure that syequally between paralleled generators, it is necessary to ensure tat the same output frequency and voltage. The first is achieved the CSDU, and the second by control of generator field excitatio

Real Load Sharing132. For paralleled generators to equally share real load their It is generator speed that determines output frequency, and parawith respect to frequency.

133. The higher frequency of a higher speeding generator willgenerator will take more than its share of the total system loaparalleled generators.



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phase C of the output from eachto form a load-sharing loop (seetroller, comprising an error detector

r. The function of the error detectorrnor in proportion to the magnitude amplifier boosts the strength of the

FIGUReal LGroup

Electrics-AC


134. Current transformers are used to sense the real load atgenerator. These transformers are then connected in series Figure 8-21). The output of each transformer is fed to a load conin parallel with the current transformer, and a magnetic amplifieis to produce an electrical output signal to the CSDU speed goveand direction of current flow through the detector. The magneticsignal to improve control accuracy

RE 8-21oad Sharing



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ces a proportional voltage in eachries a current will flow in the load-d by each transformer.

each transformer will be the same

a higher speed than the remainingwill induce an output voltage in itsrrent transformers, producing an

produces current flows in the load-causes the number 1 generator errorhilst the other three error detectorsrs.

its output frequency and the othereach of the four output frequenciescurrent (load sharing). The inducedrmer. The current flow in the load-

rmer and so there will be no currentd a ‘steady-state’ (constant speed)

Electrics-AC


135. The current in phase C of each generator output inducurrent transformer. Since the transformers are connected in sesharing loop and this will be the average of the currents produce

136. When all the generator loads are equal, the current fromand no current will flow through the error detectors.

137. At Figure 8-20, generator 1 is assumed to be running atgenerators and is consequently taking an increased load. This current transformer higher than that in the other three cuunbalanced situation in the load-sharing loop.

138. The higher voltage in the number 1 current transformer sharing loop and error detectors as shown at Figure 8-20. This detector to send a decrease speed signal to its CSDU governor, wwill send increase speed signals to their respective CSDU governo

139. The result is that generator 1 will slow down, reducingthree will speed up, increasing their output frequencies. When match, all four generators will be carrying exactly the same load voltage, and current flow, will be the same in each current transfosharing loop will be the same as the current flow in each transfoflow through the error detectors, which will consequently sensignal to their respective CSDU governors.



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heir output voltages must be equal,tion current. Suppose the excitatione its voltage regulator is set slightlyent of current flowing in oppositionload is increased whilst the loads ofe load sharing

FIGUReactiSharin

Electrics-AC


Reactive Load Sharing140. For paralleled generators to equally share reactive load tand these are dependent upon voltage regulators and field excitacurrent is higher in one generator than in the remainder, becausabove mean system value. This will produce a reactive componto the reactive loads of the other generators. Consequently, its the other generators are reduced, resulting in unbalanced reactiv

RE 8-22ve Load g Loop



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hich measure the current flowing inond load sharing loop. The outputransforming device, called a mutual0°.

sformers detect this and voltagestual reactors. These voltages will

ted generator current a reactive load reactive load. If the reactor voltage less than its share of total reactive

ore than its share of reactive load.rrent flow in the load-sharing loop

gure 8-21.

ds current and its associated errorreduce excitation current. In the

ow is in the opposite direction, such detectors consequently signal their

urrent transformer induced voltagesring loop and no adjusting signals to

Electrics-AC


141. Referring now to Figure 8-21, the current transformers wphase C of each generator are connected in series to form a secfrom each transformer is fed to an error detector by way of a treactor, which produces a phase displacement of approximately 9

142. If a reactive load imbalance occurs, the current tranproportional to the differential currents are induced in the mueither lead or lag the respective generator currents by 90°.

143. When the voltage in a particular reactor leads the associaexists, indicating that it is taking more than its share of the totallags generator current this indicates that the generator is takingload.

144. At Figure 8-21, generator 1 is over-excited and taking mThe excess voltage induced in its current transformer causes a cuand through the mutual reactors as indicated by the arrows at Fi

145. The result is that number 1 mutual reactor voltage leadetector signals the number 1 generator voltage regulator to remaining three mutual reactors, the load-sharing loop current flthat mutual reactor voltage lags current. The associated errorrespective voltage regulators to increase excitation current.

146. When all generators are carrying the same load, all four cwill be identical and there will be no current flow in the load-shathe voltage regulators - a steady-state situation.



Chapte

with engine driven generators. Thisn the APU gas turbine speed, which1011 Tristar is a notable exception.

ircraft, especially those required torbine, can be extended into the airpm. A governor system maintainsis a 3-phase AC machine, similar toormal system values of voltage and

rovided with controlling circuits toe) and frequency (400 Hz). In theon of over- or under-voltage couldnd generators. Similarly, a fault in aency, which may cause overheatingard against the potential hazards of AC distribution systems.

Electrics-AC


Auxiliary Generators147. APU driven generators can rarely be operated in parallel is because their speed (and therefore frequency) is dependent upois not controlled on the same basis as a CSDU. The Lockheed L-

148. Ram air turbine driven generators are fitted in many ameet extra safety standards such as EROPS. A fan, or air tustream to drive a generator through gearing at about 12,000 rconstant rpm by adjustment of fan blade angle. The generator the main engine-driven generators and its output is regulated to nfrequency.

Generator Control and Protection149. As has been demonstrated, AC generation systems are pmaintain constant and correct output voltage (115v single phasevent of a fault developing in the voltage regulators a conditioccur, with potentially damaging results to system components aCSDU, or its control system, could lead to under- or over-frequdue to excess current flow and faulty operation of motors. To gusuch faults, various protection systems are usually provided with



Chapte

eration system is usually the resultor earthing of the field windings orant frequency AC systems it is usualolid-state circuit which will trip the

its bus bar. The over-voltage relay

a generator is being shut down andg of the voltage sensing circuit. Thection in that an incorrect detected

ection devices this also causes an

ge less than 100 ± 3 volts. A timelay is not tripped due to transient under-frequency condition during

r-voltage protection units operate inenerators.

ad sharing circuit of a paralleled ACems have sensing, warning and (inDU's or generators.

rt-circuit of a feeder line or bus bars current transformers measures thestream side of the bus bar. A short

ich causes the differential protection

Electrics-AC


Over-voltage Protection. Over-voltage in any electrical genof a fault in the field excitation circuit, such as short-circuiting open-circuiting of the voltage sensing lines. In three phase constfor the voltage in all three phases to be detected and fed to a sgenerator circuit breaker (GCB), disconnecting the generator fromis set to operate at a voltage greater than 130 ± 3 volts.

Under-voltage Protection. Under-voltage occurs whenevercould also be due to a field excitation circuit fault such as earthinprotection circuit is similar to that used for over-voltage protevoltage trips a relay which opens the GCB. In both protannunciation of the condition on the flight deck.

150. The under-voltage protection circuit operates at a voltadelay of 7 ± 2 seconds is included to ensure that the GCB revoltages. In addition, it allows the CSDU to slow down to anengine shut down, thereby inhibiting tripping of the GCB. Undeconjunction with the reactive load sharing circuits of paralleled g

Under- and Over-Frequency Protection is provided by the real logenerating system. Non-paralleled constant frequency AC systsome cases) automatic disconnect systems associated with the CS

Differential Current Protection. This system protects against shoand the very high current flow which would result. In AC systemcurrent flow on the return side of the generator and on the downcircuit would cause a significant difference between the two, whdetector to trip a relay, opening the generator circuit breaker.



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onnected to the same synchronisingquency and for the AC wave formsing lights is provided to indicate theconnected to the synchronising busd C of each generator between the

to be paralleled and that Gen. 1 is bar and Gen. 2 load bus bar via itsing and both generators have beenynchronising light selector switch isnising bus bar via the synchronisingts are sensitive to phase differencee longer the time interval betweening’ Gen. 2 is now finely adjusted tois closed while both lights are out,

scribed in DC Components, but the to DC, since it is essentially a DC

Semiconductor devices.

A shunt is essentially a low valueeter.

Electrics-AC


Monitoring EquipmentSynchronising lights. Before two or more generators are cbus bar it is necessary for them to be at the same voltage and freto be in phase with each other. In some aircraft a system of flashphase comparison between generator outputs, before they are bar. These synchronising lights are connected into phases A angenerator output and the synchronising bus bar.

151. With reference to Figure 8-19, assume Gen. 1 and 2 arerunning and connected to its load bus bar, the synchronising busGCB and the two BTBs. Gen. 2 GCB is open. Gen. 2 is runnadjusted as closely as possible to the "master" frequency. If the smoved to the Gen. 2 position, Gen. 2 is connected to the synchrolights whilst Gen. 1 is connected to it directly. Hence the lighbetween the two generators. The less the phase difference thflashes of the synchronising lights. The frequency of the ‘oncomobtain the greatest time interval between flashes and its GCB indicating that the two generators are in phase with each other.

AC voltmeters are usually of the moving coil type previously deinstrument additionally contains a bridge rectifier to convert ACdevice. Bridge rectifiers are described in the section dealing with

AC ammeters contain a bridge rectifier, transformer and shunt.resistance connected in parallel with the moving coil of the amm



Chapte

ey too are moving coil instruments.supplied to both fixed and movingnes the extent by which the moving

n some cases, real power (kW) and may be combined in a Watt/VAR

the frequency meter. A current coil of the moving coil instrument,

g will therefore be calibrated in kW.across phases A and C. This causesoil current. If the power factor is 1in the instrument, and the reading iser the interaction between the twoe instrument.

e that they will continue to operatesually incorporate a ‘press-to-test’

a circuit through the bulb filament.

ditions (generator overheat, engine

l conditions (power not available to

Electrics-AC


Frequency meters are a standard item of AC instrumentation. ThA potential, dependent upon supply voltage and frequency, is coils. Interaction between the magnetic fields produced determicoil is deflected against the control springs.

Power meters indicate total power being generated (kVA) and, ireactive power (kVAR). Both real and reactive power displaysmeter.

152. Functioning of the power meter is similar to that oftransformer senses generator phase B load and energises the fieldproducing a field proportional to load. The instrument's readinWhen the VAR reading is selected, the moving coil is connected the current in the moving coil to be 90° out-of-phase with field cthere is no interaction between the two magnetic fields produced zero (no reactive load). The lower the power factor, the greatfields and the higher the reactive (kVAR) reading displayed by th

Failure warning lights are supplied from a DC bus bar to ensurwhen the AC service they are indicating has failed. They ufunction, activated by pressing in the bulb holder to complete Indicating lights fall into three categories:

(a) Warning, coloured red and indicating unsafe confire warning, low CSDU oil pressure).

(b) Caution, coloured amber and indicating abnormaa bus bar).



Chapte

g safe/operable (power available).tion only.

ge and frequency to AC at anotheric induction. A basic transformerre 8-23.

FIGUBasic T

uit for an alternating magnetic flux a primary winding connected to a a secondary winding, connected totransference of electrical energy by

power transformers, and currentransformers are discussed in greater

Electrics-AC


(c) Advisory, coloured green or blue and indicatinWhite indicating lights or captions are for informa

Transformers153. Transformers are devices for converting AC at one voltavoltage, but the same frequency, by means of electro-magnetconsists of three major components, which are illustrated at Figu

RE 8-23ransformer

154. A laminated soft iron core provides a low reluctance circfield. This field is created by alternating current flow throughpower source. The alternating magnetic field induces an emf inthe output circuit. Thus, the transformer is a device for the mutual induction.

155. There are two classes of transformer; voltage or transformers. Various types of power transformers, and current tdetail later in this chapter.



Chapte

s. When an AC voltage is applied toalmost equal to the supply voltage.ion current sufficient to set up an secondary winding and, by mutual

e secondary winding the secondaryndary winding. This current flowreducing it. This in turn reduces theary current to flow and maintaining

upply current also increases. Eddyer loss, but this is minimised byhin insulating coat of varnish.

hrough both primary and secondaryemf induced by electro-magnetism Thus, if the secondary winding hascondary terminals will be twice thatut to input ratio, known as the

Electrics-AC


Transformer Function156. The principle of operation of the transformer is as followthe primary winding self-induction establishes a voltage that is The difference between the two voltages produces an excitatalternating flux in the transformer core. This flux cuts across theinduction induces a voltage in the secondary winding.

157. When a load is connected to the output terminals of thvoltage causes current to flow through the load and the secoproduces a magnetic flux that tends to oppose the primary flux, self-induced voltage in the primary winding, allowing more primthe core flux at an approximately constant value.

158. Thus, as secondary load current increases, the primary scurrents induced in the iron core lead to heating and powconstructing the core from soft iron laminations separated by a t

159. The alternating field in the core of the transformer cuts twindings. It will be remembered that the magnitude of the depends upon the number of turns of conductor ‘cut’ by the flux.twice as many turns as the primary winding, the voltage at the seof the voltage applied to the primary winding. The outptransformation ratio, is expressed in the equation:

EsEp-------

NsNp-------=



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former is said to be a ‘step-up’ type,ll power loss due to eddy currents,

r out. Thus, if the secondary outputormer has a 5:1 step down functionst be 10a.

ortly), step up transformers give an therefore lower current, to remote equipment at the requisite voltage. example.

t amperes. Because the resistance ofnductive unit. It has already beenquency is reduced the current flowarer to magnetic saturation, which

Electrics-AC


Where:

(a) Es = secondary voltage

(b) Ep = primary voltage

(c) Ns = number of turns in secondary winding

(d) Np = number of turns in primary winding

160. When secondary voltage is greater than primary the transthe reverse function is known as a ‘step-down’ type. If the smaplus other minor losses, is ignored the power in is equal to poweis 100v, 50a (5000w) the input will also be 5000w. If this transfthe primary voltage must be 500v and so the primary current mu

161. As well as step down applications in TRUs (discussed shinvaluable cable weight saving by supplying higher voltage, andequipment. A local step-down transformer can then supply theStep-down transformers are used to supply fluorescent lights, for

162. Transformers are usually rated in volt-amperes or kilovolthe primary winding is so low it may be considered a purely ishown that in such a circuit if voltage remains constant and frewill rise. The increased current brings the transformer core nereduces inductance and increases current flow further.



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at Figure 8-24, together with itsat it is a cored transformer; in someting the same thing.

Electrics-AC


Types of Transformer163. A drawing of a standard power transformer is shownelectrical symbol. The two parallel lines in the symbol indicate thpublications the symbol is shown with three parallel lines, indica



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FIGUPowerTransfo

Electrics-AC


RE 8-24 rmer



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a simpler type of transformer calledis the same as with a conventional combined. Auto transformers areers, however their use in aircraft

primary/secondary isolation makesangement for an auto-transformer is

Electrics-AC


164. Where only a small voltage step, up or down, is required an auto-transformer can be used. The principle of operation transformer, but now the primary and secondary windings areconsequently smaller and lighter than dual winding transformsystems is usually restricted to lighting supplies, since the lack ofthem suitable for use in low current systems only. The circuit arrshown at Figure 8-25, together with its electrical symbol.



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FIGUAuto T

Electrics-AC


RE 8-25ransformer



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e as voltage transformers, but use an order to give a secondary outputt-carrying primary conductor. The

FIGUCurrenTransfo

power meters, to monitor currentlation and protection systems.

Electrics-AC


165. Current transformers operate on the same basic principlcircuit's normal insulated conductor as the primary winding, ivoltage proportional to the current that is flowing in the currenprinciple of the current transformer is illustrated at Figure 8-26.

RE 8-26t rmer

166. Current transformers are used to supply ammeters andflows for fault detection systems and in many AC generator regu



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s)craft's regulated three phase powerbined TRUs, to supply the 28v DCree phase transformer consists of ad delta form. The reduced voltageode bridge rectifiers, which converthown at Figure 8-27.

Electrics-AC


Transformer Rectifier Units (TRU167. These are used in the conversion of AC to DC. An airsupply is stepped down by transformers and then rectified in combus bars for battery charging and other DC loads. A typical thstar wound primary winding and secondary windings in star anthree phase output from the secondary windings is fed to six dithe AC to DC. The circuit diagram for such an arrangement is s



Chapte

FIGUTransfoRectifi(T.R.U

ven in the section dealing withreciate that a diode will only permitymbol.

Electrics-AC


RE 8-27rmer -

er Unit .)

168. A description of bridge rectifiers and diodes is gisemiconductor devices, but for the moment it is sufficient to appcurrent flow in the direction indicate by the arrow in the diode s



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s that convert DC to AC. On oldery electrical power source, constant employed to supply the AC power

as the principal source of electricalsource. 28v DC is supplied to am, and then reshapes this into the at Figure 8-28.

FIGUSquareForm

parts, and therefore being far morend therefore the need for a cooling

‘arcing’ at low air pressures, as theirnecessity to locate them within the

Electrics-AC


Inverters169. Whereas rectifiers convert AC to DC, inverters are deviceaircraft, employing DC engine-driven generators as the primarspeed, DC motor driven, AC generators (rotary inverters) wereneeded by avionic equipment.

170. Later generations of smaller aircraft, whilst still using DCpower, employ static (solid state) inverters as an AC power transistorised circuit that produces 400 Hz in square pulse forrequisite sine wave form. A square wave pulse form is illustrated

RE 8-28 Wave Pulse

171. Static inverters have the advantage of having no movingreliable. With static inverters there is very little heat generated aair supply is removed. Finally, static inverters do not suffer from rotary forebears did, and consequently there is no overriding pressure hull.



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rs) as the primary electrical powers is provided by a battery fed static are not available.

able speed control, then DC motors aircraft applications in which theseumstances AC motors prove to bes of AC motor are brushless andand expensive AC/DC conversion

as a DC motor in construction andwith high frequency supplies and is00Hz AC supplies.

that is synchronous with the applied and the stator a soft iron shell with magnet rotor follows the rotatingf current to the stator coils.

Electrics-AC


172. Most large modern aircraft use AC generators (alternatosupply. Emergency AC power supply to essential AC consumerinverter when the normal 115v AC power sources have failed or

AC Motors173. Where there is a requirement for high torque and/or variare generally superior to AC motors. There are, however, manyrequirements are either absent or non-rigorous. In these circsuitable, or even preferable, since the most widely used typetherefore require less maintenance. Also the need for heavy equipment is reduced.

Types of AC MotorThe universal AC/DC commutator motor is essentially the sameis found in many household appliances. It is unsuitable for use therefore not normally found in aircraft, which usually employ 4

The synchronous motor is so called because it rotates at a speed AC frequency. In small motors the rotor is a permanent magnetcoils wound around it and supplied with AC. The permanentmagnetic field set up in the stator by the sequential application o



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net excited by DC from an externaltype of motor. None of them areconstant speed under varying loadrloaded it will stall. The principalacteristic and the fact that they tenday as a capacitor as far as the AC

ary speed indicators (tachometers). its AC output is connected to aof the alternator output is directlytor is synchronous with its supplymeter through a permanent magnet in operation, except that a turbine,ator.

rotate at a speed synchronous with versions; the two and three-phaseose speed as load increases and viceontrol, if necessary. This type of AC motor.

s basically simple. By sequentiallying magnetic field is produced. This magnetic field resulting from thistraction.

Electrics-AC


174. In larger synchronous motors the rotor is an electromagsource. There are one, two and three-phase versions of this inherently self-starting although, once running, they maintain conditions. Unfortunately, should a synchronous motor be oveadvantages of synchronous motors are their constant speed charto correct the power factor because they behave in the same wsupply is concerned.

175. Synchronous motors are commonly used as remote rotThe aircraft engine drives a small three-phase alternator andsynchronous motor inside the rpm indicator. The frequency proportional to engine speed and speed of rotation of the mofrequency. The motor drives the indicating needle of the tachoand spring-loaded drag cup arrangement. A flowmeter is similarthe speed of which is dependent upon fluid flow, drives the altern

The induction motor is an asynchronous motor since it does notthe applied AC frequency. There are one, two and three-phasemotors are inherently self-starting. Once running, they tend to lversa, but this tendency can be overcome with automatic speed cmotor is commonly found in aircraft and is normally a brushless

The Induction Motor176. The principle of operation of an AC induction motor ienergising the windings in the casing (stator) of the motor a rotatinduces a current flow in the conductors of the rotor and thecurrent flow ‘follows’ the rotating stator field due to magnetic at



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rent is carried in windings of copperype the rotor is drum-shaped and isnged axially and connected to flate of low resistance, so that the emf and, therefore, a strong magnetic

igure 8-29.

FIGUSquirreMotorConst

nd are typically only used where

Electrics-AC


177. The rotor may be wound, in other words the induced curwire, or it may be of ‘squirrel cage’ construction. In the latter tmade up of a ring of closely spaced parallel copper bars arraconducting rings at either end. In either case the conductors arinduced by the rotating stator field sets up a large current flowfield. The construction of a squirrel cage rotor is illustrated at F

RE 8-29l Cage

- ruction

178. Squirrel cage motors are very low-torque machines amechanical loads are light.



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hree-phase, squirrel cage inductionh a rotating magnetic field is set upave relationship of the three supply

FIGUSquirreMotorOpera

Electrics-AC


179. Figure 8-30 illustrates the principle of operation of a tmotor. The upper part of the diagram shows the means by whicin the stator of the machine. The lower part shows the sine wphases.

RE 8-30l Cage

- tion



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C are both carrying current, but ofetic fields set up by phases A and Crrow. If we now progress to the of opposite polarity, and phase A isC produce a resultant field at 60° towise 60°. By supplying three-phase the stator field will rotate through

uctors in the rotor, inducing an emfl be attracted to ‘follow’ the rotatinghe greater the current flowing in thefield they produce, consequently theherefore, the greater the magnetic

ction between the rotor conductorsf and torque, because their relativee reached where the relative rotor/

mechanical load on the motor.

ld, since if it did there would be noere would be no current flow in the the torque necessary for rotation.nous) speed is called slip speed, ando produce the necessary increases ofociation with these motors.

Electrics-AC


180. Consider then the situation at instant 0°. Phases A and opposite polarity, and phase B is carrying no current. The magnproduce a vertical resultant field, as indicated by the heavy asituation at instant 60°, phases B and C are carrying current, butcarrying no current. The magnetic fields set up by phases B and the vertical. In other words the resultant field has rotated clockAC to the stator windings they will be sequentially energised and360° at supply frequency.

181. This rotating stator field will cut through the axial condand current flow. The magnetic field due to this current flow wilstator field, producing torque and causing the rotor to rotate. Tstator windings the greater the strength of the rotating magnetic greater the induced emf and current flow in the rotor and, tattraction, and torque produced.

182. As the speed of rotation of the rotor increases, the interaand the stator field decreases causing a decrease of induced emmovement decreases. Eventually a steady-state condition will bstator field speed is producing just sufficient torque to match the

183. The rotor never rotates at the same speed as the stator fierelative movement and no induced emf. With no induced emf throtor and therefore no rotor electro-magnetic field to maintainThe difference between the rotor speed and stator field (synchroit increases as the torque load on the motor increases, in order trotor emf and current flow. Hence the term asynchronous in ass



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tor runs more slowly. Eventually, ifcally this is called the pullout point.

termined by the number of polesply. It is defined as:

ad is applied to an induction motor.ctance of the rotor. When the rotorator field, the frequency of the rotor

re is a corresponding increase in theactance of the rotor and the powerr is equal to the cosine of the phase this phase angle.

ively fine wiring of the field coils, iss the motor will continue to run at. The extent of the speed reductiontor runs at about half speed unless it altogether. In an aircraft the fault motor, when it will not re-start.

ween the stator windings, causesally greater with increased load.

(Hz) 60×s of poles------------------------

Electrics-AC


184. As the torque load increases, so slip increases and the mothe torque load becomes excessive, the motor will stop. TechniPullout torque is at least twice the normal full load torque.

185. Synchronous speed of an induction motor field is deproduced by the stator windings and the frequency of the AC sup

186. There is another effect that must be considered when a loThis is a lowering of the power factor caused by the inductive reais turning at almost synchronous speed, that is the speed of the stcurrent and the inductive reactance of the rotor are low.

187. As a load is applied to the rotor, the slip increases and thefrequency of the rotor current. This increases the inductive refactor of the motor consequently decreases since the power factoangle between voltage and current. Inductive reactance increases

188. Open circuiting of one phase, due to breakage of the relata fairly common fault in induction motors. When this happenreduced speed (and will emit a humming noise whilst doing so)will, theoretically, be to 60% of normal rpm. In practice the mois particularly heavily loaded, when it may slow down and stopmay well not be apparent until the next occasion of requiring the

189. Short-circuiting between phases, in other words betoverheating and loss of power and speed, the effect is proportion

Synchronous Speed (rpm)Frequency No. of pair--------------------------=



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ction of rotation of the motor. Fortation of the stator field, then theld and therefore of the rotor. From be reversed to reverse the directiontem.

rcraft use and is commonly found inos, gyro rotors, gyro torque motorsmbient conditions than DC motors.

r winding, split into two parts set ate energised and a rotating magneticen reached a centrifugally operated continues to run as a single-phaseor.

phase AC split-phase motor. Whendirect to the clockwise field coil andct of causing the current in the anti- a clockwise rotating field. Placingct. The rotor of such a motor is of

Electrics-AC


190. Reversing the supply phase sequence will reverse the direexample, if the phase sequence A, B, C produces clockwise rosequence B, A, C must produce anti-clockwise rotation of the fiethe foregoing it can be seen that only two of the three phases needof rotation and this is easily achieved with a simple switching sys

191. The induction motor is the most suitable AC motor for aisuch applications as fuel pumps, hydraulic pumps, autopilot servand AC actuators. AC motors are less vulnerable to changes in a

192. Single-phase induction motors usually have a single stato90° to each other. During start up, both parts of the winding arfield is set up in the stator. Once sufficient rotor speed has beswitch cuts out one part of the stator windings and the motormachine, albeit less powerful than an equivalent three-phase mot

193. Figure 8-30 shows the basic circuit for a reversible singlethe control switch is in the clockwise position current is supplied by way of a capacitor to the anti-clockwise coil. This has the effeclockwise coil to lead the current in the clockwise coil, creatingthe control switch in the anti-clockwise position reverses the effesquirrel cage construction.



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FIGUReversPhase Phaser

powered devices such as equipmentding/taxi light.

asychronous motors) where a servoat 90° to each other, but they aresupply which is a constant voltage.n be varied, thus allowing the speed

depends upon its atomic structure.an electric current will flow easilyeadily give up, or receive, electrons.as free electrons.

Electrics-AC


RE 8-31able Single AC Split Motor

194. Single-phase induction motors are used for relatively lowcooling fans, aircraft rotating navigation lights and retracting lan

195. Some AC systems employ two phase induction motors (control of synchronous devices is required. The windings are connected to different voltage sources. One source is the main The other source serves as a control voltage whose amplitude caof the motor to be varied.

Semiconductors196. The ability of a substance to conduct an electric currentSome elements, mostly metals, are good conductors because through them. This is because the atoms of these elements will rElectrons that freely move from one atom to another are known



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The nucleus of the atom comprisesarged electrons. The total electron

harge. Thus, the atom is electrically electrical charge and they play no

concentric shells. A simple atom,mplex atoms may have many shells

t Figure 8-31.

FIGUAtomiExamp

Electrics-AC


197. An atom is made up of electrons, protons and neutrons. positively charged protons surrounded by orbiting negatively chcharge is equal to, but of opposite polarity to, the total proton cneutral. Neutrons have the same mass as protons, but have nopart in electrical conduction.

198. The electrons orbit the nucleus of the atom in layers, orwith only one or two electrons, has a single shell, whilst more coof orbiting electrons. Examples of atomic structure are shown a

RE 8-32c Structure le



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orbit and the electrons in this orbite orbit which take part in electricalo move within the valence orbit, orbit, the less freedom of movement electrons in their valence orbits are

electrons in the valence orbit offernductors.

the valence orbit of any element'sinsulator (a non-conductor) or a

ent atoms are able to overlap eachell is subjected to an electrical forcecape and be attracted to the positiven an electron escapes from an atom,ive charge. This attracts any mobile

is a movement of negatively chargedctrical source) and a movement ofe charges are known as holes, intofrom positive to negative, assumesement of electrons in the opposite

Electrics-AC


199. The outermost shell of the atom is known as the valenceare known as valence electrons. It is the electrons in the valencconduction. To be able to do this, these electrons must be free tshell. The greater the number of electrons in the valence oravailable. Elements comprised of atoms with a high number ofpoor conductors of electricity, or insulators.

200. Elements whose atomic structure has a low number of greater freedom of electron movement and are good electrical co

201. The number of electrons which are actually present in atoms determines whether that element is a conductor, an semiconductor.

202. In a conductor, the outer shells (valence orbits) in adjacother. When a voltage is applied to the conductor the outer shwhich, if sufficiently strong, will cause some of the electrons to espole of the electrical supply (unlike poles attract). However, whethe atom is no longer electrically neutral, but now has a net positelectron in the vicinity.

203. Thus, when current is flowing through a conductor there electrons in one direction (towards the positive pole of the elepositive charges in the opposite direction. These mobile positivwhich mobile electrons can fall. ‘Conventional’ current flow, movement of holes in the direction of current flow with movdirection.



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ll of each atom is too great for theugh the material extremely difficult.

r shell of their atomic structure thature state such a material offers highectrons (electrons in the outer shell)ons is added, or removed, the outer’ electrons, capable of mobility, andalue. Such elements are called

itting current flow, in one directione direction. A rectifier is, in effect,

ectifiers employ semiconductors to

nductor material. Semiconductorce) shell. Germanium atoms have a. Both of these elements are highlyalence bond. In other words, the

se of adjacent atoms. Consequentlynt carriers. This is illustrated at

Electrics-AC


204. In an insulator the number of electrons in the outer sheouter shells to overlap, which makes movement of electrons thro

205. Some materials contain a number of electrons in the outeis midway between that of a conductor and an insulator. In its presistance to current flow because there are sufficient valence elto prevent ‘overlapping’. However, if a precise number of electrshells can be persuaded to overlap such that there are a few ‘freethe element's resistance can be reduced to a very low vsemiconductors.

Semiconductor Rectifiers206. A rectifier is a device that offers very low resistance, permand very high resistance, preventing current flow, in the oppositthe electrical equivalent of a non-return valve. Solid-state rachieve this effect.

207. Typically, germanium or silicon is used as the semicomaterials in their pure state have 4 electrons in their outer (valentotal of 32 electrons. Silicon atoms have a total of 14 electronsresistive because the atoms of these materials have a strong velectrons in the outer shell of each atom naturally pair with thothere are no free electrons in these materials to act as curreFigure 8-32.



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FIGUSemicoValenc

er germanium or silicon, a processble of conducting current. Suppose,. Antimony atoms have 5 electrons

becomes a free electron. Hence, the

N-type germanium, because it nowe material is still electrically neutral,l number of electrons.

es, or holes, left in the valence bondonly have 3 electrons in their outercan then fill these holes. This leavesresents a net positive charge, and

germanium but as with N-type, the

Electrics-AC


RE 8-33nductor

e Bond

208. If a small quantity of another element is added to eithcalled doping, the resulting impure semiconductor becomes capafor example, a small amount of antimony is added to germaniumin their outer shell, so the fifth electron cannot pair, or bond, andgermanium becomes conductive.

209. When germanium is doped with antimony it is known ashas some free electrons and electrons are negatively charged. Thsince the total number of protons in each atom balances the tota

210. If germanium is doped with indium there are vacant spacbetween germanium and indium atoms, because indium atoms shell. Electrons that break away from an adjacent valence bond a hole elsewhere, and so the process is repeated. A hole repconsequently germanium doped with indium is known as P-typematerial is still electrically neutral.



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nd an N-type material will give up

joined together the P-type materialmaterial as a cathode (an emitter ofce that permits the flow of electronsde (from the N-type to the P-typede, (from the P-type to the N-type

menon would occur. The positivelyually attracted and drift towards thection to fill holes.

and become negatively charged and charged. Note, however, that the, the total number of electrons still

he potential difference between theged P-type materials (0.3 volts innt of electrons across the junction.

of P-type and N-type germanium is

Electrics-AC


211. Remember that a P-type material will accept electrons aelectrons.

212. Were P-type and N-type semiconductor materials to be would act as an anode (a gatherer of electrons) and the N-type electrons). This device would in effect be a diode, which is a deviin one direction only. Electrons will flow from cathode to anomaterial), but not in the opposite direction from anode to cathomaterial).

213. At the junction of the two materials an interesting phenocharged holes and the negatively charged electrons would be mutjunction, and some of the electrons would migrate across the jun

214. Thus, the P-type material would gain (negative) electronsthe N-type material would lose electrons and become positivelymaterial would still be electrically neutral overall (that is to sayequals the total number of protons).

215. At a point between the P-type and the N-type material tnow positively charged N-type and the now negatively chargermanium) creates a potential barrier which stops the movemeFigure 8-33 illustrates the condition that exists when a junctionmade.



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FIGUP typeGermaJunctio

d state diode works, except that the in the middle. It is rather a single way that one end is P-type and the

the terminals of a battery, it is found in the other.

sitive battery terminal connected to charged) N side connected to the conventional current flow from at a net migration of electrons in thee. The diode has become a goodrward direction has been reduced to

Electrics-AC


RE 8-34 and N type nium n

216. The preceding paragraphs describe accurately how a solidiode is not in fact two pieces of semiconductor material joinedpiece of germanium (or silicon) which has been doped in such aother end N-type.

217. If the diode illustrated at Figure 8-33 is connected across that current will flow through the diode in one direction, but not

218. If the connection is made as at Figure 8-34, with the pothe (negatively charged) P side of the diode, and the (positivelynegative battery terminal, current will flow. (Remember that apositive terminal through a circuit to a negative terminal is in facopposite direction). This is known as a forward-biased diodconductor in the direction shown because its resistance in the foa low value.



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FIGUForwaDiode

, P side to negative and N side ton towards the positive connection

as illustrated at Figure 8-36. Withrent flow through the diode. This isn has been increased sufficiently to a diode is illustrated at Figure 8-36.

Electrics-AC


RE 8-35rd Bias

219. If the connections to the battery terminals are reversedpositive, the negatively charged electrons in the N side are drawand the positively charged holes towards the negative connectionno movement of electrons across the junction there can be no curknown as a reverse-biased diode. The resistance at the junctioprevent current flow through the diode. The symbol representing



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FIGUSymboDiode

FIGUReversDiode

Electrics-AC


RE 8-36l for a

RE 8-37e Bias



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tion of alternating current to directcycle of alternating current can passction of flow is reversed during theses. Consequently, the DC output is to the AC supply frequency. This isiagram at Figure 8-37.

Electrics-AC


Half-Wave and Full-Wave Rectification220. A single diode can be used to achieve half-wave rectificacurrent. Placed in series with a single phase AC supply, one half through the diode to the DC output. When supply current diresecond half cycle the diode is reverse-biased and current flow ceain a series of pulses, the number of pulses per second being equalknown as half-wave rectification and is illustrated in the upper d



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FIGUHalf anRectifi

Electrics-AC


RE 8-38d Full Wave cation



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es three diodes (one per phase) withrity DC terminal is connected to thes, the number of pulses per second

in the lower diagram at Figure 8-37,f AC to DC is achieved. In aircraft

, in which case the rectifier needs six.

Electrics-AC


221. Half-wave rectification of a three-phase AC supply requirtheir DC outputs connected to one terminal. The opposite polaAC neutral return. This arrangement also produces DC pulsebeing three times the frequency of the AC supply.

222. If four diodes are connected in a bridge circuit, as shown and supplied with single phase AC, then full-wave rectification oit is usually necessary to obtain DC from a three-phase alternatordiodes connected in such a way as to provide one-way paths only



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FIGU3 PhasWave

Electrics-AC


RE 8-39e AC Full Rectification



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AC to DC is shown at Figure 8-38.oint in the AC cycle, diodes D3, D4-biased. Consequently, current flowrom the DC load is through D4 ands will be forward-biased and threethe DC load.

follows :

Figure 8-39 will conduct electricityf the zener diode is designed aroundis a set voltage at which the zeners a normal diode. If the breakdown

oduces what is known as avalanchero. They are ideal for use in voltage

Electrics-AC


223. A circuit to provide full-wave rectification of three-phaseCurrent flow at instant 90° is indicated. At this instantaneous pand D5 are forward-biased and diodes D1, D2 and D6 are reversefrom all three phases is to the DC load through D3 and return fD5. At any instant in the three phase supply cycle three diodereverse-biased to ensure constant direction current flow through

224. Some typical diodes and there applications are listed as

(a) Zener Diode. The Zener Diode as illustrated at only under certain voltage conditions. Operation oits breakdown voltage. The breakdown voltage diode will conduct, below this set voltage it acts avoltage is reached in the reverse bias mode, it preffect. At that point the diode resistance falls to zeregulation.



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FIGUZener

ten connected in parallel with theon eliminates any transient voltages relay or solenoids electro-magnetic

ener will conduct current above a

Electrics-AC


RE 8-40 Diode

(b) Bi-directional Zener Diode. These diodes are ofmagnetic coil of a relay or solenoid. Their inclusithat could be created during the rise and fall of thefield. Like the zener diode, the bi-directional zcertain voltage level. (Figure 8-40).



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FIGUBi-direZener

tems as indicator lights to displaypplied voltage must be connected inwhen the diode is forward biased are available in LEDs which ishorus, gallium and arsenic used in0 to 20mA to produce the required

Electrics-AC


RE 8-41ctional Diode

(c) Light-emitting Diode (LED). Used in aircraft sysletters and numbers. For an LED to conduct, the athe forward biased condition. Energy given off produces light and heat. A variety of coloursdetermined by the active elements such as phosptheir construction. LEDs require 1.5 to 2.5V and 1light.



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FIGULight-eDiode

can provide indication of letter andare commonly grey but some colour materials that contain moleculeshrough the crystal, it is ‘bent’. If aes align and the light passes straight to align light waves with polarisedlight to form the pattern for display.

n at Figure 8-43. In this example,

Electrics-AC


RE 8-42mitting s

(d) Liquid Crystal Display (LCD). This type of displaynumber patterns or even form a full picture. LCD systems are available. Liquid crystals are fluidarranged in crystal forms. When light is passed tvoltage is applied to the liquid crystal the moleculthrough the material. LCDs use this phenomenonfilters. The polarised filters will pass or block the A typical 7-segment liquid crystal display is show



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a number 3. The segments that do waves pass through these segmentsolariser. The segments that form theing the light. The liquid crystals of

FIGUElectro1,2,3,4Energithe Nu

Electrics-AC


voltage is applied to the individual segment to formnot have voltage applied are light grey. The lightand reflect off a mirror mounted behind the rear pnumber 3 are dark grey, because they are reflectthese displays are aligned by an applied voltage.

RE 8-43des No and 6 are sed to Form mber 3



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tance internally to either permit orcircuit, and to increase the outputor as an amplifier. The name is a

s opposed to two in a diode) and is or silicon, two of which are N-typeistors. They may be either n-p-n or

FIGUJunctioTransis

of N-type germanium (or silicon). One end of the transistor is called section is called the base. A p-n-pnd the emitter and collector P-type

Electrics-AC


Transistors225. The primary function of a transistor is to transfer resisprevent current flow, depending upon its connection within a current. In other words, its function is as a switching device combination of transmitter and resistor.

226. A semiconductor transistor is a triode (three electrodes, atypically formed of three sections of a single piece of germaniumand one P-type, or vice versa. These are known as junction transp-n-p type transistors, as illustrated at Figure 8-43.

RE 8-44n tors

227. An n-p-n junction transistor comprises two sections separated by a very thin section of P-type germanium (or silicon)the emitter and the other is called the collector. The thin P-typetransistor is basically the same, except that the base is N-type amaterial.



Chapte

xcept that their connections are of the type of transistor (pointing outw. Conventional current flow is, of

een the N-type and P-type material.orrectly biased. The base-emittermust be reverse-biased.

e emitter is negative with respect toA p-n-p transistor will be biased initter is positive with respect to the

lectronic circuits. Their advantagenerate little heat and consequentlynt function. They are cheaper to

n with other semiconductor devices, provide heat sinks and adequate air

Electrics-AC


228. Functionally n-p-n and p-n-p transistors are the same, eopposite polarity. The arrow in the transistor symbol indicatesfor n-p-n and in for p-n-p) and defines the direction of current flocourse, opposite to electron flow.

229. In the transistors described there are two junctions betwTo permit the transistor to conduct both junctions must be cjunction must be forward-biased and the base-collector junction

230. In an n-p-n transistor this condition is achieved when ththe base and the collector is positive with respect to the base. conducting mode (both junctions forward-biased) when the embase and the collector is negative with respect to the base.

231. Transistors have replaced valves, or electron tubes, in eover valves is that they are considerably smaller and they gerequire much less power than a valve performing an equivaleproduce than valves and are generally more reliable. In commotransistors are easily damaged by heat and it is often essential tocirculation, or even air conditioning, for transistorised circuits.





Computer, Binary and Logics

Binary Arithmetic

The AND Gate

The OR Gate

The NOT (Inverter) Gate

The NAND Gate

The NOR Gate

Inhibited or Negated Gates

The Exclusive OR Gate

Positive and Negative Logic

Integrated Circuits



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ogics

very system in whole or in part. Inontrol system is used in which theough computer-controlled electrical

When a computer is an analogueare represented with physical meansis not broken into discrete parts ore or two wires. By contrast digital

ete parts and usually several wires



9Computer, Binary and L

Computer Basics1. On modern large aircraft computers operate practically esome, such as the later series of Airbus aircraft, a fly-by-wire cpilot’s commands are transmitted to the control surfaces only thrsystems.

2. Computers can be analogue systems or digital systems.system, the quantities to be computed, transmitted or controlled that vary in a smooth continuous fashion. The representation separated into set levels and the signal is usually carried on onsystems represent a system quantity by breaking it into discrbundled in a group are needed to carry the signals (digital bus).



Chapte

FIGUAnalogDigital

example of an analogue system.or tape pass from the pick-up to theor tape are converted to continuous by the loudspeaker. The electricalignals. The circuits used to handle



RE 9-1ue and Systems

3. A system that plays records or cassette tapes is an Continuously varying signals of the audio tones from the record amplifier and to the loudspeaker. Sound impulses on the record electrical signals that are amplified and reconverted to soundsignals have provided an analogy or ‘analogue’ of the sound sanalogue information electronically are called linear circuits.



Chapte

re 9-1. The computer is designed tothe human thought process. These

ation, converts it to a form that can information can be data aboutssure, light or it can be commands and tell it where to start. By having from humans via a key-board. The’ are simply called inputs or input

e reasoning function in the humannd decisions are made here. Theseands and information about the the computer, the ‘decide function’sic arithmetic and logical decisionsation of the computer by turning thete times.

ers memory. It must remember what the function or task and the results number of rules used in makinghe system.



Digital Computer Components4. The basic building blocks of a computer are shown in Figuperform the same sequence of steps to execute a command as steps are as follows:

(a) Sense Function. The ‘sense function’ senses informbe read by the ‘decide function’. This inputsurrounding conditions such as temperature, prethat set the machine in a given mode of operationsense elements, the system can receive informationfunctional units that perform the ‘sense functioninterface.

(b) Decide Function. The ‘decide function’ is like thbrain. All the computations, logical operations adecisions take into account the inputs (commsurroundings) and information in the memory. Inis provided by the processor. It performs the barequired by the computer. It also controls the operother functional units on and off at the appropria

(c) Store Function. The ‘store function’ is the computit is asked to do, information on how it performsof what it has done. It must also remember adecisions, performing arithmetic and controlling t



Chapte

y the processor, the ‘act function’e termed the computer output. The format to be displayed on a visual

stem or motor for example.



(d) Act Function. Once a decision has been made bcarries out the command. The ‘act function’ may boutput interface may convert the information to adisplay unit or provide commands to operate a sy



Chapte

FIGUDigitalComp



RE 9-2 Computer onents



Chapte

It has access to the arithmetic logic only memory (ROM) and random

control instructions for computeranged. The random access memoryas required. The clock provides the



The Central Processing Unit (CPU)5. The central processing unit is the brain of the computer.unit (ALU) to carry out calculations, the memory block for readaccess memory (RAM). The read only memory contains theoperation and also data for reference purposes that cannot be chis the memory used by the computer to store and retrieve data timing signal for coding the digital signals.



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FIGUCentraUnit



RE 9-3l Processing



Chapte

tware are bound to be encountered. circuit chips, transistors, resistors,ke up the computer. Software is thehe system. The name software came

ment of bits form codes to identify a byte and 4 bits a nibble. Memory

digital computer the information iserter. A digital output signal to analogue (D/A) converter.

stem, which is made up of only two a unique combination of 1s and 0s, For example, in a particular systemay represent PRESSURE and 0 NOts positive voltage and 0 represents

true.



6. When discussing computers, the terms hardware and sofHardware is the assembly of printed circuit boards, integratedcapacitors and all the wiring, connections and switches that materm applied to the step-by-step programme written to control tabout because most software is written or printed material.

7. Digital signals are termed bits, ‘1s’ and ‘0s’. The arrangeeach number. A group of 16 bits is called a word, 8 bits is calledcapacity in computers is expressed in bytes.

Analogue Signals and Digital Computers8. Where analogue sensors supply input information to a converted to digital form by an analogue to digital (A/D) convanalogue system is converted to analogue form by a digital to an

Binary Numbers9. The language of the computer is the binary numerical sybasic numbers, 1 and 0. Each decimal number is represented bywhich can be made to mean a variety of things to the computer. 1 may represent ON and 0 may represent OFF; in another, 1 mPRESSURE. In what is called a positive logic circuit 1 represenzero or negative voltage; in a negative logic circuit the reverse is



Chapte

system and the numerical languageent conversion between the two. Asding columns to the left of the firstto 9, so any value above 9 is showninning with 10.

so the number of columns required,imal system. The first or right hand

he digit to the left of this equals 21,

f adding the individual values of the Thus the binary number 1111 is

+0+20 = 8+0+0+1 = 9.

lent by repeatedly dividing by twoxample, the binary equivalent of 25

+0+20 = 16+8+0+0+1 = 25).

62--- 3 0 remaining( )=

esult :- 11001



10. Since the language of the computer is the binary numberof humans is the decimal number system, there must be a consistwith the decimal system, the binary system increases value by addigit. In the decimal system there are 10 basic numerals from 0 by adding a column, or columns, to the left of the first digit, beg

11. In the binary system there are only two basic numerals, representing any given decimal number, is greater than in the dec

digit of a binary number is represented by 20, which equals 1. T

or 2. The third digit to the left equals 22, or 4 and so on.

12. To find the decimal value of a binary number it is a case obinary 1s, remembering that binary 0s have a value of zero.

23+22+21+20 = 8+4+2+1 = 15. The binary number 1001 is 23+0

13. A decimal number can be converted to its binary equivaand recording the remainder sequentially from right to left. For eis found as follows:

Hence, the binary equivalent of 25 is 11001. (Proof: 24+23+0

252

------ 12 record 1 remaining( )122------ 6 0 remaining( )= =

32--- 1 1 remaining( )1

2--- 0 1 remaining( ) R= =



Chapte

are the basis for the whole of the letters of the alphabet, words orre in an air conditioning system anderature is measured. For example,perature of 5°C at location number

h logic level and 0 a low logic level., form what is known as a byte. Aown as a word. In many computer

d, subtract, multiply and divide innput data. With binary addition, in each case.

ultiplication & Division

× 0 = 0

× 1 = 0

× 0 = 0

× 1 = 1



14. The digital 1s and 0s that make up the binary numberscomputer code, or language, and may be made to representcomponents of a system. A number may represent the temperatuanother number may represent the location at which the temp0101 (5) and 0111 (7) could be the digital code indicating a tem7, which might be the mixing valve outlet.

15. One binary digit (0 or 1) is known as a bit, 1 being a higA group of bits, such as a five digit binary number for examplegroup of bits forming standard computer information data is knsystems words are of standard length, typically 16 or 32 bits.

Binary Arithmetic16. As with any numerical system it must be possible to adorder for the computer to perform calculations using the isubtraction, multiplication and division there are four basic rules

Rule No. Addition Subtraction M

1 0 + 0 = 0 0 – 0 = 0 0

2 0 + 1 = 1 1 – 0 = 1 0

3 1 + 0 = 1 1 – 1 = 0 1

4 1 + 1 = 10 10 – 1 = 1 1



Chapte

EXAM

into letters of the alphabet, decimale the binary-coded decimal (BCD)

whether the results are correct in



PLE 9-1

Computer Binary Code Systems17. There are a number of code systems used to convert bits numbers, symbols and so on. The principal systems in use arsystem, octal notation and the hexadecimal number system.

EXAMPLE

Addition

Subtraction

Multiplication

Try converting the above binary numbers to decimal, and checkeach case.

0010

00110101------------

0101

01011010------------

1000

00100110------------

1111

01011010------------

0010

00110110------------



Chapte

decimal system from 0 to 9. Each which can be used for alphabeticnge between computers and their

rogramming, which tends to involves decimal numbers to the base eightps.

and its primary function is theystems. The hexadecimal numbersers 0 to 15. A discrete 4-bit binary

by computerised circuits. Each gate will depend upon the state of theter and forms part of an integrated

AND, the OR and the INVERTERR.

inary form the relationship between of a discrete shape, designed so thatuts are shown on the left side of thes 1 or 0, where 1 represents ON, or



18. In the BCD system, four bits represent each digit of thebyte (eight bits) therefore has four unused digit combinations,symbols. The BCD system is used mainly for data interchaperipherals.

19. The octal notation system is primarily used in computer pmanipulation of large quantities of binary numbers. It represent(octal numbers) in a binary language comprised of three-bit grou

20. The hexadecimal number system uses base sixteen representation of very large numbers in computerised memory sare 0 to 9 and the letters A to F, representing the decimal numbnumber represents each hexadecimal number.

Logic Circuits and Symbols21. Logic gates represent a fundamental function performed may have several inputs and will have only one output, whichinputs. The logic gate is a basic function performed by a compucircuit. There are six commonly used logic gates, known as the(NOT) - the basic gates - and NOR, NAND and EXCLUSIVE O

22. For each logic gate there is a truth table that shows in bthe gate inputs and output. Each gate is represented by a symbolit points in the direction of logic ‘flow’. Conventionally the inpgate and the output on the right. Inputs and outputs are shown apositive voltage and 0 represents OFF, or negative voltage.



Chapte

to the gate must be ON to producet (1) at C. If either A or B is 0 the an output when both inputs are in’ gate.

table (b), the series switching circuitresents a simple AND circuit where Relay R1. Figure (d) represents ansitive (I) or negative (O) output at

put then point C will be positive (I)then Point C will be negative (voltst (O) then point C will be negative



The AND Gate23. The AND gate symbolises a situation where both inputs an output. Hence, both inputs (A and B) must be 1 for an outpuoutput at C will be 0. Because the AND gate will only producethe logic 1 state, it is sometimes referred to as the ‘all or nothing

24. Figure 9-4 shows the AND gate logic symbol (a), its truthit represents (c) and a solid-state AND circuit (d). Figure (c) repclosure of both switches A and B will provide a cureent flow inintegrated circuit (IC) where the requirement is to produce a poPoint C. If both diodes are reverse biased with a positive (I) in(i.e. no volts drop across R1). If either diode is forward biased drop across R1). if both diodes are foward biased with no inpu

(i.e. volts drop across R1).



Chapte

FIGUThe A



RE 9-4ND Gate



Chapte

or B) being in logic state 1 (ON) will inputs are in logic state 0 will ther, or both, inputs are in logic state 1

th table (b), the parallel switching (c) represents a simple OR circuit (d) represents an integrated circuitput the dioed A or B as appropriate C will be positive as a result of theen C will be negative. There is nold either A or B on A or B inputs



The OR Gate25. The OR gate represents a situation where either input (A produce an output, that is to say a logic 1 output. Only if bothoutput be 0. Because the OR gate will produce an output if eitheit is sometimes referred to as an ‘any or all’ gate.

26. Figure 9-5 shows the OR gate logic symbol (a), its trucircuit it represents (c) and a solid-state OR circuit (d). Figurewhere any input of (1) will produce an ON (1) output. Figure(IC). In this circuit, if a positive voltage (1) is applied to either inwill be forward biased and current will flow through R1. Pointvolts drop across R1. If there is no current flow across R1 thcurrent flow across R1 when both A and B are negative. Shoubecome positive the output is positive.



Chapte

FIGUThe O

ERTER, gate has only a single inputhe output will be 0, and vice versa.a solid state circuit that could be



RE 9-5R Gate

The NOT (Inverter) Gate27. The third of the three basic logic gates, the NOT, or INVand output and is used to invert a function. If the input is 1, tThe logic symbol, truth table, a simple electrical circuit, and represented by a NOT gate are shown at Figure 9-6.



Chapte

FIGUThe N

er gates, rather than on its own, to

. The output of this gate will be 1 ife added on the output side (the termal outputs of the AND gate will be table, is shown at Figure 9-7.



RE 9-6OT Gate

28. The NOT gate is generally used in conjunction with othalter their basic functions.

The NAND Gate29. The NAND gate is an AND gate with an inverted outputany input is 0. In other words it is an AND gate with a NOT gatNAND is the diminutive of NOT and AND). Thus, the norminverted. The NAND gate symbol, and the derivation of its truth



Chapte

FIGUThe N

e with a NOT gate added, the NOTutput side. The NAND truth tableented by a NAND gate is shown at



RE 9-7AND Gate

30. The symbol is abbreviated, rather than show an AND gat(INVERTER) function is symbolised by a small circle on the oshows the result of the combined gates. A simple circuit represFigure 9-8.



Chapte

FIGUSimpleGate

T added (that is to say a NOT ORy and truth table derivation for the



RE 9-8 NAND

The NOR Gate31. As with the NAND gate, a NOR gate is an OR with a NOgate). It is an OR gate with an inverted output. The symbologNOR gate are shown at Figure 9-9.



Chapte

FIGUThe N

igure 9-10.



RE 9-9OR Gate

32. A simple circuit represented by a NOR gate is shown at F



Chapte

FIGUSimple

mbinations to symbolise sequentialput of an AND or OR gate to invertld be changed further by adding a



RE 9-10 NOR Gate

Inhibited or Negated Gates33. The basic logic gates can be used in a wide range of cofunctions. Having seen how a NOT gate can be added to the outthe output, it follows that the functions of these two gates couNOT gate to one or other of the inputs.



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anner, the NOT gate position being truth table has been derived. It is ‘state symbol’, in the form of a on the logic gate, as shown.



34. Figure 9-11 shows the two basic gates modified in this mshown by a small circle in each case. For each combination thecurrent practice to indicate an inverted input or output by ahorizontal bar above the appropriate letter, in the truth table and



Chapte

FIGU



RE 9-11



Chapte

ilar inputs produce a 0 output andtable for an exclusive OR gate are

FIGUThe ExGate

s, referred to as being binary 1 and OFF, but in reality they represent

nts a higher voltage level and binaryoth negative or one positive and thes represents the lower voltage level,and it.



The Exclusive OR Gate35. The exclusive OR gate symbolises a situation in which simdissimilar inputs produce a 1 output. The symbol and truth shown at Figure 9-12

RE 9-12clusive OR

Positive and Negative Logic36. As previously stated, logic circuit signals are at two levelbinary 0. In simplistic terms these may be regarded as ON orvoltage levels.

37. Positive Logic circuits are those in which binary 1 represe0 a lower one. The actual voltage values may be both positive, bother negative. Negative Logic circuits, in which binary 1 alwayare rarely used and then only when system design parameters dem



Chapte

mbinations of the basic logic gatesuits are Adders, Subtracters, Clocks,

es suggest. An Adder circuit addscuit produces a constant alternatingme reference. Latches and Flip-Flopes the output signal is retained after

ber’ input data.

al at the input when a clock pulse clock pulse causes the output of a

being presented at the input.

an Adder circuit is explained in the

almost always necessary to carry as binary 1 produces binary 10. Two), the 1 of the result being the digit

numbers to be added), plus a thirder order addition. Hence, a numberary numbers.



Basic Logic Circuits38. Every computer and its peripherals uses a variety of codescribed in the preceding text. The five most common logic circLatches and Flip-Flops.

39. Adders and Subtracters perform the functions their nambinary digits and a Subtracter circuit subtracts them. A Clock cirsequence of 1s and 0s to create a square waveform, useful as a ticircuits are used to perform basic memory functions. In both casthe inputs have been removed, enabling the computer to ‘remem

40. The flip-flop circuit remembers or stores the digital signarrives. Each flip-flop can store one bit of information, 0 or 1. Aflip-flop to ‘flip’ to a 1 or ‘flop’ to a 0, depending on which bit is

41. In order to provide an example of a working logic circuitfollowing paragraphs.

42. As has been shown, when binary numbers are added it isdigit to the next higher order column. For example, binary 1 plusingle digits were added, producing a two-digit result (1+1 = 10carried to the next higher order column.

43. An Adder circuit is capable of accepting two inputs (theinput where necessary that is a carried digit from a previous lowof Adder circuits in series is required for the addition of large bin



Chapte

ircuit (a) and the combination ofit. Inputs A and B are the two bitsvious lower order column, whereut CO is the bit to be carried to the

FIGUAdderCircuit

nary addition 1+1. Inputs A and Bried from a lower order column. Byic gate, using the truth tables, it will



44. Figure 9-13 shows the symbol for an Adder logic cEXCLUSIVE OR, AND and OR gates that make up such a circuto be added and input CI is the carried digit from the preappropriate. Output S is the summation of the inputs and outpnext higher order column, where appropriate.

RE 9-13 Logic

45. Let us assume that the circuit is required to make the biwill both be 1, input CI will be 0, since there is nothing to be carapplying these inputs and following the process through each logbe seen that the outputs are CO = 1 and S = 0. 1+1 = 10.



Chapte

cuit containing transistors, diodes,uctors processed on and contained

the actual circuits connecting thefer by a process similar to that usedple one, such as that making up anuter system.

ariety of tasks, for example from aapplications are almost endless. Thetion.



Integrated Circuits

The Integrated Circuit46. An integrated circuit (IC) is a complete electronic cirresistors and capacitors along with their interconnecting condentirely within a single chip of silicon.

47. A process known as photolithography is used to printcomponents. The conducting circuit is etched onto a silicon wain photographic developing. The circuit may be a relatively simAdder, or a highly complex one containing an entire digital comp

48. Integrated circuits can be produced to carry out a wide vsimple timing device to a voltage regulator or an amplifier. The advantages of IC’s are small size, low weight, reliability of opera




ory


Basic Radio The

Phase Notation

Phase Comparison

Wavelength

Radio Wave Emission

Radio Wave Reception

Polar Diagrams

The Ionosphere

Attenuation

Modulation Techniques

Bandwidth

Interference

Q Code and Radio Bearings

Aerials




adiation Patterns


Instrument Landing System Aerial R

Glidepath Transmitter

VOR Aerial Radiation Patterns



Chapte

of conveying electrical energy thanlow for half the total time in oneirection. The pattern of the change

battery. Low frequency AC may beuired for radio and radar equipmentnating current.

FIGUSimple

Basic Radio Theory


10Basic Radio Theory

1. Alternating current (AC) provides a more efficient meansdirect current (DC). With alternating current the electrons fdirection, and for the other half of the total time in the opposite dof direction of the current flow is sinusoidal.

2. It is not possible to produce an AC current flow from a produced by using alternators. The much higher frequencies reqare produced using oscillators, which electronically produce alter

RE 10-1 AC Circuit



Chapte

FIGUCurrenSimple

laced by an alternator. Figure 10-2imes A, C and E no current flows, opposite directions. Note also thattimes B and D.

Basic Radio Theory


RE 10-2t Flow in a

AC Circuit

3. In Figure 10-1 the battery in a DC circuit has been repshows the sinusoidal variation of current flow. Note that at twhilst at times B and D the current flow is at maximum, but inthe amplitude of the wave is measured at its maximum value, at



Chapte

cle. The term frequency is used to UK National Grid system the timency of 50 cycles per second, or 50Hz. The frequencies used for radioto simplify matters the units shown

, some of the power is radiated intorallel to the first, will have a smallnetic energy passes over it. Theied to the transmitting wire. This is

each cycle decreases accordingly.n below.

ilohertz (1KHz)

egahertz (1 MHz)

igahertz (1 GHz)

illisecond (1 m sec)

Basic Radio Theory


4. Time A to E at Figure 10-2 represents one complete cydenote the number of cycles occurring in one second. With thetaken for each cycle is one-fiftieth of a second, giving a frequeHertz (50 Hz). Aircraft AC supply frequency is normally 400 wave transmissions are very much higher than 50 Hz. In order below are used.

5. If a wire (or antenna) is supplied with alternating currentspace as electromagnetic energy. A similar wire (antenna), paalternating current induced in it as the radiated electromagcharacteristics of the induced AC will be identical to those supplthe basis of all radio transmitting and receiving systems.

6. As the frequency increases the time taken to completeConsequently it is necessary to simplify the units of time, as show

1,000 Hz = = 1 K

1,000,000 Hz = = 1 M

1,000,000,000 Hz = = 1 G

= = 1 m

1 103 Hz×

1 106 Hz×

1 109× Hz

11 000,--------------- second 1 10

3–×



Chapte

ineated into convenient bands, eachcteristics, as will be seen shortly.

rosecond (1 µ sec)

Basic Radio Theory


7. The frequencies used for radio and radar systems are delfrequency band having, to a degree, its own unique charaFigure 10-3 shows the frequency bands which need concern us.

= = 1 mic1

1 000 000, ,--------------------------- second 1 10

6–×



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FIGUFreque nge

z

MHz

z

GHz

z

Basic Radio Theory


RE 10-3ncy Bands Band Frequency Ra

VLF

Very Low Frequency

3 –30 KHz

LF

Low Frequency

30 – 300 KH

MF

Medium Frequency

300 KHz – 3

HF

High Frequency

3 – 30 MHz

VHF

Very High Frequency

30 – 300 MH

UHF

Ultra High Frequency

300 MHz – 3

SHF

Super High Frequency

3 – 30 GHz

EHF

Extremely High Frequency

30 – 300 GH



Chapte

magnetic force field, it is convenientotation. Sine waves are constructed on the circumference of a circle, thephase point always lies on the zero

notation. Of course the time taken) is dependent upon the frequency

ments subsequently studied employion.

Basic Radio Theory


Phase Notation8. When considering alternating current flow, or an electroto denote a particular point on the sine waveform using phase nas shown at Figure 10-4, by tracing the progress of a given pointradius of which determines the amplitude of the wave. The 0° amplitude line at the start of the positive half of the sine wave.

9. Considering the construction it is easy to visualise phasefor one complete 360° cycle (the horizontal axis of the graphconsidered, the time being in fact 1/F seconds.

10. Phase notation is important since several of the equipphase comparison techniques in order to determine aircraft posit



Chapte

FIGUConstSine W

ction of a diagram showing a fixedse difference is to remain fixed, the

Basic Radio Theory


RE 10-4ruction of a ave

Phase Comparison11. It is necessary to adopt a precise approach to the construphase difference between two sine waves. Of course, if the phasine waves must be of exactly the same frequency.



Chapte

EXAM

f phase comparison diagrams.

me amplitude but 90° apart.

lid sine wave shown below.

rm starts at the 90° phase point of e wave is at the 180° phase point position B). When the reference s (270 – 90) 180° phase point,

Basic Radio Theory


PLE 10-1

12. Figure 10-5 and Figure 10-6 show two further examples o

EXAMPLE

Draw a diagram to illustrate two sine waves which are of the sa

SOLUTION

First draw one of the sine waves as a reference waveform, the so

In this case the phase difference is 90° and so the second wavefothe reference waveform, (point A). Similarly when the referencthe second waveform will be at its (180 – 90) 90° phase point, (wave is at is 270° phase point the second waveform will be at it(position C) and so on.



Chapte

FIGUTwo Si135° PDiffere

Basic Radio Theory


RE 10-5ne Waves, hase nce



Chapte

FIGUTwo Si45° PhDiffere

aerial it is assumed to travel at the

Basic Radio Theory


RE 10-6ne Waves, ase nce

Wavelength13. Once an electromagnetic wave has left the transmittingspeed of light (C), which is:

300,000,000 metres per second

or



Chapte

finite time taken to transmit oneength (λ). The higher the frequency cycle, and consequently the shorter

he formula:

8 m/sec

Basic Radio Theory


3 x 108 m/sec

14. The wave will travel a given distance in the small butcomplete cycle of energy, and this distance is known as the wavelthe shorter the length of time required to transmit one completethe wavelength.

The relationship between frequency and wavelength is given by t

or

where

F is the frequency in hertz

λ is the wavelength in metres

C is the speed of propagation, 3 x 10

FCλ----=

λ CF----=



Chapte

EXAM

108 m/sec×

30 m------------------------------

107 Hz×

0 Mhz

Basic Radio Theory


PLE 10-2EXAMPLE

Given a wavelength of 30 metres, determine the frequency.

SOLUTION

F=

or F=

F=

or F=

F = or F=

F = F =

Cλ----

Cλ----

300,000,000 m/sec30 m

--------------------------------------------- 3---

10,000,000 Hz 1

10 Mhz 1



Chapte

EXAM

108 m/sec

103× m

---------------------------

8571 105 Hz×

71 Khz

Basic Radio Theory


PLE 10-3EXAMPLE

Given a wavelength of 35 kilometres, calculate the frequency.

SOLUTION

F=

or F=

F=

or F=

F = or F=

F = or F =

Cλ----

Cλ----

300,000,000 m/sec35,000 m

--------------------------------------------- 3 ×35

------

8571 Hz 0.0

8.571 Khz 8.5



Chapte

EXAM

108 m/sec

5 102–× m

----------------------------

67 1010

Hz×

7 Ghz

Basic Radio Theory


PLE 10-4EXAMPLE

Given a wavelength of 4.5 cm, calculate the frequency.

SOLUTION

F=

or F=

F=

or F=

F = or F=

F = or F =

Cλ----

Cλ----

300,000,000 m/sec0.045 m

--------------------------------------------- 3 ×4.-----

6,666,666,666 Hz 0.6

6.67 Ghz 6.6



Chapte

EXAM

108 m/sec

103× Hz

---------------------------

105 m×

km

Basic Radio Theory


PLE 10-5EXAMPLE

Given a frequency of 15 KHz, calculate the wavelength.

SOLUTION

= or

=

= or

=

= or=

= or =

λ CF----

λ CF----

λ 300,000,000 m/sec15,000 Hz

---------------------------------------------λ 3 ×

15------

λ 20,000 m λ 0.2

λ 20 km λ 20



Chapte

EXAM

ds previously shown at Figure 10-3ncies shown.

108 m/sec

109× Hz

---------------------------

102– m×

cm

Basic Radio Theory


PLE 10-6

15. Figure 10-7 reproduces the table of radio frequency banbut now includes the wavelengths pertinent to the various freque

EXAMPLE

Given a frequency of 9.7 GHz, calculate the wavelength.

SOLUTION

= or

=

= or

=

= or=

= or =

λ CF----

λ CF----

λ 300,000,000 m/sec9,700,000,000 Hz ---------------------------------------------

λ 3 ×9.7------

λ 0.031 m λ 3.1

λ 3.1 cm λ 3.1



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FIGUFrequeand AsWavel

ange Wavelength Denomination

km Myriametric

Kilometric

Hectometric

Decametric

Metric

Decimetric

Centimetric (Microwave)

Millimetric

Basic Radio Theory


RE 10-7ncy Bands sociated engths

Band Frequency Range Wavelength R

VLF Very Low Frequency

3 – 30 KHz 100 km – 10

LF Low Frequency

30 – 300 KHz 10 km – 1 km

MF Medium Frequency

300 KHz – 3 MHz 1 km – 100 m

HF High Frequency

3 –30 MHz 100 m – 10 m

VHF Very High Frequency

30 – 300 MHz 10 m – 1 m

UHF Ultra High Frequency

300 MHz – 3 GHz 1 m – 10 cm

SHF Super High Frequency

3 – 30 GHz 10 cm – 1 cm

EHF Extremely High Frequency

30 – 300 GHz 1 cm – 1 mm



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to a transmitting element (aerial orugh the aerial will result in an lower the frequency the larger the

d by means of a radio frequencyng aerial via a modulator, whichcy (more of which later).

ce a simple sinusoid at a particularvariable. Mechanical tuning can bew very rapid variation of frequency

d from a central cathode, move out the anode block. The E field ise. A powerful permanent magnet isic field. Under the influence of theseeen cathode and anode. The anodeies.

Basic Radio Theory


Radio Wave Emission16. If an alternating current at a suitable frequency is fed antenna) of the required dimensions, the current flow throelectromagnetic force field being radiated from the aerial. Theaerial required for efficient transmission (and reception).

17. A suitable frequency for radio transmission is obtaineoscillator. This is then amplified and fed to the transmittisuperimposes the signal to be transmitted upon the radio frequen

Oscillators18. The most common oscillators are circuits, which produfrequency. They may be fixed in frequency (FIXED TUNE) or used to vary frequency fairly slowly. Electronic tuning will alloand fine control. (VCO = Voltage Controlled Oscillator).

The Magnetron19. Magnetrons are cross-field oscillators. Electrons, emitteradially under the influence of an electrical (E) field towardsassociated with a high voltage applied between anode and cathodused to supply a strong magnetic field at right angles to the electrcrossed-fields, electrons travel in curved paths in the space betwblock is made of copper and contains a number of resonant cavit



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ronism with a radio wave travellingr with total feedback; the output ison and a pick-up loop in one of the.

KW are common. Magnetrons areency is determined by the shape andruction of a magnetron.

Basic Radio Theory


20. The rotating cloud of electrons is made to move in synchround the anode block structure. The Magnetron is an amplifieconnected directly to the input. This feedback results in oscillaticavities delivers the signal to an output coaxial line or waveguide

21. Pulse outputs of several MW and CW outputs of a fewoften used as the master oscillator in pulse radars. Output frequsize of the resonant cavities. Figure 10-8 shows the general const



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FIGUSchemMagne

Basic Radio Theory


RE 10-8atic of a tron



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tron amplifier. A powerful electrong electrode on the right. The beamro-magnetic fields alternating at the a waveguide or coaxial cable andisturbs the uniform electron beamn intensity as they move along thee electron beam induce alternating

.

vities (multi-cavity klystron). Theseossible using just two cavities. Ader to stop the electron beam from

d pulse outputs of several MW. It iso the klystron may be a continuousoff, which ensures a coherent pulse amplifiers in radar receivers as they

Basic Radio Theory


The Klystron Amplifier22. Figure 10-9 is a schematic diagram of a very simple klysbeam passes between an electron gun on the left and a collectinpasses through hollow cavities which are tuned to support electsignal frequency. An input signal enters the input cavity alongcreates an alternating field pattern in the cavity. This field dcreating ripples or concentrations of electrons which increase ibeam. As they pass through the output cavity, the ripples on thfields, which are stronger replicas of those fed to the input cavity

23. In practice there are usually a number of intermediate caresult in a larger gain and wider bandwidth than would be pnumber of coils carrying direct current surround the valve in orspreading. A longitudinal magnetic field exerts a focusing effect.

24. CW power outputs of several KW are readily available animportant to notice that in a pulse radar application the input twave. The pulses are created by turning the amplifier on and stream. Klystron amplifiers are entirely unsuitable for use as RFare far too noisy for this application.



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FIGUSchemTwin CKlystro

Basic Radio Theory


RE 10-9atic of a avity n



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by connecting the output cavity totor features only one cavity but thissses through the cavity and is then

egative potential. The stream passeseflector voltage in order to achieveg a coaxial cable or waveguide. No

tron, typically a few tens of mW. Itund in the role of the local oscillator

ground transmitter are that thesignal (see later) of high stabilitylled remotely via telephone lines.

talline substances have the ability torsa. This property is known as the

Basic Radio Theory


The Reflex Klystron Oscillator25. A two-cavity klystron amplifier can be made to oscillatethe input cavity in the correct manner. The reflex klystron oscillacavity is used twice. A stream of electrons from the cathode paturned back by the reflector electrode which is maintained at a nthrough the cavity a second time. It is necessary to adjust the roscillation. The alternating signal in the cavity is extracted alonmagnetic focusing field is required.

26. A low power CW output is available from the reflex klyscan be used as a low power transmitter but is more commonly foin a microwave superhet receiver.

Quartz Crystal Controlled Circuits27. The basic requirements for a VHF aeromobile bandequipment is capable of radiating an amplitude modulated containing low noise levels and that it is capable of being contro

28. Frequency control is by quartz crystal. A number of crystransform mechanical strain into an electrical charge and vice vepiezoelectric effect.



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two conducting electrodes it will beoltage. Conversely, if the crystal iss those electrodes, thus piezoelectricrgy and vice versa. This effect is-ups, and in some headphones anded. Crystalline plates also exhibit aand to many millions of cycles, thet which the plate was cut from theelectric effect the plates also exhibitficient, tuned circuit. Such crystalsd in highly selective filters.

e frequency determining element,ossible to manufacture a crystal that even now these crystals are fragilefrequency is therefore often used, to the final operating frequency. Inquiring multiplication factors of 18and mounting have produced much. The crystal multiplier chain istput power.

the electromagnetic field emanating

Basic Radio Theory


29. If a small plate or bar of such material is placed between mechanically strained when the electrodes are connected to a vcompressed between two electrodes a voltage will develop acroscrystals can be used to transform electrical to mechanical enefrequently used in inexpensive microphones, gramophone pickloudspeakers. For these purposes crystals of Rochelle salt are usmechanical resonance of frequencies ranging from a few thousfrequency depending on the material of the crystal, the angle acrystal and the physical dimensions of the plate. Due to the piezoan electrical resonance and act as a very accurate, and highly efare used in radio equipment in high-stability oscillator circuits an

30. In ground transmitters a quartz crystal is used as thcontrolling the oscillator circuit. Until recently it has not been pwill resonate in the 110 MHz to 136 MHz aeromobile band andand expensive. A crystal on a sub-multiple of the required followed by circuits designed to multiply the oscillator frequencyearly days, crystal frequencies of 5 MHz to 7 MHz were used reor 24 times, but more recently, improved techniques in cutting higher frequencies, requiring far less frequency multiplicationfollowed by amplification stages which generate the necessary ou

Radio Transmitters31. A simplified block diagram of a radio transmitter, and from the aerial, is shown at Figure 10-10.



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nsists of two distinct elements, thehese two force fields exist in planesy at right angles to the direction ofses and collapses at the same rate as

FIGUSimplifDiagraRadio

Basic Radio Theory


32. As the name suggests, an electromagnetic force field coelectric field (the E field) and the magnetic field (the H field). Twhich are at right angles to each other and which are mutualltravel of the radio wave. The amplitude of each force field increathe alternating current, which is producing the radiated signal.

RE 10-10ied Block m of a Transmitter



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ial, would exist only on the surfacealue of current is high at the surfaceses. For a good conductor, such asnt, in effect, to a thin skin at the

t skin depth decreases as frequency frequency of 3 GHz.

tion of the electrical field of anhe electromagnetic wave lies in the. The electromagnetic wave shown

eing vertically polarised, requires aSimilarly, a signal transmitted by al part of the electromagnetic fieldl would need to be horizontal for

er ranges at frequencies up to andve better ranges at UHF and above.

Basic Radio Theory


Skin Effect33. The current, which flows to support the H field in an aerof a perfect conductor. In the case of a practical conductor, the vand falls exponentially as distance within the conductor increacopper, the current value falls very rapidly confining the curresurface. This is the effect known as SKIN EFFECT. Note thaincreases. For copper, skin depth is approximately 1 micron, at a

Polarisation34. The term polarisation describes the plane of oscillaelectromagnetic wave. At Figure 10-10 the electrical field of tvertical plane and the radio wave is said to be vertically polarisedat Figure 10-10 is transmitted from a vertical antenna and, bvertical antenna at the receiver to ensure efficient reception. horizontal aerial would be horizontally polarised (the electricawould oscillate in the horizontal plane), and the receiver aeriaoptimum reception.

35. In general, a vertically polarised signal will achieve bettincluding VHF, however horizontally polarised signals will achie



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ort of electrical resonant circuit, the The following paragraphs describe

Basic Radio Theory


Radio Wave Reception

Electrical Resonant Circuits36. The first stage of most radio receivers comprises of some sworkings of which will need to be understood by the student. how such a circuit is constructed and its method of operation.



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FIGUPhase for PurResista

0-11. From ohms law we know thatcross it, VR, is IN PHASE with the

Basic Radio Theory


Pure Resistive CircuitRE 10-11Diagrams e nce

37. An ac flows through a resistance of R ohms, see Figure 1VR = IR. Thus for a pure resistance, the potential difference acurrent flow through it.

I IN PHASE with VR



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, which is flowing through a coil ofplied voltage by 90°.

FIGUPhase Pure In

Basic Radio Theory


Pure Inductive Circuit38. Figure 10-12 shows the curve for a sinusoidal current (I)inductance L henrys. It can be shown that the current lags the ap

RE 10-12Diagram for ductance

I LAGS VL



Chapte

ge to current gives the resistance R.

– frequency graph for an inductor.

is measured in ohms

Basic Radio Theory


Inductive Reactance. In a pure resistance the ratio of voltaIn a pure inductance the ratio of voltage to current is:

Reactance/Frequency Graph. Figure 10-13 is the reactance

VL

I------- 2πFL=

VL

I-------is called the INDUCTIVE REACTANCE, XL and,

XL 2πFL=

XLα F



Chapte

FIGUEffect FrequeInductReacta

eloped across a pure capacitor ofapplied voltage by 90°.

Basic Radio Theory


RE 10-13of ncy on ive nce

Pure Capacitive Circuit39. Figure 10-14 shows the curve for a voltage (VC) devcapacitance C farads. It can be shown that the current leads the



Chapte

FIGUPhase Pure C

oir in that in a capacitorrent.

age to current is:

Basic Radio Theory


RE 10-14 Diagram for apacitance

NOTE:

Students may find the word ‘CIVIL’ a useful aid memcurrent leads voltage, in an inductor voltage leads cur

Capacitive Reactance. In a pure capacitance the ratio of volt

I LEADS VC



Chapte

frequency graph for a capacitor.

nd is measured in ohms

Basic Radio Theory


Reactance-Frequency Graph. Figure 10-15 is the reactance

VC

I-------

12πFC--------------=

VC

I------- is called the CAPACITIVE REACTANCE, XC, a

XC1

2πFC--------------=

XCα1F---



Chapte

FIGUEffect FrequeCapacReacta

Basic Radio Theory


RE 10-15of ncy on

itive nce



Chapte

FIGUSummResistiCapacInduct

Basic Radio Theory


SummaryRE 10-16ary of Pure ve, itive and ive Circuits



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s, is connected to a capacitor of C to the circuit. Figure 10-17 shows

Basic Radio Theory


Series Circuit40. A coil, of self inductance L henrys and resistance R ohmfarads. An emf of e volts and of variable frequency is connectedthe circuit details.



Chapte

FIGUSeries and Ph

Basic Radio Theory


RE 10-17LCR Circuit ase Diagram



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-17. In this the potential differencee the applied voltage leads the input

asured in ohms.

HASE with the applied voltage. In

llows:

Basic Radio Theory


41. A phase diagram for the circuit is also shown in Figure 10(pd) across L is taken as greater than that across C and thereforcurrent by the phase angle, .

Impedance

42. is called the IMPEDANCE, Z of the circuit and is me

Resonance

43. i.e., the input current is IN P

this special condition, the circuit is said to be at RESONANCE.

As VL = VC then I XL = I XC i.e. XL = XC

Z is a minimum and is equal to R (from above equation).

Resonant Frequency44. The frequency at which XL = XC may be determined as fo

XL = XC

φ

EI---

Z R2

XL~XC( )+2

=

When VL VC φ, 0= =



Chapte

t equal to VC and the impedance ofFor an applied voltage of constante supply changes, see Figure 10-19..

Basic Radio Theory


This value of frequency is denoted by .

Response Curve45. At frequencies other than the resonant frequency, VL is nothe circuit is higher than that at resonance, see Figure 10-18. amplitude, the current (rms value, I) varies as the frequency of thThe curve shown in Figure 10-19 is called a RESPONSE CURVE

2πFL1

2πFC--------------=

FO

FO1

2π LC------------------=



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FIGUVariatiImpedFreque

Basic Radio Theory


RE 10-18on of ance with ncy



Chapte

FIGUVariatiCurrenFreque

ween the two frequencies either sideimum value, see Figure 10-20.

Basic Radio Theory


RE 10-19on of t with ncy

Bandwidth46. The BANDWIDTH, B, of the circuit is the difference betof resonance at which the current has fallen to 0.707* of its max



Chapte

FIGUBandwSeries

Basic Radio Theory


RE 10-20idth in a LCR Circuit

The bandwidth, B = F2 – F1



Chapte

uencies near resonance indicates thed circuit to respond strongly to its

ies either side of resonance. A sharpcated by a flat response curve. Forh L\C ratio.

Basic Radio Theory


NOTE:

Selectivity47. The sharpness of the response curve over a range of freqSELECTIVITY of the circuit. Selectivity is the ability of a tuneresonant frequency and to give a poor response to other frequencresponse curve indicates high selectivity; poor selectivity is indigood selectivity, a circuit should have a low value of R and a hig

1

2------- 0.707=



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FIGUNarroBandwSelecti

Basic Radio Theory


RE 10-21w idth/Good vity



Chapte

FIGUWide Poor S

Basic Radio Theory


RE 10-22Bandwidth/electivity



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is connected across a capacitor of Co the circuit, see Figure 10-23. Thisd has many important applications.

FIGUParalleCircuit

Basic Radio Theory


Parallel Circuit48. A coil, of self-inductance L henrys and resistance R ohms,farads. An emf of e volt and of variable frequency is connected ttype of parallel ac circuit is very common in radio equipments an

RE 10-23l LCR



Chapte

equal to the pd across the capacitor, diagram for the circuit is shown inGS the applied voltage by a phase IC).

FIGUPhase a ParalCircuit

e circuit is at RESONANCE. OnceFO) is given by:

Basic Radio Theory


49. The pd across the coil, the phasor sum of VR and VL, is VC. The supply current is the vector sum of IL and IC. A phaseFigure 10-24. For the condition shown, the supply current LAangle of degrees and the circuit is therefore INDUCTIVE (IL >

RE 10-24Diagram for lel LCR

50. For a certain value of frequency, I is in phase with E, i.e. thagain it can be shown that the value of the Resonant Frequency (

φ

FO1

2π LC------------------=



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ximum and the supply current is ait.

t; namely, the ability of the circuit toant frequency, and to give a pooris a minimum and the impedance ise.g. by altering the value of C, the 10-25 and Figure 10-26.

Basic Radio Theory


51. At resonance, the impedance of a parallel circuit is a maminimum. This circuit arrangement is called a REJECTOR circu

Selectivity52. This is defined in the same way as for a series tuned circuirespond strongly to the required signal, which is at the resonresponse to all other signals. At resonance, the supply current maximum. If the circuit is mis-tuned either side of resonance, supply current increases and the impedance decreases, see Figure



Chapte

FIGUVariatiCurrenFreque

Basic Radio Theory


RE 10-25on in t with ncy



Chapte

FIGUVariatiImpedFreque

Basic Radio Theory


RE 10-26on in ance with ncy



Chapte

uency. The BANDWIDTH of the of resonance, at which the voltage

FIGUBandwParalleCircuit

Basic Radio Theory


Bandwidth53. Parallel circuits reject signals near to the resonant freqcircuit is the difference between the two frequencies, either sidehas fallen to 0.707 of its maximum value. See Figure 10-27.

RE 10-27idth in a l LCR



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achieve this is to connect a resistor

nd of a radio receiver the incomingulation. Demodulation is the point

r wave and therefore made available

fading. This is caused by the radiond skywave) before meeting at therce each other, if they arrive in anti-e, the ionosphere is fluctuating iny continuously, leading to fading in

esentation of the strength of the aerial.

Basic Radio Theory


Bandwidth B = F2 – F1

In some circuits a wide bandwidth is required and one way to across the parallel circuit.

Demodulation54. Having been filtered by the resonant circuit at the front esignal is then amplified before undergoing a process called demodat which the intelligence/information is separated from the carrieto the recipient.

Fading55. Occasionally radio reception suffers from the effects of wave travelling by two alternative routes (e.g. surface wave areceiving aerial. If the two signals arrive in phase they will reinfophase they will cancel out; if therefore, in the above examplintensity and height the path length taken by a skywave will varand out of the signal.

Polar Diagrams56. A transmitter polar diagram is simply a pictorial prelectromagnetic energy field in all directions from the transmitter



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iew associated with a single verticaldio waves omni-directionally in thee circumference of the circle, the(b) that the vertical antenna is notsilence which occurs overhead, fordiagrams do not define the limit ofal strength.

FIGUTransmDiagraVertica

r receiver aerial arrays. Now the transmitted from any point on the in the receiver aerial.

Basic Radio Theory


57. Figure 10-28 shows the polar diagrams in plan and side vantenna. Figure 10-28 (a) shows that the aerial is propagating rahorizontal plane, that is to say that at any point around thtransmitted signal strength is the same. Notice at Figure 10-28transmitting vertically upwards. This gives rise to the cone of example, VOR transmitters. Appreciate that transmitter polar coverage of the transmitted signal, but rather points of equal sign

RE 10-28itter Polar

ms - Single l Antenna

58. It is equally convenient to produce polar diagrams fosituation is reversed in that a signal of a given strength which ispolar diagram will induce a current of constant amplitude to flow



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just the shape of the polar diagram. whereas VOR transmitters produceact made to rotate, but more of this

he mesospheric layer. The gaseous the sun cause electrons to become

uently left with a positive charge arecontains the greatest concentrationsd electrons tend to re-combine with

rs, known as the D, E and F layers.

yers to become established, D, E F1these layers will depend upon suchion intensity increases with increase lower layers.

rge into only two distinctive layers,

heric layers are generally speaking

n entering a layer of ionised gases,

Basic Radio Theory


59. By modifying the antenna configuration it is possible to adThe ILS transmitters give polar diagrams which are lobe shaped,a polar diagram which is known as a limacon, and which is in flater in this section.

The Ionosphere60. The ionosphere exists in the upper atmosphere above tcomposition of the ionosphere is such that ultra-violet rays fromseparated from their parent atoms. The atoms which are conseqknown as ions. The ionosphere is most intense, that is to say it of ions, during the daylight hours. During the night the displacetheir parent atoms, resulting in some degree of de-ionisation.

61. The areas of ionised gases tend to exist in distinctive laye

62. During the daylight hours it is normal for four distinct laand F2, see Figure 10-29. The height and thickness (depth) of factors as latitude, time of the year and sun spot activity. Ionisatin height and therefore the F layer(s) tend to be stronger than the

63. During the hours of darkness the four layers tend to methe E and the F layers, again see Figure 10-29.

64. It is important to appreciate that, at night, the ionosphigher, and are less intensely ionised than during the day.

65. The significance of the ionosphere is that radio waves, owill tend to both bend and weaken.



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FIGUThe IoLayers

as follows:

Basic Radio Theory


RE 10-29nospheric

66. The average heights of the various ionospheric layers are

(a) D Layer 75 km.

(b) E Layer 125 km.

(c) F Layer 225 km.



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comes weaker, or is attenuated, due

mitted in all directions, that is to sayed within a spherical wave front inrea of the wave front also increases.eld strength of the signal decreases. and it can be proved that, to doubleupled.

s surface, the wave will attenuate byworthy of note. Firstly, the rate ofan it is over sea. Secondly, the ratees.

layers it will be weakened by theay be totally attenuated within anreases as the transmitted frequency

Basic Radio Theory


Attenuation67. As any radio wave travels away from the transmitter it beto some or all of the reasons discussed below.

Range from the Transmitter68. Assume that an electromagnetic radio wave is being transit is being transmitted omni-directionally. The energy is containthis case, and as the distance from the transmitter increases the aConsequently as the range from the transmitter increases the fiField strength is inversely proportional to the square of the rangethe range of a transmission the transmitter power must be quadr

Surface Attenuation69. If a radio wave is constrained to travel close to the Earth’virtue of its contact with the surface. Two important facts are surface attenuation is likely to be three times greater over land thof surface attenuation increases as the frequency of signal increas

Ionospheric Attenuation70. As a radio wave travels through any of the ionosphericpositively charged ions. In the extreme case a radio signal mintensely ionised layer. The rate of ionospheric attenuation decincreases.



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turated air would not suffer anywever the atmosphere contains solidnd these particles, droplets and iceequencies.

ttenuated, merely redirected in anred, it may never reach the receiver

be appreciated that the particles,mpared to the wavelength of theonsiderable attenuation under these

lines, that is to say along great circleexist as a result of refraction of the

ough any of the ionospheric layers.y of the radio signal increases.

Basic Radio Theory


Atmospheric Attenuation71. A radio wave travelling through unpolluted and unsaattenuation due to the medium through which its travelling. Hoparticle pollutants and water in both liquid and solid form, acrystals will reflect and scatter radio waves at sufficiently high fr

72. The purist might say that the signal is not in fact aunwanted direction. The fact remains that, if the signal is scatteat a given range, and is therefore of no use.

73. When considering atmospheric signal scatter it shoulddroplets or ice crystals must be of significant size when cotransmitted signal. Radio signals in the EHF band may suffer cconditions.

Refraction74. It is normal to assume that radio waves travel in straight paths. Any departure from such a straight line path is said to radio wave.

75. The principal causes of refraction are discussed below.

Ionospheric Refraction76. Radio waves tend to bend, or refract, as they travel thrThe rate at which this refraction occurs decreases as the frequenc



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travel at a constant speed. In facttly faster over the sea than over the other than 90° will bend slightlyl refraction. Again, the degree of.

waves tend to refract to an extente the curvature of the Earth. There a very useful phenomenon since it

orological conditions, and producesr stage.

ave are discussed below.

Basic Radio Theory


Coastal Refraction77. It was previously convenient to accept that radio wavesradio waves travelling close to the Earth’s surface will travel slighland. In consequence any radio wave crossing a coastline attowards the land mass and this phenomenon is termed coastarefraction decreases as the frequency of the radio signal increases

Diffraction78. In the low frequency and medium frequency bands radiosuch that they remain in contact with the Earth’s surface, despitare several theories as to why this occurs, suffice to say that isincreases the surface range of these frequencies.

Atmospheric Refraction79. Atmospheric refraction sometimes occurs in certain metea situation known as ducting. This is discussed in detail at a late

Propagation Paths80. Six of the possible paths which may be taken by a radio w



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le path for radio waves in the VHFhich remains close to the ground (ae. Conversely, at these frequencies ay refracted to return to the Earth (as

maximum theoretical range of any

er and/or the receiver the greater the

given by the formula:

validate the above formula.

H2 )+

Basic Radio Theory


The Direct Wave81. This is the simplest case, and is normally the only possibband or above. This is because, at VHF and above, a signal wsurface wave) will be totally attenuated over a very short distancsignal which is beamed at the ionosphere would not be sufficientla sky wave).

82. The curvature of the Earth is the factor which limits thedirect wave, see Figure 10-30.

83. Subject to the power transmitted, the higher the transmittdirect wave range.

84. The maximum theoretical range of a direct wave signal is

where

H1 is the height of the transmitter in feet, amsl

H2 is the height of the receiver in feet, amsl

85. Obviously the presence of intervening high ground will in

MAX RANGE (NMS) 1.25 H1(=



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FIGUThe D

hich radio signals can be received ise LF and MF bands diffraction has

h that some will more or less followds, surface wave (or ground wave)

Basic Radio Theory


RE 10-30irect Wave

The Surface Wave86. Fortunately, the maximum range from a transmitter at wnot always limited to the direct wave range. At frequencies in ththe effect of altering the direction of travel of the radio wave sucthe Earth’s curvature. Thus in the LF and MF frequency banpropagation is possible, see Figure 10-31.



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FIGUThe Su

of attenuation suffered by the radioblem of surface attenuation can bet lower frequencies the amount oface.

is approximately three times greaterrface wave ranges of 1000 nms are

s for radio waves in the VHF bandes). The obvious gap in the middleerties of the ionospheric layers may

Basic Radio Theory


RE 10-31rface Wave

87. At the relatively low frequencies concerned, the amount wave as it travels across the surface is not too great, and the proovercome if transmitters of sufficient power are used. Also, adiffraction increases, so increasing the reception range at the surf

88. Remember that the rate of attenuation of a surface wave over the land than it is over the sea. Consequently maximum suachievable over the sea, but only 300 nms over the land.

The Sky Wave89. So far we have considered the normal propagation pathand above (direct waves) and the LF and MF bands (surface wavis the HF band, and it is within this band that the refractive propbe usefully employed.



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d is directed towards the ionosphereertical and the outgoing radio wavein the ionosphere to return to theal and the radio wave is termed the sky wave, see Figure 10-32.

FIGUThe Sk

Basic Radio Theory


90. If a radio wave of a suitable frequency within the HF banit will refract within the ionosphere. As the angle between the vincreases, the radio wave will eventually bend sufficiently withEarth’s surface. When this happens, the angle between the verticcritical angle, and the returning wave is termed the first returning

RE 10-32y Wave



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ansmitted signal, as well as the statehe rate of refraction decreases, andwave must travel further within theproduce a returning sky wave.

f the critical angle, sky waves willower, until the angle coincides with

gnal is shown refracting within the

FIGUEarth TWaves

Basic Radio Theory


91. The critical angle will depend upon the frequency of the trof ionisation of the ionosphere. As the frequency is increased ttherefore, the critical angle increases. This is because the radio ionospheric layer in order for there to be sufficient refraction to

92. If the signal enters the ionosphere at angles in excess ocontinue to return to the Earth’s surface, subject to transmitter pthe Earth’s tangential wave shown at Figure 10-33.

93. Please note that at Figure 10-32 and Figure 10-33 the siionosphere, and not bouncing off the bottom of the ionosphere.

RE 10-33angential



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one ionospheric layer, however thef the tangential wave the theoreticalm from the ‘F’ layer.

of surface attenuation is fairly high. It is quite possible that the firstrange, which is well in excess of 100l be received, and this is termed theitter and the point at which any skyce.

which the first returning sky wavestance, and defines the far end of the

Basic Radio Theory


94. In all of these diagrams it is convenient to show only height of the ionosphere is significant. For example, in the case oskip distance is 1300 nm from an ‘E’ layer refraction and 2500 n

95. Because the signals used lie within the HF band, the rateand therefore surface wave ranges will rarely exceed 100 nmreturning sky wave will not arrive at the Earth’s surface below a nm. There will therefore be an area within which no signal wildead space, see Figure 10-34. The distance between the transmwave returns to the surface of the Earth is termed the skip distan

96. The distance between the transmitter and the point at returns to the surface of the Earth is termed the minimum skip didead space, as shown at Figure 10-34.



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FIGUThe Dand MDistan

aves. Providing that the signal haseturn journey to the Earth’s surfacelti-hop propagation.

Basic Radio Theory


RE 10-34ead Space inimum Skip ce

97. The preceding diagrams have shown only single hop sky wsufficient power there is no reason why it should not make the rvia the ionosphere two or even three times. This is known as mu



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y wave propagation is the diurnally in using a frequency such that theeen transmitter and receiver should so that the strongest possible signal

istance between the transmitter andst returning sky wave (the minimumitted in order to achieve this, for a, is termed the optimum frequency.ere, the maximum usable frequencyble frequency will be slightly lowerorter minimum skip distance. Thisll not cause the receiver to lie within

m usable frequency for a low andituation with the same frequency at. The receiver now lies within the

s shown at Figure 10-37.

Basic Radio Theory


98. The final problem for your consideration in terms of skchanging state of the ionosphere. Obviously there is no pointreceiver lies within the dead space. Conversely the range betwideally be only slightly in excess of the minimum sky wave range,is received.

99. In an ideal world the perfect situation would be for the dthe receiver to be exactly the same as the skip distance for the firskip distance). The frequency which would have to be transmgiven transmitter-receiver distance set of ionospheric conditionsSince it is not possible to predict precisely the state of the ionosphis used rather than the optimum frequency. The maximum usathan the optimum frequency and will therefore give a slightly shensures that slight variations in ionospheric intensity or height withe dead space.

100. Figure 10-35 shows the ideal situation with the maximuintensely ionised layer during the day. Figure 10-36 shows the snight, when the ionosphere is higher and is partially de-ioniseddead space, and a new maximum usable frequency is required, a



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FIGUDay TiWave

Basic Radio Theory


RE 10-35me Sky Propogation



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FIGUNight Wave - same

Basic Radio Theory


RE 10-36Time Sky Propogation Frequency



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FIGUNight Wave with NUsable

be approximately half the daytimealler critical angle, to overcome theuency will refract at a similar rate,rtunately, the less intense level of, despite the lower frequency, which

st by day, since the lower frequencyit is normal for some sky waves tois poses the problem of sky wave

Basic Radio Theory


RE 10-37Time Sky Propogation ew Max Frequency

101. The maximum usable frequency required by night will maximum usable frequency. This lower frequency will have a smgeometry of the higher ionosphere. Additionally the lower freqdespite the partial de-ionisation of the layer. Finally and foionisation will mean that the signal will not be unduly attenuatedis being used.

102. In the LF and MF bands sky waves do not generally exisignals are totally attenuated within the ionosphere. By night survive, since the ionosphere is now somewhat weaker. Thinterference by night in such equipments as ADF.



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wing general statements concerning

nd) waves. Within these bands skyir presence by night reduces useful

ves, which are present both by day

of sight, since the propagation path

t VHF or UHF and now a processce of the Earth and the underside ofd as a reflecting surface. The nameh is the conduit wave. As with VHFuation and consequently ranges oftputs in the region of ten kilowatts.-38.

Basic Radio Theory


Propagation at VLF103. By now you should be absolutely familiar with the follopropagation paths:

(a) LF and MF signals are propagated as surface (grouwaves are not normally present by day, and theequipment range due to night effect.

(b) HF signals are propagated principally as sky waand by night.

(c) At VHF and above, signal range is limited to line is direct wave only.

104. At VLF the wavelength is obviously much longer than awhich is similar to ducting (see below) occurs between the surfathe lowest layer of the ionosphere, which may now be considerewhich now seems to be accepted for this type of propagation patducting the VLF signals travel great distances with little attenmore than 11,000 nm can be achieved with transmitter power ouThe principle of conduit wave propagation is shown at Figure 10



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FIGUConduPropo

Basic Radio Theory


RE 10-38it Wave gation



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ssed, namely direct waves, surfacetherefore useful. The fifth option,sefully employed.

n the VHF, UHF and SHF bands, which is at first sight similar to sky

agation are a marked temperature10-39 shows ducting which, in this. The signal is effectively trapped

attenuation. In this way, when highant VHF communications stations

Basic Radio Theory


The Ducted Wave105. The four propagation paths which have now been discuwaves, sky waves and conduit waves are all predictable and ducting, is unpredictable and ducted waves cannot therefore be u

106. Under certain meteorological conditions, radio waves iwhich normally travel only in straight lines, may behave in a waywaves.

107. The meteorological conditions required for duct propinversion and a rapid decrease in humidity with height. Figure case, is occurring between the surface and a low-level inversionunder the inversion and may travel hundreds of miles with little pressure systems prevail, interference may be heard from distwhich are far beyond the normal direct wave range.



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FIGUThe D

tween two inversion layers.

sly affect receipt of signals and will VHF communication transmissionreceived by aircraft flying above the to short surface ranges because of of 1000 nm.

Basic Radio Theory


RE 10-39ucted Wave

108. Ducting may also occur above the surface of the Earth, be

109. Whilst ducting cannot be used to advantage, it can serioucause unexpected station interference. For example, a VOR ormay be trapped beneath an inversion layer so that it cannot be layer. Similarly, transmission which would normally be limitedtheir direct wave propagation may travel distances well in excess



Chapte

e index of the atmosphere can cause10 GHz. Provided that sufficient

ards the receiver from a volume ofTropospheric scatter offers extendedperating from the upper VHF bands manner is considered to be in thest scattering a beam of sunlight in a

re or less halted research into the scatter for better than line-of-sightary over the horizon radars using

n paths in this section, it is perhapsnd radar systems which you willncy band in which each of these

FIGUIndividEquipmOperaFrequeFreque

Frequency Band

LF

LF

LF/MF

Basic Radio Theory


Tropospheric Scatter110. Small random irregularities or fluctuations in the refractiva scatter of radio waves in the frequency band 200 MHz to transmitter power is available, the signal will be scattered towatmosphere approximately 3 to 8 km above the Earth’s surface. over-the-horizon range for high power communication systems othrough to the SHF band. The maximum range achieved in thiregion of 400 nm. This propagation phenomenon is akin to dudark room or hallway.

111. The advent of satellite communication systems has modevelopment of communications systems relying on troposphericrange at VHF and above, however the development of milittropospheric scatter continuous.

112. Having discussed radio frequency bands and propagatioappropriate at this stage to produce a table of the radio asubsequently study, showing the frequency range and frequeequipments operate, see Figure 10-40.

RE 10-40ual ent

ting ncies and ncy Bands

System Frequency Range

Decca 70 to 130 KHz

Loran C 100 KHz

ADF 190 to 1750 KHz



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HF

VHF

VHF

VHF

VHF

UHF

UHF

UHF

UHF

UHF

aft to Satellite)

ite to Ground)

UHF

SHF

SHF

SHF

SHF

UHF

SHF

Basic Radio Theory


HF Communications 2 to 25 MHz

ILS Markers 75 MHz

ILS Localiser 108.1 to 111.95 MHz

VOR 108.0 to 117.95 MHz

VHF Communications 118 to 136 MHz

ILS Glidepath 329.15 to 335.0 MHz

DME 960 to 1213 MHz

SSR 1030 and 1090 MHz

GPS 1575.42 MHz (L1)

1227.6 MHz (L2)

Satcom (Inmarsat) 1500 to 1600 MHz (Aircr

4000 to 6000 MHz (Satell

Radio Altimeter 4200 to 4400 MHz

Weather Radar 9375 MHz

MLS 5031 to 5091 MHz

ATC Surveillance Radars 600 to 1300 MHz

ATC Ground Manoeuvre Radars 10 to 16 GHz



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gation options for transmittingw necessary to consider the various on to the basic radio wave. Whentelligence is said to be modulated on

ence is to vary the amplitude of thehich is at a lower frequency (the

gence is to vary the frequency of theis technique, logically, is known as

lse modulation.

odulation (AM) are coded, and the below.

Basic Radio Theory


Modulation Techniques113. The preceding section covered the various propaelectromagnetic radio waves from one point to another. It is notechniques which may be employed to superimpose intelligencethis is done the radio wave is termed the carrier wave, and the into it.

114. One way of modulating a carrier wave to convey intelligcarrier wave in sympathy with the modulating wave form wintelligence). This technique is known as amplitude modulation.

115. Another way of modulating a carrier wave to carry intellicarrier wave in sympathy with the modulating intelligence. Thfrequency modulation.

116. The third modulation technique which is considered is pu

Amplitude Modulation117. For convenience the various sub-divisions of Amplitude Mparticular codes which are pertinent to this syllabus are discussed



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rrier wave is simply a radio wavefrequency, and therefore carries notype of signal is ideal for directionnecessary to superimpose a morsenows which transmitter he is tunedN. An unmodulated carrier wave is

FIGUUnmoCarrie

Basic Radio Theory


Unmodulated Carrier Wave118. As the name suggests, unmodulated (or continuous) cawhich is transmitted as a constant amplitude and a constant intelligence. In other words the wave is unmodulated. This finding equipment such as the ADF, although it is obviously identifier onto the wave from time to time, so that the operator kinto. The designation normally given to this type of signal is NOillustrated at Figure 10-41.

RE 10-41dulated r Wave



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f modulation. Here the radio signalred intervals, see Figure 10-42. Thecode. The designation given to thisDBs, the advantage being that all ofdulating wave, since there isn’t one., however the disadvantage is that a into the receiver to make the morse

FIGUKeyed Wave

the amplitude of the carrier waveorm. The significance of simplestant in amplitude and constant in

this type of transmission would be a

Basic Radio Theory


Keyed Carrier Wave119. Keyed (or interrupted) carrier wave is the simplest form ois not continuously transmitted but is switched on and off at desiprimary use of keying is to convey intelligence using the morse type of signal is A1A. It is principally employed on long range Nthe power is contained in the carrier wave, and none in the moRange for a given transmitter power output is therefore enhancedBFO (Beat Frequency Oscillator - see later) must be incorporatedaudible.

RE 10-42Carrier

Simple Amplitude Modulation120. Now to, as it were, real amplitude modulation where consistently varies in sympathy with the intelligence wave famplitude modulation is that the modulating wave form is confrequency, see Figure 10-43. The intelligence normally used for simple steady audio frequency tone.



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carrier wave itself, the audio tone

n of the Aeronautical Information the continuous carrier wave, whichF receiver a nice steady signal forNDB identifier, which in this case isne.

2A with an NDB (where it is keyedpth of modulation is made to vary

Basic Radio Theory


121. By keying either the modulating signal, or the modulatedcan be used to convey simple morse idents.

122. If you check any short range NDB in the COMS sectioPublication you’ll find that it is given as NONA2A. The NON isoccurs between the ident sequences, and which gives the ADdirection finding. The A2A which is tacked onto the end is the achieved by keyed amplitude modulation, using a steady audio to

123. The designation given this type of transmission is either Ato achieve station identification) or A8W in ILS (where the deacross the transmitted lobes).



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FIGUSimpleModul

Basic Radio Theory


RE 10-43 Amplitude ation



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as simple in nature, that is to sayuces a complex waveform which is

eform in that it is varying in bothoan as against a scream). You mayrophone and an oscilloscope handyulation is shown schematically at

ommunications, and J3E when used).

Basic Radio Theory


Complex Amplitude Modulation124. The modulating waveform in the previous paragraph wconstant in amplitude and in frequency. The human voice prodoften modulated on to a carrier wave as intelligence.

125. The human voice produces a complex modulating wavamplitude (a shout as against a whisper) and in frequency (a grwell be either groaning or screaming right now, if you have a micyou can prove the complexity of the wave form! Voice modFigure 10-44.

126. This type of signal is designated A3E when used in VHF cin HF communications (normally on single sided band networks



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FIGUCompAmplitModul

Basic Radio Theory


RE 10-44lex ude ation



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techniques it is necessary to considerplex amplitude modulated signals.

ude of the modulating waveform to as a percentage, or:

is:

wave. The carrier wave (as always with the modulating wave. In thisolts and the minimum amplitude of:

lowest amplitude which governs theof modulation when extreme range

ting Waveform 100×ave (before Modulation)-----------------------------------------------------------

imum Amplitude 100×Minimum Amplitude----------------------------------------------------------

100× 54 %=

Basic Radio Theory


Depth of Modulation127. Before proceeding on to frequency and pulse modulation briefly depth of modulation as it applies to both simple and comQuite simply, the depth of modulation is the ratio of the amplitthe amplitude of the carrier wave prior to modulation, expressed

An alternative formula for determining the depth of modulation

128. Figure 10-45 shows a simple amplitude modulated carrierwith amplitude modulation) is varying in amplitude in sympathycase maximum amplitude of the modulated carrier wave is 10 vthe modulated carrier wave is 3 volts. Using the second formula

129. Since it is the power contained in the carrier wave at its range of the signal, it is normal to reduce the percentage depth reception is required.

Depth of Modulation %Amplitude of the Modula

Amplitude of the Carrier W------------------------------------------------------------------=

Depth of Modulation %Maximum Amplitude Min–

Maximum Amplitude +------------------------------------------------------------------=

Depth of Modulation %10 3–10 3+--------------- 100× 7

13------= =



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FIGUDepthModul

conventional receiver without anyin sympathy with the intelligenceplitude of the carrier wave remains

ble output from a receiver using

the BFO function is selected. Theof which differs from the incomingal and the BFO-generated signal arer output frequencies. The output ofm of the two input frequencies, andHz) which is audible, and this is fede frequency is known as the Beat

Basic Radio Theory


RE 10-45 of ation

The Beat Frequency Oscillator (BFO)130. An amplitude modulated signal is demodulated in a difficulty since the amplitude of the carrier wave is varying waveform. With a NON or A1A (see earlier text) signal the amconstant and therefore it is impossible to achieve an audiconventional demodulation techniques.

131. Figure 10-46 shows how the receiver is modified whenBFO is made to generate an alternating current, the frequency carrier wave frequency by, typically, 2 KHz. The incoming signboth fed to the heterodyne unit which mixes the two to give fouthe heterodyne unit comprises the two input frequencies, the suthe difference frequency. It is only the difference frequency (2 Kto the loudspeaker, producing the audio tone. The differencFrequency.



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FIGUThe BeFrequeOscilla

ployed the amplitude of the carriere carrier wave is made to vary inFigure 10-47. The amplitude of the the frequency of the carrier wave rate of change of the carrier wave

Basic Radio Theory


RE 10-46at ncy tor

Frequency Modulation132. When pure frequency modulation (FM) techniques are emwave normally remains constant, however the frequency of thsympathy with the modulating wave form (the intelligence), see modulating waveform is represented by the amount by whichchanges and the frequency of the modulating waveform by thefrequency.



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FIGUFrequeModul

Basic Radio Theory


RE 10-47ncy ation



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uency modulated. The designation

amplitude modulation, frequency, the necessary modulating power isast benefit is due to the fact that theit is normally an amplitude-orientedlex and the modulated transmissionands (see later). This is why FM

frequency bands would not permitand, as a side benefit they can cover thus provide high fidelity receptioning inside the limit of a spread of 10

nge 118-136.975 MHz and use thegence onto the Carrier Wave. A 10-48:

Basic Radio Theory


133. With VOR the carrier wave is both amplitude and freqgiven to this type of signal (as it applies to VOR) is A9W.

AM Versus FM134. Comparing the technique of frequency modulation withmodulation (FM) transmitters are simpler than AM transmittersrelatively lower and the reception is practically static-free. This lVHF band is practically free from static, and where it is present, disturbance. Of the disadvantages, FM receivers are more compcalls for a much wider frequency band to cover its multi-sidebbroadcasters operate in the VHF band; the congestion in loweraccommodation of the necessary bandwidth. Being in the VHF ba complete range of human audio frequencies (up to 15 KHz) andwhereas in the MF band they would have to be content with stayKHz.

VHF Communications135. VHF Communication Systems operate in the frequency rafrequency modulation technique for superimposing the intellibreakdown of the VHF Communication Band is shown in Figure



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FIGUVHF CommFreque

nal allotments will be determined ional agreement.

e a guard band for the protection l emergency frequency, the nearest cies on either side of 121.5 MHz

nd 121.6 MHz, except that by t it may be decided that the frequencies are 121.3 MHz and

nd movement, pre-flight checking, clearances, and associated

nal allotments

nal allotments

Basic Radio Theory


RE 10-48

unication ncy Band

Block Allotment Of Frequencies (MHz)

World-Wide Utilisation Remarks

118 to 121.4 Inclusive.

International and National Aeronautical Mobile Services

Specific internatioin the light of reg

121.5 Emergency Frequency. In order to providof the aeronauticaassignable frequenare 121.4 MHz aregional agreemennearest assignable121.7 MHz.

121.6 to 121.975 Inclusive.

International and National aerodrome surface communications.

Reserved for grouair traffic servicesoperations.

122 to 123.05 Inclusive.

National Aeronautical Mobile Services.

Reserved for natio

123.1 Auxiliary Frequency SAR.

123.15 to 123.675


Reserved for natio



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nal meeting (Montreal/April 1995)mmunications band should be splitsure was a short term improvementongestion, in anticipation of the system.


nal allotments.


Air-Ground Data Link

Basic Radio Theory


136. The ICAO Special Communications/Operations Divisiodecided that, in order to increase channel capacity, the VHF cofrom 25 to 8.33 KHz channel spacing. It was noted that this meafor regions experiencing severe VHF frequency spectrum cdevelopment and implementation of the future digital VHF radio


International and National Aeronautical Mobile Services.




Reserved for natio






Reserved for VHFCommunications.

Block Allotment Of Frequencies (MHz)

World-Wide Utilisation Remarks



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z capable radios was required forom 1 January 1999. However, as ain July 1998 to delay operational99.

nnel spacing are Austria, Belgium,e United Kingdom (known as 8.33

tates in the ICAO EUR Region, theAO EUR Region.

els for the same airspace sector willt plans of all aircraft not equippedng the provision of an Air Trafficspacing, will be rejected.

egion will still require to be capablee in an environment which uses off- take into account that these offsetany years.

Basic Radio Theory


European Implementation137. As detailed above, the mandatory carriage of 8.33 KHoperation above FL 245 in the ICAO EUR Region with effect frresult of current fitment rates, Eurocontrol took the decision implementation of 8.33 KHz channel spacing until 7 October 19

138. The States that will initially implement 8.33 KHz chaFrance, Germany, Luxembourg, Netherlands, Switzerland and thStates).

139. Although the initial 8.33 States do not comprise all the Smandatory carriage above FL 245 applies to the whole of the IC

140. Parallel operation of 25 and 8.33 KHz spaced VHF channnot be achievable. Accordingly, from 7 October 1999, the flighwith radios compatible with the new channel spacing, requiriService as GAT in the airspace designated for 8.33 KHz channel

141. Aircraft VHF radio equipment used in the ICAO EUR Rof tuning to 25 KHz spaced channels and, additionally, to receivset carrier systems (CLIMAX operation). Airspace users shouldcarrier systems may continue to be used throughout Europe for m



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l not implement 8.33 kHz channeles. Therefore, the UK will file ament for the carriage of 8.33 kHzaircraft intending to fly through UKequirements detailed in the relevant

introduce 8.33 kHz channel spacinge that initial implementation will be

wave (A1A). The difference is thatn, typically one microsecond, whilstillisecond. The designation given to

Basic Radio Theory


UK Implementation142. The UK, although an 8.33 kHz participating state, wilspacing until June 2000 due to related technical dependencidifference with ICAO to the effect that the UK has no requireradios prior to June 2000. However, all non 8.33 kHz equipped airspace above FL 245 are to comply with the flight planning rAIC in order to prevent rejection of flight plans by IFPS.

French Implementation143. Notification had been given that France had intended to above FL 195. However, a decision has now been taken by Francabove FL 245 with effect from 7 October 1999.

Pulse Modulation144. Pulse modulation is similar in principle to keyed carrier the pulses used in typical radar systems are of very short duratiothe period between pulses is relatively very long, typically one mpulse modulation is PON.



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FIGURadar

tudent. The continuous power of acount the time interval between the ‘duty cycle’ of the transmitter).

he individual pulses of a transmittedined code. This allows additional

d Tx, and is called Pulse Coded

lly AM and FM, the carrier wave is called Multiplex Modulation.

Basic Radio Theory


RE 10-49Terminology

145. The terms in Figure 10-49 should be understood by the sradar can be considered to be the average power taking into acpulses, the pulse width and the pulse power (i.e. it allows for the

Pulse Coded Modulation146. In some equipments (e.g. SSR) the time interval between tpulse sequence is made to vary in accordance with a pre-determinformation to be passed between a Tx and Rx, or Rx anModulation.

Multiplex Modulation147. When two types of modulation are used together, typicaable to carry two separate sets of information. This technique is



Chapte

ey apply to some of the equipments

transmitter is, for example 123.2.2 MHz. This is not in fact the case.

ephony (VHF communications)

phony (HF communications).

nterrupted by keyed carrier wave DBs).

keyed amplitude modulation ntifier (short wave NDBs)

h of the modulation is made to

R).

Basic Radio Theory


Modulation Designators148. We can now summarise the modulation designators as thwhich we will study in the Radio Navigation section:

Bandwidth149. We tend to assume that if the published frequency of aMHz, that the frequency transmitted is 123.2 MHz and only 123

A3E Amplitude modulated double side-band radiotel

J3E Amplitude modulated single side-band radiotele

NONA1A Unmodulated carrier wave (NON) periodically i(A1A) to give the Morse identifier (long range N

NONA2A Unmodulated carrier wave (NON) with a simpleperiodically superimposed to give the Morse ide

PON Pulse (radars)

A8W Simple amplitude modulation, however the deptvary (ILS).

A9W Composite amplitude/frequency modulation (VO

NOX.G1D Microwave Landing System



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y modulated signal must necessarilyidth of the transmitter. For perfectame bandwidth. That is to say thatrequencies as that transmitted. Theabove and below the published spot

nsequently the number of channelsectrum, without risk of overlap and

sight that only one frequency isple of simple amplitude modulationulated with a 2 KHz audible tone. be 298 KHz to 302 KHz, giving a

ce the value of the highest frequencytransmitted is the basic carrier waveg waveform.

Basic Radio Theory


150. With a little thought it should be obvious that a frequenccover a band of frequencies, this band being termed the bandwreproduction of the intelligence the radio receiver must have the sthe receiver must possess the ability to accept the same band of fupper and lower limits of the frequency band are equally spaced frequency (except in single side band systems).

151. With FM signals the bandwidth is quite broad, and cowhich can be fitted into any given part of the radio frequency spconsequent interference, is somewhat limited.

152. With amplitude modulation it would appear at first transmitted. Unfortunately life is never that simple. The examillustrated at Figure 10-43 shows a 300 KHz carrier wave modThe range of frequencies actually transmitted in this case wouldbandwidth of 4 KHz.

153. The bandwidth of any amplitude modulated signal is twiof the modulating waveform. The range of frequencies actually frequency plus and minus the highest frequency of the modulatin



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narrower than the bandwidth of aDespite this narrow bandwidth thespectrum, particularly the HF band, from stations on adjacent channelsrefore required of the receiver, thelitude modulation requires a 3 KHzrder to alleviate this problem single

hed that the frequencies actuallyThe intelligence is contained in two the upper side band (USB), and one. The same information is conveyed one or the other without losing anyth, enabling closer channel spacingbands absorb at least 25% of theore of the total transmitter powerincreasing the range at which the

e bandwidth is halved).

(a better signal to noise ratio).

Basic Radio Theory


154. The bandwidth of an amplitude modulated signal will befrequency modulated signal conveying the same intelligence. situation still arises whereby certain parts of the radio frequency have become congested, and that as a result of this, interferenceoften occurs. The greater the bandwidth transmitted, and thegreater the scope for interference. Speech transmission using ampmodulating waveform, whilst music requires about 15 KHz. In oside band (SSB) systems are now widely used.

Single Side Band Transmission155. With amplitude modulated signals we have establistransmitted are equi-spaced about the carrier wave frequency. bands of frequencies, one above the carrier wave spot frequency,below the carrier wave spot frequency, the lower side band (LSB)in the USB as in the LSB, and it is therefore possible to suppressof the intelligence. The effect of this is to halve the bandwidwithout risk of interference. Furthermore, because the side transmitter power, suppressing one of the side bands leaves mavailable for transmission of the carrier wave – significantly transmitted signal can be received.

156. In summary, the advantage of SSB transmission are that:

(a) It occupies less of the available radio spectrum (th

(b) Interference from other transmissions is less likely



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ame power is concentrated into the

ch interference is generally termedsphere is correctly known as static,correctly termed noise. Station

This can occur when stations arequally station interference can occurographic spacing between them is

orm cells cause the storm clouds toarticularly troublesome in the VLF,

nospheric disturbances such as are frequency the greater the problem.

e in summer than in winter and willspheric disturbances will pose morewaves spend more time within the

Basic Radio Theory


(c) Greater range is achieved for a given power (the snarrower bandwidth).

Interference157. Radio interference may arise from several sources. Sunoise. It should be noted that interference generated in the atmowhereas interference generated by man-made equipment is interference was briefly discussed in the preceding paragraphs.operating on adjacent channels and bandwidth overlap exists. Ewhen two stations operate on the same frequency, and the geinsufficient (see ADF and VOR).

Static158. The vast amounts of energy contained within thunderstemit high levels of electromagnetic energy. These emissions are pLF, MF and to a degree the HF bands.

159. In the HF band, static may additionally result from iocaused by sun spot activity. Generally, below VHF, the lower theFor a given location, ionospheric static will be more troublesombe more troublesome at low latitudes than at high latitudes. Ionoof a problem at night than during the day, since at night sky ionised layers.



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given electrical potential strikes anll flow between the airframe and thee same effect. Whenever a currenttatic. Precipitation static, includingd MF bands.

dily occurs at generator and electricns. As we probably all know from.

above. When alternating currentquipment circuitry, the wiring itselfemissions from causing interference

eless telegraphy (Morse) operators.wireless operators disappeared fromtions still survive, and four of themted below.

tation from the aircraft. Sometimeshe station in still-air conditions.

Basic Radio Theory


160. Precipitation static occurs, for example, when rain at a aircraft at a different electrical potential. On impact a current wiwater droplet; dust and sand in the atmosphere can produce thflows it produces a magnetic field which is termed precipitation sthat caused by snow, is particularly troublesome within the LF an

Noise161. Electrical noise is primarily caused by sparking, which reamotor commutators, at relays, and at poor electrical connectioexperience, this noise can seriously interfere with radio reception

162. Electronic noise is troublesome at VHF frequencies andelectron flow occurs at these very high frequencies within the etends to emit electro-magnetic energy. In order to prevent these it is necessary to screen sensitive areas of the equipment.

Q Code and Radio Bearings163. The Q code was introduced as a shorthand to assist wirWith the advent of voice communications networks (telephony), flight decks and with them much of the Q code. Some Q notawhich are particularly pertinent to this part of the syllabus are lis

(a) QDM. The magnetic great circle bearing of the sdefined as being the great circle heading to fly to t



Chapte

ircraft from the station. The term

from the station.

rom the aircraft.

with a combination of electric andcy of the alternating current andby the transmitter. A transmitting

a passive conductor, an alternating as a receiving aerial. The receiveronvert it into a form appropriate to

practical aerial. It is a straighting current is fed to the aerial alonge rod.

Basic Radio Theory


(b) QDR. The magnetic great circle bearing of the aradial is often used as an alternative to QDR.

(c) QTE. The true great circle bearing of the aircraft

(d) QUJ. The true great circle bearing of the station f

Aerials164. Alternating current flowing in a conductor is associatedmagnetic fields. The fields vary in time with the frequenelectromagnetic waves radiate away carrying energy supplied aerial is simply a conductor supplied with alternating current.

165. If the radiated electromagnetic wave is intercepted by current is generated in the conductor. This conductor is actingwhich it feeds is designed to detect the current, amplify it and cthe desired display.

Half Wave Dipole166. The half wave dipole is probably the most commonconducting rod, approximately half a wavelength long. Alternata transmission line, which is usually connected at the centre of th



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FIGUHalf WAerial

a half-wave dipole. The slot aerialcal dipole. Slot aerials are polarisedllel to the rods.

Basic Radio Theory


RE 10-50ave Dipole

Slot Aerials167. A half-wave slot cut in a sheet of metal behaves very likein Figure 10-51 has a radiation pattern exactly like that of a vertiat right angles to the slots, whereas rod aerials are polarised para



Chapte

FIGUHalf WAerial

slot. In these aerials, the slot is cutide, feature in the design of certain

Basic Radio Theory


RE 10-51ave Slot

168. Some types of aircraft suppressed aerial are based on thein the aircraft skin. Half-wave slots cut in the wall of a wavegumicrowave aerials.



Chapte

on design frequency and, at lowsists of a single conductor a quarterrface. This conducting surface actsrformance very like that of the half-bove the conducting surface but inle. The radiation resistance of the

FIGUQuartUnipo

Basic Radio Theory


Quarter Wave Unipole169. The physical length of a half-wave dipole depends frequencies, it may be excessive. The quarter-wave unipole conof a wavelength long mounted at right angles to a conducting suas a reflector and creates the image of the unipole leading to a pewave dipole. Obviously radiation is only present in the space athis space the radiation pattern is identical to that of the dipounipole is about half that of the dipole.

RE 10-52er Wave le Aerial



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mple modification, which improvesic radiating elements. Parasites aree signal transmitted from the dipoleeived. By adjusting the lengths anderference can be encouraged in onenger than a half wavelength placeddipole. Such a parasite is called adirection of the parasite and such a

FIGUThe Ya

eflector and a number of directors.w tens of degrees are available.

Basic Radio Theory


Parasites170. The half-wave dipole is not particularly directional. A sithe directional qualities of a dipole, involves the use of parasitconducting rods placed near to the energised (driven) dipole. Thinduces currents in the parasites, which reradiate the energy recspacings of the parasites relative to the dipole, constructive intdirection and destructive interference in others. A parasite, lonear to a dipole enhances the radiation in the direction of the reflector. A short parasite tends to enhance the radiation in the parasite is therefore called a director.

The Yagi ArrayRE 10-53gi Array

171. The Yagi aerial array is formed from a driven dipole, a rThe polar diagram is similar in all planes and beamwidths of a fe



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tion resistance of the driven dipole.stem losses. For these reasons, theer in a Yagi array. It has a higherrasites, offers a reasonable match to

FIGUThe Fo

g device or transducer between free-tion. When the radar transmits, the a beam of the desired shape, whichcollects the energy contained in the

Basic Radio Theory


The Folded Dipole172. An effect of the parasite elements is to reduce the radiaThis can cause serious mismatch problems and an increase in syfolded dipole, shown in Figure 10-54, is often used as the drivradiation resistance which, when reduced by the effect of the pathe feeder.

RE 10-54lded Dipole

Radar Aerials173. The basic function of a radar aerial is to act as a couplinspace propagation, and guided-wave (transmission line) propagafunction of the antenna is to concentrate the radiated energy intopoints in the desired direction in space. On receive, the aerial signal and delivers it to receiver.



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ls used for communications at other between radar aerials and mostt a narrow beam of energy in any focussing the transmitted energy by a searchlight beam is formed, or bydiation from a number of individual may be achieved by mechanical or

e. The aerial may be designed to all planes; this is usually called arrow in azimuth, but much wider in achieved by using an aerial with anw beam, and narrow in the other

e in the radiation pattern of smallrections other than that of the mainey represent wasted power and can

Basic Radio Theory


174. Fundamentally radar aerials do not differ at all from aeriafrequency bands. The most important practical differencecommunications aerials, is the requirement to be able to direcdirection at will. The narrow beam property is often achieved bymeans of a specially designed reflector, in much the same way asutilising an array configuration in which it is arranged that the raelements adds up in one direction only. Scanning of the beamelectronic means.

175. The shape of the beam is a function of the aerial shapproduce a narrow-beam pattern with a uniform beamwidth inpencil beam. Other applications may require a beam which is naelevation, or vice-versa. Such a beam is called a fan beam, and isaperture which is wide in the dimension requiring the narrodirection.

Sidelobes176. An unavoidable feature of radar aerials is the existencsubsidiary beams, known as sidelobes, which are generated in dibeam. The existence of sidelobes is usually undesirable since thgive rise to spurious echoes.



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a reflector because of two basicr a simple parabola. Firstly, all thebola, are reflected as parallel rays.ll equal. Thus the reflected wave iss does not imply that the beam willtenna, the beam will diverge. Thet source of energy at the focus into a

FIGUProperParabo

Basic Radio Theory


Parabolic Antennas177. The paraboloid or parabolic dish is widely used as geometrical properties. These are illustrated in Figure 10-55 forays from a fixed point (called the focal point, F) to the paraSecondly, in Figure 10-55, the path lengths FXA, FYB, etc. are amade up of parallel rays which are all in phase. Notice that thinot diverge. In fact, except in the region very near to the ansignificant effect of a parabolic reflector is that it converts a poinplane wavefront of uniform phase.

RE 10-55ties of the la



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s of the paraboloid and making anctor (Figure 10-56 (c)). Notice thatrture is greater. Another common, and the parabolic ‘cheese’ reflectorlat plates.

FIGUTypes Reflect

om a radar aerial, depends on theppropriately configured aerials. For mapping radars, produces a beamrovide equal strength returns from

Basic Radio Theory


178. The beam shape may be modified by cutting away part‘orange-peel’ (Figure 10-56(d)) or a ‘truncated’ ‘paraboloid’ reflethe beam is narrower in the plane in which the reflector apeconfiguration is the parabolic cylinder shown in Figure 10-56 (a)which is a narrow parabolic cylinder enclosed on either side by f

RE 10-56of Parabolic or

Cosecant Squared Aerial179. The shape of the vertical radiation pattern required frfunction of the radar. Special beam shapes can be produced by aexample the cosecant squared reflector, widely used in airbornewhich is narrow in azimuth but wide in elevation, in order to psimilar ground targets at different ranges (Figure 10-57).



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FIGUCosecRadiat

distorting the upper portion of arange targets than towards those at

Basic Radio Theory


RE 10-57ant Squared ion Pattern

180. This specialised type of pattern is often produced by parabolic reflector, so that more energy is directed at the longer short range (Figure 10-58).



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FIGUCosecReflect

Basic Radio Theory


RE 10-58ant Squared or



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ry aerial placed at the focal point, orng the axis. The feed element ofteninder a length of slotted waveguide.

FIGUFeeds ParaboReflect

inciple used to reduce the length of

Basic Radio Theory


Feeds for Parabolic Reflectors181. Parabolic Reflectors are usually energised by an elementain the case of a parabolic cylinder, a line of such aerials placed aloconsists of a dipole or a waveguide horn, or for the parabolic cylTypical arrangements are shown at Figure 10-59.

RE 10-59for lic ors

The Cassegrain Aerial182. The aerial shown at Figure 10-60 is developed from a proptical telescopes, and is known as the Cassegrain Aerial.



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FIGUThe CAerial

de horn. The waveguide energy islic reflector. This ‘double-focussing’from a conventional aerial of similarr mounting, the rear feed means that waveguide losses are thus reduced.

ations where a low-noise amplifier making the waveguide run as short

Basic Radio Theory


RE 10-60assegrain

183. A parabolic reflector is fed from the rear by a waveguireflected from a small convex sub-reflector into the main paraboresults in a beam considerable narrower than would be achieved dimensions. As well as the advantages of smaller size and simplea shorter waveguide run between aerial and radar is required, andThe Cassegrain configuration is often used in low-noise applicmay conveniently be mounted directly behind the main reflector,as possible.



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are spaced half a wavelength apartme directions and cancels in othersl array is the linear broadside array,erials (Figure 10-61).

FIGUThe BrArray

half wavelength slots cut in the walls equal amounts of in phase energy.cally the slots would be cut in the

Basic Radio Theory


Slotted Waveguide Arrays184. It has already been shown that when a number of aerialsand fed in phase, the energy radiated from the aerials adds in sodepending on the phase relationships involved. This type of aeriawhich gives a beam at right angles to the line of the constituent a

RE 10-61oadside

185. A similar effect may be achieved by means of a series of of a length of waveguide, and so arranged that each slot radiateSuch a device constitutes a microwave broadside array. Typinarrow dimension of the waveguide as shown at Figure 10-62.



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FIGUInclineWaveg

l Radiation

ch aid which provides the pilot withach in bad weather.

nded centreline of the instrumenttreline.

ent slope (normally three degrees),proach.

Basic Radio Theory


RE 10-62d Slot uide Array

Instrument Landing System AeriaPatterns186. The Instrument Landing System (ILS) is a runway approaaccurate guidance both in azimuth and elevation during an appro

ILS Ground Equipment187. The ground installation consists of:

(a) A localiser transmitter which defines the exterunway, and indicates any deviation from this cen

(b) A glidepath transmitter which defines a safe descand again indicates any deviation from this safe ap



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eacon transmitters for a typicallations the inner marker is omitted,

specified ranges from the runway

wide and four metres high, and isd of the instrument runway, see

FIGUILS AeLocatio

the extended centreline, it may beffset ILS. In this case, the QDM ofM by a few degrees.

Basic Radio Theory


(c) Normally two (occasionally three) marker binstallation. That is to say that with many installeaving only the middle and outer markers.

The primary purpose of the markers is to definethreshold.

Localiser Transmitter188. A localiser antenna array is approximately 25 metres normally situated some 300 metres beyond the upwind enFigure 10-63.

RE 10-63rial ns

189. Should it not be possible to locate the localiser aerial onlocated to one side of the runway, giving what is known as an othe localiser centreline will differ from the runway centreline QD



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-magnetic energy (designated A8W)rlap area, the equisignal, defines the

FIGUILS LoRadiat

Basic Radio Theory


Localiser Radiation Pattern190. The localiser transmits two overlapping lobes of electroon the same VHF carrier wave frequency. The centre of the oveILS QDM, see Figure 10-64.

RE 10-64caliser ion Pattern



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ude modulated at 90 Hz, whilst theulation of both the lobes is made tobes. The airborne localiser receiverves. When they are of equal depth

eedle is deflected in the appropriatehe greater the displacement of the

with normal glidepath transmittersnces of:

(centre) line.

entre) line.

Basic Radio Theory


191. The lobe on the pilot’s left during the approach is amplitright lobe is amplitude modulated at 150 Hz. The depth of modvary, being greatest at the centre and least at the sides of the locompares the depth of modulation of the 150 Hz and 90 Hz wathe localiser needle will be centralised.

192. When the depth of modulation is uneven the localiser ndirection. The greater the difference in modulation depths, tlocaliser needle from the centre of the instrument.

193. In the United Kingdom ILS localisers which are associatedprovide coverage from the centre of the localiser antenna to dista

(a) 25 nm within plus or minus 10° of the equisignal

(b) 17 nm between 10° and 35° from the equisignal (c

As illustrated at Figure 10-65.



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FIGUILS LoCover

ciated with steep angle glidepathenna to distances of:

(centre) line.

Basic Radio Theory


RE 10-65caliser age

194. In the United Kingdom ILS localisers which are assotransmitters provide coverage from the centre of the localiser ant

(a) 18 nm within plus or minus 10° of the equisignal



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entre) line.

mal glidepath transmitter should beroximately three degrees above thene which defines an angle from the

reas, even on the approach side, caneceived. Such use should not besion for localiser Back Beams to bet be ignored.

on a mast approximately ten metreseline and 300 metres upwind of the

two overlapping lobes of electro-uency. The frequency range used toe lobes overlap in the vertical plane.z and 150 Hz. Figure 10-66 shows glidepath at a typical value of 3°

.

Basic Radio Theory


(b) 10 nm between 10° and 35° from the equisignal (c

195. As far as the above coverage areas are concerned, a norconsidered to be one which produces a glidepath angle of apphorizontal, and a steep glidepath should be considered to be ohorizontal of 4° or more.

196. Pilots are warned that use of the localiser outside these alead to False Course and Reverse Sense indications being rattempted. In particular it must be noted that there is no proviused in the United Kingdom, and any indications from them mus

Glidepath Transmitter197. There are two glidepath aerials which are both mounted tall, which is displaced some 150 metres from the runway centrthreshold markings.

Glidepath Radiation Pattern198. As with the localiser, the glidepath transmitter emits magnetic energy (designated A8W) on the same carrier wave freqglidepath transmissions lies in the UHF band, and in this case thAgain the lobes are continuously amplitude modulated at 90 Hthe idealised radiation pattern with the equisignal defining theabove the horizontal plane passing through the touchdown zone



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FIGUILS GliRadiat

face ground-reflected waves result,roduce additional equisignals andhs will be situated above the maingerously low during the approachns that the aircraft is flying a false

Basic Radio Theory


RE 10-66depath ion Pattern

199. Since the lower (150 Hz) lobe lies adjacent to the surgiving side lobes (Figure 10-67). These side lobes may pconsequently false glidepaths. Fortunately, these false glidepatglidepath and cannot therefore result in an aircraft flying danshould the false glidepath be inadvertently followed. Indicatioglidepath are listed below:



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res the glidepath from below. Thiswest) will be the first one to beed a warning to pilots emphasising

around the world where proceduresom above.

t slope which is inclined at least 6°r 5° to the horizontal for a 2.5°t least twice the expected value.

S approach shows check heights andcons (see later) if appropriate. If aaltimeter will verify this. A typical00 feet (QFE), whereas on the first (QFE) or above.

Basic Radio Theory


(a) During a normal ILS procedure, the aircraft captubeing the case, the true glidepath (being the lointercepted. The Civil Aviation Authority has issuthat special care must be taken at certain airfields are published involving capture of the glidepath fr

(b) The first (lowest) false glidepath will give a descento the horizontal for a normal 3° glidepath, oglidepath. This will result in a rate of descent of a

(c) The approach plate used by the pilot during an ILaltitudes at the marker beacons, and locator beafalse glidepath has been captured, a check of the check height over the outer marker would be 15false glidepath the altimeter would read 3000 feet



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FIGUILS FalGlidep

allations) is provided through an arc of 10 nm from the threshold, as

Basic Radio Theory


RE 10-67se aths

200. Glidepath coverage in azimuth (for United Kingdom instof 8° on either side of the localiser centreline out to a rangeillustrated at Figure 10-68.



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FIGUILS GliCoverHorizo

h, the coverage is reduced to a range side of the localiser centreline.

rc of 1.35° above the horizontal toglidepath installation, and are basedation) is provided through an arc.45 and glidepath angle x 1.75, as

Basic Radio Theory


RE 10-68depath age - ntal Plane

201. For glidepath transmitters which produce a steep glidepatof 8 nm from the threshold, again through an arc of 8° on either

202. Glidepath coverage in elevation is provided through an a5.25° above the horizontal. These figures apply to a standard 3° on the formulae which state that glidepath coverage (in elev(measured from the horizontal) of between glidepath angle x 0illustrated at Figure 10-69.



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FIGUILS GliCoverVertica

Basic Radio Theory


RE 10-69depath age - l Plane



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limits can lead to intermittent andlidepath at very shallow approach

be attempted when the promulgated

ch angle is so shallow as to put theown of 10 nm or more.

correct deflection sensitivity to one or other problems and can lead toituation exists a warning will beriate columns of the COM2 section

vertically upwards. Figure 10-70inner marker is not often used theseHz. Notice from Figure 10-63 andons because of the narrow extent of

Basic Radio Theory


203. Pilots are warned that use of the glidepath outside theseincorrect indications being received. In particular, use of the gangles, (that is below 1500 feet aal at 10 nm range), should only glidepath intercept procedure requires such use.

204. The glidepath indication must be ignored if the approaaircraft at a height of 1000 feet or below at a range from touchd

205. Certain glidepaths in the United Kingdom do not exhibitside of the localiser course line. This effect is caused by terraininadequate fly up indications being received. When this spromulgated by NOTAM and subsequently appear in the appropof the UK AIP.

Marker Beacons206. Marker beacons radiate fan-shaped patterns of energyshows an installation using three marker beacons, although the days. All market beacons transmit on a set frequency of 75 MFigure 10-70 that there is no interference between adjacent beacthe radiation patterns along the glidepath.



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FIGUILS MaRadiat

Basic Radio Theory


RE 10-70rker Beacon ion Patterns



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ed with dots and/or dashes at givented with a given marker beacon, thet Figure 10-71.

FIGUMarkeAural Indicat

way centrelines to denote reportingrs radiate a fan-shaped pattern on aransmitted by an airways marker isral identifier is a single Morse letterarker) light on the aircraft marker

nisation with the audible dashes at

ate dots and dashes.

ronisation with the audible dots haracters per second.

onisation with the audible dots at

Basic Radio Theory


207. The marker beacon transmissions are amplitude modulattones. As the aircraft flies through the radiation pattern associapilot will receive both aural and visual indications as described a

RE 10-71r Beacon - and Visual ions

Airways Fan Markers or Z Markers208. Marker beacons are still sometimes found straddling airpoints. As with the ILS marker beacons, the airways fan markefixed carrier wave frequency of 75 MHz, however the power tconsiderably greater to facilitate high altitude reception. The auof high pitch tone (3000 Hz) which activates the white (inner mbeacon panel.

Outer Marker Aural: Low Pitch (400 Hz) Dashes

Visual: A Blue light flashing in synchrothe rate of two per second.

Middle Marker Aural: Medium pitch (1300 Hz) altern

Visual: An amber light flashing in synchand dashes at the rate of three c

Inner Marker Aural: High Pitch (3000 Hz) dots.

Visual: A White light flashing in synchrthe rate of six per second.



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which gives accurate bearings withomparison between two waveforms.

in a manner previously described as say that the single carrier wave is

0 Hz waveform, the reference signalt a given range from the station willss of the aircraft bearing from the

onally), however the signal strength in time. The polar diagram for theigure 10-72.

Basic Radio Theory


VOR Aerial Radiation Patterns209. VHF Omni-directional Radio Range (VOR) is a system reference to ground-based stations using the principle of Phase C

Principle of Operation210. VOR stations transmit a carrier wave which is modulatedA9W in the section entitled Modulation Techniques. This is toboth frequency and amplitude modulated at the same time.

211. By frequency modulating the carrier wave with a simple 3is achieved. This signal is so named since all airborne receivers areceive a reference signal which is at the same phase, regardlestation.

212. The VOR station transmits in all directions (omnidirectivaries depending on the bearing from the station at a given pointVOR transmitter, which is known as a limacon, is illustrated at F



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FIGULimacoDiagra

per second and this has the effect ofeceiver. The phase of the amplitudeceiver from the ground station. The

Basic Radio Theory


RE 10-72n Polar m

213. The limacon itself is rotated at the rate of 30 revolutionsamplitude modulating the carrier wave arriving at an airborne rmodulated signal will depend upon the bearing of the airborne reamplitude modulated is therefore known as the variphase signal.



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flat terrain to minimise site errors.ployed to transmit the VOR signal.on and produces a signal which is hilly terrain.

different from conventional VOR,he airborne receiver will operate onl is amplitude modulated at 30 Hz, Because this is the reverse of a

de to lead the reference signal by aR ground station.

t 50 antennae surrounding a single signal, whilst the circle of antennaeolutions per second (30 Hz). From

vancing and retreating at 30 Hz – inp between the doppler shift and then a bearing of 000° (M) from theency (30 Hz), at 180° (M) from the0° and so on.

Basic Radio Theory


Doppler VOR214. Conventional VOR transmitter aerials should be sited onIf such a site is not available, a complex aerial system may be emThis type of station is know as a Doppler VOR (DVOR) beacreasonably free of site errors even when the transmitter is sited in

215. The way in which the bearing signal is produced is quitethe received signals are indistinguishable from each other and teither with equal facility. In Doppler VOR the reference signawhilst the bearing signal is frequency modulated at 30 Hz. conventional VOR, the bearing (or variable) modulation is maphase angle equal to the aircraft’s magnetic bearing from the VO

216. The Doppler VOR transmitter comprises a circle of abouomni-directional antenna. The latter transmits the AM referenceare sequentially energised in an anti-clockwise direction at 30 revany given direction, it will appear as though the transmitter is adother words there will be a Doppler shift. The phase relationshisteady reference signal is arranged to be zero when received otransmitter. Since both signals have the same modulating frequVOR the phase difference will be 180°, at 270° (M) it will be 27





Piston Engine Principles and Construction

Introduction

Working Cycles

Power and Efficiencies

Engine Design Types

Mechanical Components



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and

ed at one end and in which a pistonsmitted to the crankshaft by meanson pin and to the crank at the otherhaft. The function of the connectingiston into rotary movement of the

ead, includes an inlet valve and anhe inlet valve admits a combustiblealve allows the waste products ofs attached to rotating gear wheels,m drive gears are meshed to a spur

ansion of hot gas, heated by thel engine, by an electrical dischargercraft piston engines usually employectrical energy to the sparking plug,



11Piston Engine PrinciplesConstruction

Introduction1. A single cylinder piston engine consists of a cylinder, closis free to slide up and down. The movement of the piston is tranof a connecting rod, hinged to the piston at one end by the gudgeend by a crank pin. The crank arm is fixed to the rotating cranksrod and crank is to convert the reciprocating motion of the pcrankshaft.

2. The closed end of the cylinder, known as the cylinder hexhaust valve, each of which is held closed by strong springs. Tmixture of air and fuel to the cylinder whilst the exhaust vcombustion to escape to atmosphere. Push rods, driven by camopen the valves through the lever action of rocker arms. The cagear on the crankshaft.

3. The force to move the piston is derived from the expcombustion of fuel. The fuel is ignited, in the case of a petrocreating a large spark across the electrodes of a sparking plug. Aia magneto, driven from the engine crankshaft, to provide the elwhich is located in the upper part of the cylinder.



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m of a single cylinder, spark ignitionity, the inlet and exhaust valves are two cams would be mounted on a



4. These features are illustrated in the cross-sectional diagrapiston engine illustrated at Figure 11-1. In this diagram, for clarshown as being driven from separate cam drives. In practice thecommon camshaft.



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FIGUCross-througCylindEngine



RE 11-1Section h Single er Piston



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which the heat energy is converteduel-air mixture is inducted into thee the fuel will readily ignite. At closeand the heat of combustion causes aste products of combustion are then

chanically related operations eachression, power (the result of ignitionf the four-stroke engine is describedroke cycle.



Working Cycles

The Four-Stroke Cycle5. In aircraft piston engines the sequence of operations by into mechanical energy, is known as the four-stroke cycle. A fcylinder and compressed to raise the temperature to a point wherto maximum compression an electrical spark ignites the mixture rapid increase in pressure, driving the piston downwards. The waejected during the final stage of the cycle.

6. The cycle can be divided into four separate, but merequiring one full stroke of the piston. They are induction, compand combustion) and exhaust. The basic principle of operation oin the following paragraphs. Figure 11-2 summarises the four-st



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FIGUFour-S



RE 11-2troke Cycle



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et valve is open, and the piston isnown as the top dead centre (TDC) dead centre (BDC) position. The

such that cylinder pressure is lowercylinder through the inlet pipe, oron system.

e induction stroke depends upon air mixture in the inlet manifold isylinder at the start of the inductionh sides or sharp corners that impedeis dependent on the mass of mixtureimise the resistance to the flow ofder by means of superchargers or

d induction stroke, the piston nowthe connecting rod by the inertia of-driven push rod and the mixture is the mixture into a smaller space, itsessed it is heated adiabatically, andre rises to a higher value than thaturpose of the compression stroke is

hich the fuel will readily ignite and



The Induction Stroke. During the induction stroke the inltravelling down the cylinder from the upper extreme of travel, kposition, to the lower extreme of travel known as the bottomincreasing volume above the piston causes a pressure decrease than ambient atmospheric pressure. Air therefore enters the manifold, where fuel is added from the carburettor or fuel injecti

7. The mass of mixture that enters the cylinder during thnumber of factors. Whilst the inlet valve is closed the fuel-aeffectively at rest and will not instantly begin to flow into the cstroke. This problem is accentuated if the inlet manifold has rougthe flow of the mixture. Since the power achieved by the engine being burnt, it is normal to polish the inlet manifold to minmixture. Additionally the mixture can be blown into the cylinturbochargers, as discussed in a later section.

The Compression Stroke. Having completed the downwarstarts moving upwards from the BDC position, driven through the crankshaft. The inlet valve is closed by the action of the camcompressed in the enclosed space of the cylinder. By compressingpressure and temperature is increased. As the mixture is compralso gains heat from the hot surroundings. The pressure therefowhich would be expected from volumetric reduction alone. The pto raise the temperature of the fuel-air mixture to a value at wburn efficiently.



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e compression stroke, a spark at thee mixture, the intense heat rapidly

t past TDC. The mixture burns, andthe cylinder to complete the powerrough the connecting rod and crank

er stroke the piston now starts up action of its cam and push rodhere through the exhaust manifold.arp corners which would otherwiseny exhaust gases remaining in theof fuel-air mixture at the start of theis known as scavenging.

ke cycle is that the inlet valve opens at the end of the induction stroke. closes at TDC at each end of the to account for a number of factors.

he flow rate of gases past the valvesen before the piston commences its



The Power Stroke. Just before the piston reaches TDC on thspark plug ignites the mixture. As the flame spreads through thraises the pressure to a peak value, ideally when the piston is juspressure falls as the hot gases expand, forcing the piston down stroke. During this stroke the piston is driving the crankshaft thmechanism.

The Exhaust Stroke. Having descended to BDC on the powthe cylinder once more. The exhaust valve is opened by themechanism and the combustion gases are forced out to atmospIdeally the exhaust manifold should be polished and free from shimpede the exit of the burnt gases. This is important, since acylinder at the end of the exhaust stroke would impede the entry subsequent induction stroke. Expulsion of the combustion gases

Valve and Ignition Timing8. The inference in the preceding description of the four-stroat TDC at the start of the induction stroke, and closes at BDCSimilarly it appears that the exhaust valve opens at BDC andexhaust stroke. In fact it is necessary to adjust the valve sequence

9. During the time that the valves are opening and closing tis less than maximum and so the valves must be arranged to opstroke and close after it has completed its stroke.



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g the induction stroke is determined as the piston reached BDC on theheric pressure, because of the inertiaer, and thus increase the mass of thest part of the compression stroke. Inalised when the inlet valve closes.

TDC on the exhaust stroke therein due to gas inertia. By keeping thestroke, residual gas pressure ensures

nd exhaust valves during the four-



10. The mass of fuel-air mixture that enters the cylinder durinby the instant of closure of the inlet valve. If this were to closeinduction stroke the cylinder pressure would be less than atmospof the gas. In order to allow a little more mixture into the cylindinduced fuel-air charge, the inlet valve is kept open during the firthis way the cylinder and atmospheric pressures have almost equ

11. If the exhaust valve were to close as the piston reachedwould inevitably be some waste gas remaining in the cylinder, agavalve open a little beyond TDC, during the start of the induction that the last of the waste gas is exhausted.

12. Typical timing of the opening and closing of the inlet astroke cycle is illustrated in the timing diagram at Figure 11-3.



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FIGUValve T



RE 11-3iming Cycle



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urs shortly before the end of theum pressure early on in the power

ce is illustrated at Figure 11-3. Then stroke when both valves are open

rankshaft during which the linears very small. These times are termedDC and BDC, and are shown at

FIGUIneffecAngle

gle periods does a lot to achieve thehe valves and valve gear.



13. It will be seen from Figure 11-3 that the spark occcompression stroke, so that the burning mixture achieves maximstroke. The sequence of valve lead, valve lag and ignition advanperiod at the end of the exhaust stroke/beginning of the inductiois termed the valve overlap.

14. There are two periods during each revolution of the cmovement of the piston per degree of rotation of the crankshaft ithe ineffective crank angles. They occur at either side of TFigure 11-4.

RE 11-4tive Crank

15. Opening and closing valves during the ineffective crank anaims listed above, whilst preventing undue stresses occurring to t



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the rotational speed of the enginelve lead, lag, and overlap. With such spark, in other words to cause themaximum combustion pressure stilled will affect the amount of ignitionak mixtures.

rpm is achieved automatically.

at sufficiently high speed to requiretant over the operating speed range ignition of the mixture takes place,.e. nearer to TDC) with increasing constant, but the time taken for thesing rpm.

operates to the two-stroke cycle, inkes of the piston instead of four. A



16. The various factors that affect valve timing increase asincreases. Consequently high-speed engines have considerable vaengines, as engine speed increases it is necessary to advance thespark to occur at a greater angle before TDC. This ensures that occurs early in the power stroke. The strength of the mixture usadvance that is required, since rich mixtures burn faster than we

17. This advance of the ignition timing with increasing engine

18. Modern, air-cooled aircraft piston engines do not rotate automatic ignition advance and the ignition timing remains consof the engine. It is worth noting however that the point at whichafter the spark has occurred, will be progressively retarded (iengine rpm. This is because the time taken for the fuel to ignite iscrankshaft to rotate a given number of degrees is less with increa

The Two-Stroke Cycle19. A very few small single-engine aircraft use an engine thatwhich the complete cycle of operations is completed in two strosimple diagram of a two-stroke engine is shown at Figure 11-5.



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FIGUTwo-S

in the side of the cylinder admit thee movement of the piston uncoversw of the gases into and out of the

he two-stroke cycle with the pistonl-air mixture is being compressed ine piston is creating reduced pressurere difference, between the outsideinlet manifold to open and admit ampression-induction stroke.



RE 11-5troke Cycle

20. There is no inlet or exhaust valve. Inlet and exhaust portsfuel air mixture and allow the expulsion of exhaust gases as ththem. The crown of the piston is specially shaped to aid the flocylinder.

21. Referring to Figure 11-5, let us start the description of tabout halfway up the cylinder, as in Figure 11-5a. A charge of fuethe upper part of the cylinder whilst the upward movement of thin the lower part of the cylinder and crankcase. The pressuatmosphere and the crankcase causes a non-return valve in the fresh charge of fuel-air mixture into the crankcase. This is the co



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he sparking plug ignites the mixtureown, driving the crankshaft throughcompresses the new charge in the intake non-return valve to close, as

the exhaust port, allowing the wasteuncovers the inlet port. The slightinder through the inlet port and theylinder. The incoming charge assistsscavenging. This is illustrated at

gine, the two-stroke spark ignitionre unsuitable for use as an aircraft such as in powered gliders and

f a force moves and is measured byof the force.

tion of the force)

work.



22. As the piston approaches the TDC position a spark from tand the heat of combustion expands the gas to force the piston dthe connecting rod. The downward movement of the piston crankcase, raising its pressure above atmospheric and causing theshown in Figure 11-5b.

23. Continuing its downward travel the piston first uncovers combustion gases to begin escaping to atmosphere, and then overpressure in the crankcase forces the new charge into the cylshape of the piston crown directs its flow towards the top of the cwith the final expulsion of the exhaust gases, known as Figure 11-5c.

24. Although simpler in construction than the four-stroke enengine is much less efficient and is difficult to cool. It is therefoengine except where the power required is relatively small,microlights.

Power and Efficiencies

Power 25. Work is said to be done when the point of application othe product of the force and the distance moved in the direction

Work = force x distance moved (in the direc

Power is defined as the rate of doing



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eting his steam engines he used theomparison of power. He estimated

of a rate of doing work of 550 footrarily used this rate to define one20 HP it means that it is capable oflb/min. One horsepower is equal to

nergy of a burning fuel/air mixturetly mechanically converted to rotary

fied in terms of the theoretical or terms of the effective power, after

work done within the engine to



Horse Power. When James Watt was developing and markhorse, the most familiar ‘engine’ of the day, as a reference for c(perhaps rather generously) that the average horse was capable pounds per second (ft lb/s) or 33,000 ft lb/min. and he arbithorsepower (1HP). Thus, when an engine is said to be rated at doing work at the rate of (20 x 33,000 ft. lb/min) 660,000 ft. 745.7 watts

26. A piston engine produces power by converting the heat einto the reciprocating motion of a piston in a cylinder, subsequenmotion by the connecting rod and crank.

27. The power produced at the crankshaft may be quantiindicated power available from the burning mixture (IHP), or infrictional losses, (BHP).

28. Indicated Horsepower (IHP) takes no account of anyovercome friction. It is calculated using the formula:

PowerWork DoneTime Taken-----------------------------=

IHPPLANK33 000,

---------------------=



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measured in the cylinders by a spe-

inute.

e and is reduced by the frictionr actually delivered to the propeller

iency of the engine and is expressed

rted into thrust, after the various horsepower.

t great, when compared with a gasvantage.



Where:

• P = The mean effective pressure (lb/in²) ascial pressure gauge.

• L = The length of the piston stroke in feet.

• A = The area of the pistons (in²).

• N = The number of working strokes per m

• K = The number of cylinders.

29. The indicated horsepower is a purely theoretical valuhorsepower to give brake horsepower (BHP). This is the powegearing.

30. The ratio of BHP to IHP is known as the mechanical efficas a percentage. A figure of about 80% is normal.

31. To complete the power train, the power that is convepropeller losses have been taken into account, is known as thrust

32. Whilst the mechanical efficiency of a piston engine is noturbine, it is its poor thermal efficiency which is its greatest disad



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ng the conversion of the heat energyproximately 30% of the heat energyt to engine cooling (25%), the heatrcome friction and ‘pumping’ losses

t energy converted into useful workke horsepower (BHP) or indicated

= 778 ft lb)

ted" is substituted on both sides of

rmal efficiency at full power wouldmay improve slightly. By using highail at a later stage) it is possible tohough, for mechanical reasons, thise petrol engine has a better thermal

l-air charge at atmospheric pressure

33000at value (BTU) X 778

------------------------------------------------------



Engine EfficienciesThermal Efficiency is a measure of the heat losses that occur duriof the fuel to useful mechanical work. In a typical engine only apof the fuel is converted into useful work. The remainder is loscarried away in the exhaust (40%) and mechanical work to ove(5%).

33. The thermal efficiency of an engine is the ratio of the heato the heat energy of the fuel. It may be based upon either brahorsepower (IHP). It is given by the formula:

34. (778 is a conversion factor, 1 BTU (British thermal Unit)

35. For Indicated Thermal Efficiency (ITE) the word "indicathe equation.

36. For many modern aircraft piston engines an excellent thebe about 34%. At slightly reduced power the thermal efficiency compression ratios and high octane fuels (both discussed in detachieve thermal efficiencies in a petrol engine as high as 40% altis not attempted in many practical aero engines. Even at 34%, thefficiency than other types of piston engine.

Volumetric Efficiency is the ratio of the volume of the induced fueto the piston displacement.

BrakeThermal Efficiency BHP x wt/min fuel burned x he--------------------------------------------------------=



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the fuel-air charge into the cylindereric pressure) in the cylinder. Hence. Remembering the gas laws it willressure, its volume would be less.

volume at atmospheric pressure) ofiency of an engine.

factors: (1) Partial throttle openinguction systems of restricted cross-, and inversely with cross-sectional

ion of the airflow. (4) High inlet airrature, also causing low density. (6)s not completely exhausted throughg charge. (7) Incorrect valve timing.t of inflow to the cylinder and the(escape) of exhaust gases. The inleton stroke of the piston, the exhaustceases. (8) Engine rpm. Volumetricd friction in the intake, carburettor

mospheric pressureacement-------------------------------------------------



37. An un-supercharged (normally aspirated) engine induces by creating a depression (a pressure lower than ambient atmosphthe induced charge will be at a lower-than-atmospheric pressurefollow that, if this mass of induced charge were at atmospheric p

38. Factors which tend to decrease the mass (and therefore the induced charge have an adverse effect on the volumetric effic

39. Volumetric efficiency will be affected by the following causing restricted mixture flow to the cylinders. (2) Long indsectional area. Friction increases directly with length of manifoldarea. (3) Sharp bends in the induction system, causing decelerattemperature, causing low density. (5) High cylinder head tempeIncomplete scavenging. If the spent charge within the cylinder ithe exhaust valve, there is less available volume for the incominThe degree of opening of the inlet valve determines the amoundegree of opening of the exhaust valve determines the outflow valve must be opened as wide as possible throughout the inductivalve must close immediately exhaust flow from the cylinder efficiency may be limited at high engine rpm because of increaseand valve ports.

Volumetric Efficiency volume of charge at atpiston displ

----------------------------------------------------------=



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xhaust gases from the engine, this isent pressure means that the effective will escape more easily and less willhe induction stroke. Consequently,

the volume of the cylinder with ther with the piston at top dead centre

of total volume to clearance volume. diagrammatically at Figure 11-6.

FIGUCylind

hen it is at BDC.

the piston when it is at TDC.

the clearance volume, and is equale. Remember that the stroke is the



40. The ambient atmospheric pressure opposes the escape of eknown as back pressure. As altitude is increased the fall in ambiexhaust back pressure decreases. This means that the exhaust gasremain in the cylinder after the exhaust valve closes during tvolumetric efficiency improves with altitude.

Compression Ratio. The compression ratio is the ratio of piston at bottom dead centre (BDC), to the volume of the cylinde(TDC). Putting it another way, the compression ratio is the ratio Total volume, clearance volume and swept volume are explained

RE 11-6er Volumes

The total volume of the cylinder is the volume above the piston w

41. The clearance volume of the cylinder is the volume above

42. The swept volume is the difference between the total andto the cross-sectional area of the piston multiplied by the strokcrank throw multiplied by two.



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during its induction stroke. It is the

n the mass of the fuel-air mixtures and the definition of density (massdensity of the induced charge must

d pressure. If altitude is increasedwever, the reduction in air pressure decreases with altitude, the mass ofeloped decreases.

eased, its density is lowered. Use ofdensity (and therefore mass) of the

midity. The more humid the air theion charge.

above, is of the order of 3.5% pers graphically the change of power



43. Piston displacement is the volume displaced by the pistonsame as the swept volume.

44. The power developed by a piston engine depends upoinduced during the induction stroke. Bearing in mind the gas lawper unit volume) it will be seen that any factor that affects the affect the mass of the charge and therefore the power developed.

45. Factors that affect density of a gas are temperature antemperature decreases, which tends to increased air density. Hohas a much more marked effect in reducing density. Since densitythe induced charge decreases and consequently engine power dev

46. Similarly, if the temperature of the induced charge is incrheated intake air (to prevent carburettor icing) will reduce the induced charge, resulting in a power reduction.

47. The density of the fuel/air mixture will also vary with hulower its density and therefore the lower the weight of the induct

48. The power reduction with increasing altitude, referred to1000 feet of altitude increase. The diagram at Figure 11-7 showwith altitude with a normally aspirated petrol engine.



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FIGUPowerwith A(NormAspira

on of the cylinders relative to theial and Horizontally Opposed.



RE 11-7 Reduction ltitude ally-ted)

Engine Design Types49. Aircraft piston engines are typified by the configuraticrankshaft. The principal configurations are In-Line, V-type, Rad



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nged in a single line from front toeriod between the World Wars andbest known is the De Havilland

linders were mounted beneath thell frontal area of the engine enabledere usually air-cooled, the air beingers by baffles.

ich led to the major disadvantage ofpiston engine. Hence the power-to-ed four or six cylinders. Figure 11-8



In-Line Engines. In this configuration the cylinders are arrarear. The arrangement was popular for racing aircraft in the pcontinued in training aircraft for many years. Probably the Chipmunk. In this, as in a number of other examples the cycrankshaft to give what is known as an inverted engine. The smait to be housed in a narrow, low drag cowling. In-line engines wscooped in on one side of the engine and directed over the cylind

50. The cylinder arrangement required a long crankshaft, whthe type since the crankshaft is the heaviest component of any weight ratio of the in-line engine was poor. In general the type usshows a partly sectioned side view of an inverted in-line engine.



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FIGUPart-Seof InveEngine

was to mount two lines of cylinders,haft. The small frontal area of the12 versions, that is six cylinders inthe Second World War, probably thee type was liquid-cooled, since air totally inadequate. The power-to-t deal of effort was expended tryingth bank of cylinders on a commonned end view of a V-type engine.



RE 11-8ction View rted In-line

V-Type Engine. A natural development of the in-line engine set apart by 45° or 60° and operating onto a common cranksengine made it popular for use in high performance aircraft. V-each bank, were developed and improved just before and during best known being the Rolls Royce Merlin of Spitfire fame. Thcooling of such a large and powerful engine would have beenweight ratio was better than the straight in-line engine, and a greato improve this even further by adding a third and even a fourcrankshaft, but with no great success. Figure 11-9 shows a sectio



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FIGUSectioView o



RE 11-9ned End-f V-Engine



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rame and engine designers sought tof the crankshaft weight in the total

ld War One aircraft went as far as itas attached to the cylinder block insame plane. The piston connectingfrom the axis of the propeller. Thereciprocated in their cylinders. Thee pilot control problems at take-off. was sound however, and led to theful type of aircraft piston engine

st all large military and civil aircraftcylinders extend radially from the a common flat vertical plane. Theg rod to the single crank throw. Allough articulated connecting rods.r radial engine.



Radial Engines. From the earliest powered aircraft both airfimprove the power-to-weight ratio by reducing the proportion oweight of the engine. The rotary radial engine used in many Woris possible to go by having no crankshaft at all. The propeller wwhich, typically, seven cylinders were mounted radially in the rods were attached to a common point on the airframe, offset entire engine and propeller rotated and as it did so the pistons torque of this mass thrashing around can be imagined, causing thAlso lubrication presented unique difficulties. The basic principledevelopment of the static radial engine and the most powerproduced.

51. From the 1920’s until superseded by the gas turbine, almoused radial piston engines. In this type an odd number of centreline of the crankshaft, the cylinders being spaced evenly inpiston in one of the cylinders is connected by a master connectinthe remaining pistons are connected to the master rod thrFigure 11-10 shows a partly sectioned end view of a nine-cylinde



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FIGUPartly End ViCylindEngine



RE 11-10Sectioned ew of Nine er Radial



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ers, although three and five cylinderchieve even distribution of power toit takes two complete revolutions ofimproved power-to-weight ratio iscommon crankshaft and 14 and 18tart of the Second World War. Thet was a four-row, 28-cylinder radial

ircraft piston engine in present dayt or O-Type engine. For fixed wing

tal on either side of the crankshaft,wer-to-weight ratio is high becauset power in-line engine. Horizontallyht cylinders, although four and six

of the short crankshaft and becauseas a small frontal area keeping drag of a four-cylinder opposed-piston



52. Radial engines most commonly used seven or nine cylindversions were built. An odd number of cylinders is necessary to athe crankshaft. The engine operates on the four-stroke cycle, so the crankshaft for all cylinders to fire. Greater power and achieved by adding a second row of cylinders, operating on a cylinder two-row radial engines had become common by the smost powerful aircraft piston engine used in production aircrafengine, the Pratt and Whitney R-4360.

Horizontally Opposed Engines. By far the most popular ause is the horizontally opposed type, otherwise known as the Flaaircraft the engine is arranged with the cylinders lying horizonwhich is parallel to the longitudinal axis of the aircraft. The pothe crankshaft is kept short, about half the length of an equivalenopposed engines have been produced with two, four, six and eigare the most common configuration. The design is light, becausethe short engine length lends itself to air cooling. Furthermore it hto a minimum. Figure 11-11 shows a partly sectioned plan viewengine.



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FIGUPartly View oCylindEngine



RE 11-11Sectioned f Four-er Opposed



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iston engine are the crankcase, thetons and connecting rods and the

our cylinder opposed engine. It is anreline of the engine and joined byr the crankshaft bearings and valve

he passage of the piston connectingolted. The crankcase casting also

ached to the airframe in the engine



Mechanical Components53. The principal mechanical components of the aircraft pcylinders and valves, the valve operating mechanism, the piscrankshaft and its bearings.

The Crankcase. Figure 11-12 illustrates the crankcase of a faluminium alloy casting in two parts, divided along the centthreaded studs and nuts. Cast into the casing are the housings focamshaft bearings. At the sides of the crankcase are holes for trods, with machined facings to which the cylinders will be bincludes the attachment points by which the engine will be attnacelle of the aircraft.



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FIGUCrankFour-COppos

g enough to withstand the pressureust be able to dissipate the heat oflinder barrel and the cylinder head.



RE 11-12case of ylinder ed Engine

The Cylinder. The cylinder of a piston engine must be stroncaused by the expansion of the hot combustion gases and it mcombustion. The two major components of the cylinder are the cy



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and down, is usually made fromo provide the good bearing surfacesly polished surface internally, and

ist with heat dissipation. The uppernder head.

he fuel-air mixture. It contains thelugs (or bosses) for the valve rockergs are inserted. Cylinder heads areg on the outside to dissipate the

is threaded to receive the cylinderless than the outside diameter of therinking it, and the cylinder head is the temperatures equalise the headllows for the differential expansion

iew at Figure 11-13.



54. The cylinder barrel, within which the piston moves upchrome-molybdenum steel or chrome-nickel-molybdenum steel tand high structural strength required. It is machined to a highexternally fins are machined to increase the surface area and assend of the barrel is threaded externally for attachment to the cyli

55. The cylinder head forms the combustion chamber for tmachined guides for the inlet and exhaust valves, the mounting mechanism and the threaded holes into which the sparking pluusually cast in aluminium alloy and include extensive finninconsiderable heat of combustion.

56. The lower part of the combustion chamber in the headbarrel and is deliberately machined so that its inside diameter is barrel. To fit the barrel into the cylinder the barrel is chilled, shheated to expand it. The two are then threaded together and asshrinks to form a pressure-tight fit on the barrel. This process arates of the two materials when the engine is operating.

57. A typical opposed engine cylinder is shown in sectional v



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FIGUCross-Cylind



RE 11-13Section of a er



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which forms a pressure-tight seal fluid. The purpose of the inlet andnd exhaust ports as required during

rpose is the poppet valve, so-calleded with the assistance of the spring. guide, and a circular flat head set at machined to a smooth surface thatight seal. Valves are normally madelled. The sodium becomes liquid at

the valve head, through the stem,ant. Such a valve is illustrated at



Inlet and Exhaust Valves. A valve is a movable mechanismwhen closed, but which can be opened to allow the passage of aexhaust valves in a piston engine is to open and close the inlet athe operating cycle of the engine.

58. The type of valve almost universally employed for this pubecause it pops open against the action of a spring and pops closThe valve comprises a parallel-sided stem, which slides in a valveright angles to the axis of the stem. The periphery of the head iswill mate with the valve seat in the cylinder head to form a gas-tof special steel alloys, and the exhaust valve may be sodium fiengine working temperatures and assists in conducting heat fromto the valve guide where it is dissipated by the engine coolFigure 11-14.



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FIGUSodiumValve

nt as the exhaust valves since theith solid stems and flat or concavelve guides and closing springs in a



RE 11-14-Filled

59. The inlet valves do not need cooling to the same exteincoming mixture assists in cooling. They are usually made wheads. Figure 11-15 shows a typical arrangement of valves, vacylinder head.



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FIGUValvesSpringCylind



RE 11-15, Guides and s in a er



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and the inlet valve seat from bronzeouble valve springs are used, that iss from bouncing as they close due to different gauge wire and different

m is shown at Figure 11-16.



60. The exhaust valve seat is typically made from steel alloy and they are either shrunk or threaded into the cylinder head. Dtwo helical-coil springs of different diameter, to prevent the valvethe natural vibration frequency of a spring. The springs are ofpitch to dampen out the natural frequency of each other.

Valve Operating Mechanisms. A valve operating mechanis



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FIGUValve OSystem



RE 11-16perating



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ng with a ratio of 1 to 2, so that theart of the camshaft, transmits linearaft rotates. The linear movement ofver, which is pivoted about a rocker rocker lever pushes the valve open the cam rotates further and linear

he cylinder and the push rod as the length of the push rod is provided.ngine cold, a small clearance existsarance, which will be just (and onlyclearance and is often referred to asush rod, or sometimes a part of the

ee Figure 11-17). This type of tappetnd non-return valve, and a push rod picked up by a groove around thetroke. This oil lubricates the tappet in the plunger wall. It then passese socket mechanism.



The camshaft is driven by the crankshaft through reduction gearicamshaft rotates at half crankshaft speed. The cam, forged as pmotion through the cam follower to the push rod as the camshthe push rod is transferred to the valve by means of the rocker leshaft attached to the engine cylinder head. The movement of theagainst a strong spring. This spring serves to close the valve asmovement of the push rod reverses.

61. In order to accommodate the differential expansion of tengine heats up during operation, an adjustment to the effectiveThis is carefully set so that, when the valve is shut, with the ebetween the end of the valve stem and the rocker lever. This clejust) absorbed by differential expansion rates, is known as valve tappet clearance. The tappet is an alternative name given to the ppush rod.

62. Most opposed engines are fitted with hydraulic tappets (sconsists basically of a body and plunger, with an internal spring asocket. During operation, pressure oil supplied to the tappet isbody at the point when the tappet is near the outer end of its sbearing surface and enters the plunger reservoir through a portthrough the push rod socket and hollow push rod to lubricate th



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FIGUHydra

hen the tappet is resting on the cams to eliminate this clearance. At the the body reservoir. As the cam lobe a hydraulic lock is formed, thereby from the mechanism. Valve closureis much less than that of the valve



RE 11-17ulic Tappet

63. If clearance is present in the valve operating mechanism wdwell, the spring in the tappet body pushes the plunger outwardsame time the non-return valve will open to allow oil to pass intocommences to push on the tappet, the non-return valve closes andtransmitting motion to the push rod. Thus clearance is eliminatedwill be unaffected, since the force applied by the tappet spring springs.



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gas-tight plug within the cylinder, Pistons are tapered slightly towardsatures involved. The crown of theessary valve clearances, and to add

e of the piston to maintain a gas sealiron, which retains its springiness athe face which contacts the cylinder,ar to the bore of the cylinder during

l ring is fitted below the lowestcylinder walls. This is to ensure thatut not an excess of oil that would

pistons also have an oil scraper ringh oil might escape to the crankcase.hich forms a sliding fit in machinediston is illustrated at Figure 11-18.



The Piston and Connecting Rod. The piston forms a solidand is made of a forged aluminium alloy that is light but strong.the crown, to prevent distortion at the high operating temperpiston may be shaped to suit the cylinder head, giving the necswirl to the inducted air/fuel mixture.

64. Piston compression rings are fitted into grooves in the sidbetween piston and cylinder. The rings are usually made of cast high temperatures. The rings when new have a slight taper of tgiving a rapid initial wear, and therefore allowing the ring to wethe first few hours of operation.

65. In addition to the compression rings, an oil controcompression ring to regulate the thickness of the oil film on the there is sufficient oil for lubrication of the compression rings, bcause carbonisation problems in the combustion chamber. Some fitted around the piston skirt for oil retention; otherwise too mucThe piston is attached to the connecting rod by a gudgeon pin, wholes in the piston walls and is retained by means of circlips. A p



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FIGUTypica

f alloy steel. The H section gives aending under the considerable loads at Figure 11-19.



RE 11-18l Piston

66. The connecting rod is normally made in H section out ogreater inherent strength to the rod, and prevents the rod from bimposed upon it. A connecting rod and piston assembly is shown



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FIGUConnePiston

throws, or crank arms. In an in-line of cylinders. The large end bearingcrank throws. At each end of the

ing in the engine crankcase. Multi-ple, have further bearing journals



RE 11-19cting Rod & Assembly

67. The crankshaft is a solid steel forging having a number ofor opposed engine the number of throws is equal to the numberof the connecting rod fits around the crank pin between the crankshaft a journal is machined to fit into a supporting bearthrow crankshafts, as in a six-cylinder in-line engine for exambetween each pair of crank arms.



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row radial engines. In order tod to the crank arms, or webs as theythrow, or 180° crankshaft. In a six-other words a 60° crankshaft. These

FIGUTypes Crank



68. Single-throw, or 360° crankshafts are used for singledynamically balance the crankshaft, counterweights must be fitteare sometimes called. Two row radial engines require a double-cylinder in-line engine the cranks will be at 60° to each other, in various types of crankshaft are illustrated at Figure 11-20.

RE 11-20of shaft



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ally mounted at the rear end of the. In some cases the propeller end of°) crankshaft with counterweights,

FIGUCrankCountTiming

d in aircraft piston engines, plain



69. The valve timing gear, which drives the camshaft, is usucrankshaft whilst the propeller flange is fitted to the opposite endthe crankshaft is splined to take the propeller. A four-throw (90timing gear and propeller flange is illustrated at Figure 11-21.

RE 11-21shaft with erweights & Gear

Bearings. There are three types of bearing commonly usebearings, roller bearings and ball bearings.



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losely around a shaft journal and an 11-19. This type of bearing is bestg requires forced oil lubrication to

suitable for high radial and thrustontained in a cage. In radial loadedvel within an outer race, or guide, that is shrunk onto the crankshaft.

roller bearing, with balls being usedpe of bearing. In many applicationssealed for life.



70. Plain bearings consist of two semi-circular shells that fit cexample may be seen in the large end bearing shown at Figuresuited to radial loads in low-power engines. The plain bearinprevent friction wear and overheating.

71. Roller bearings are very low friction bearings and are (axial) loading. They consist of a ring of hardened steel rollers cbearings, such as a crankshaft support bearing, the rollers trawhich is attached to the crankcase and roll around an inner race

72. The construction of a ball bearing is similar to that of a instead of rollers. Ball bearings have the lowest friction of any tyball and roller bearings can be manufactured pre-lubricated and





Piston Engine Lubrication and Cooling

Lubricants

Lubrication Methods

Pressure Lubrication Systems



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n and

friction between moving surfaces,ontact with each other would soont this, a thin film of lubricant istion is dramatically reduced.

l lubricants, the source of which isd fluids. Solid lubricants have littlee lubricating oil. Examples are micae used to lubricate ball and rollertion lubrication systems as used in

ypes of internal combustion enginesm, they absorb and dissipate heat



12Piston Engine LubricatioCooling

Lubricants1. The two principal functions of a lubricant are to reduceand to act as a coolant. Obviously two dry moving surfaces in cwear, and would generate large amounts of heat. To preveninterposed between the surfaces, separating them, so that the fric

2. The lubricants used in aircraft piston engines are minerapetroleum. Lubricants may be classified as solids, semisolids anuse in aircraft engines, except occasionally as additives to enginand graphite. Semisolid lubricants, principally greases, may bbearings but are unsuitable for circulating or continuous-operaaircraft engines. Grease is a mixture of oil and soap.

3. Fluid lubricants (oils) are the principal lubricants in all tbecause they can be easily pumped around a circulating systereadily and they provide a good ‘cushioning’ effect in bearings.



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icating oil are its flash point and its

ust be heated in order to give offct with a very small flame. Aircraftatures (compared to liquid-cooledused shall not vaporise too readilyns, the engine will not be properly

s case the oil. In simple terms it is ascosity the more freely the oil flows.the oil must have sufficient viscosityble of flowing freely through theincreasing temperature, although anhole range of working temperatures running.

by a series of numbers allocated byetermined from the time taken for a

h an orifice of fixed size. Thus, theher words, SAE 10 grade oil flows



Properties of Lubricating Oil4. The most important properties of an aircraft engine lubrviscosity.

5. Flash point is defined as the temperature to which oil msufficient vapour to burn momentarily when brought into contapiston engines are air-cooled and so operate at high temperengines). Consequently it is important that the lubricating oil since, if the vaporised oil burns then, among other consideratiolubricated.

6. Viscosity is defined as the fluid friction of a liquid, in thimeasure of the ease with which the oil will flow. The lower the viIn order to properly lubricate all the working parts of an engine to give cushioning and "slipperiness", whilst still being capalubricating oil system. The viscosity of oil tends to decrease with ideal lubricating oil would maintain constant viscosity over the wof the engine, from cold, winter starting to hot, high temperature

Oil Grades7. Lubricating oils are graded, according to their viscosity, the Society of Automotive Engineers (SAE). The SAE number is dgiven quantity of oil, at a specific temperature, to flow throughigher the SAE number, the higher the viscosity of the oil. In otmore freely than SAE 50 oil.



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ire less viscous oil than one that is toselected for the former it may well warmer climates. Similarly, if SAE

uld undoubtedly prove too viscouse especially problematical on start-

the engine.

and seasons these single-grade oilss are used which provide adequatesingle-grade (or straight-grade) oil.le-grade oils of SAE 10 and SAE 30



8. An engine that is to be operated in cold climates will reqube operated in the tropics. However, if SAE 10 grade oil were prove too thin (insufficient viscosity) if the aircraft were taken to50 grade oil were selected for the warmer climates the oil wo(thick) if the aircraft were taken to a cold climate. This would bup, when the thick oil would not circulate readily to all parts of

9. Because aircraft frequently operate in a range of climatesare unsuitable. Consequently, multi-grade or multi-viscosity oillubrication over a much wider range of temperatures than a Figure 12-1 shows, graphically, a comparison between two singand multi-grade SAE 10W30 oil.



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FIGUOil Vis

) of oil is usually given. If not, it iseight of a substance compared withy distilled water). It is necessary toulated given its volume. One litre of 10lbs. Hence, 10 litres of oil at SGhs 9 lbs.



RE 12-1cosities

Specific Gravity10. In addition to its viscosity grade, the specific gravity (SGusual to assume a specific gravity of 0.9. Specific gravity is the wthe weight of an equal volume of a standard substance (usuallknow the specific gravity of a substance if its weight is to be calcdistilled water weights one kilogram, one Imperial gallon weighs0.9 weighs 9 kilograms; one Imperial gallon of the same oil weig



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mber of functions:

terposing a film of oil.

ation. Heat is transferred from hotbricating oil.

all space between piston rings andmbustion gases from combustion

e (mainly products of combustion) oil filter.

f the engine from oxidising agents

t loads occur. Large and small endples.

typical aircraft piston engine eithercombination of the two is used.



Functions of Lubricating Oil11. The lubricating oil in an aircraft piston engine serves a nu

(a) Reduction of friction between moving parts, by in

(b) Cooling various parts of the engine by heat dissipmetal engine parts (by convection) to the cooler lu

(c) Sealing the combustion chamber by filling the smcylinder walls, thus preventing the flow of cochamber to crankcase.

(d) Cleaning the engine by carrying sludge and residufrom the moving parts and depositing them in the

(e) Preventing corrosion by protecting metal parts o(oxygen & water).

(f) Providing a cushion between parts where impacbearings and crankshaft bearings are typical exam

Lubrication Methods12. The lubricating oil is distributed to the moving parts of aby a pressurised system or by splash lubrication. In most cases a



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chanical pump supplies oil underom the engine sump and the pump

haft bearings. A pressure-relief valveor main bearings oil passes throughrod bearings. Oil from the manifoldrings and cams. The overhead valvelied through hollow push rods, the

ns are often lubricated by splashnd bearings into the crankcase and

into two categories, the dry sumphen it is running is that which isich all lubricating oil is contained

a tank rather than the engine sump.cavenge pump, which passes the oilsupply of oil, for engine lubrication,re controlled by the pressure relief



13. In a pressure lubrication system an engine-driven mepressure to the bearings. Oil is supplied to the pressure pump frforces oil into an oil manifold, which distributes oil to the cranksusually controls the oil delivery pressure. From the crankshaft, holes drilled in the crankshaft to the lower (big end) connecting also supplies the hollow camshaft to lubricate the camshaft beamechanism (rocker arm bearings and valve guides) is often suppoil having first been used to pressurise the hydraulic tappets.

14. The engine cylinder walls and the piston gudgeon pilubrication, the source of which is oil spraying from the large ethrown onto the pistons by the flailing cranks.

Pressure Lubrication Systems15. Aircraft piston engine pressure lubrication systems fall system in which the only oil contained within the engine wlubricating the working parts, and the wet sump system in whwithin the engine.

The Dry Sump System16. In a dry sump engine the oil not in circulation is stored inThe sump is kept clear of oil, when the engine is running, by a sthrough a cooler and into the tank. The pressure pump draws its from the tank and delivers it to the engine bearings at a pressuvalve.



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the oil pump will not be starved ofof a dry sump system is shown at

FIGUDry PuSystem



17. The dry sump system is ideal for aerobatic aircraft, sincelubricant whilst the aircraft is inverted. A schematic diagram Figure 12-2.

RE 12-2mp Oil



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let of the engine driven pump. Thisafts, bearings and other parts to beaintained at all times, the pump is

red oil pressure is then achieved byy-passed back to the suction side of

at forms when the oil forces its way Pistons are cooled internally by oilnents, for example camshafts, arepressure supply line via a pressure-

e tank, normally via a cooler, by aat of the oil delivery pump, and so

elow;

sed.

of the engine sump.

mage or overheating, since the oil isey through the engine.

d so over-oiling cannot occur.

w discussed.



18. Oil flows from the tank, through a suction filter to the inpump forces oil under pressure through a pressure filter to the shlubricated. In order to ensure that an adequate supply of oil is mdesigned to deliver more oil than the engine requires. The requimeans of a pressure relief valve, which allows excess oil to be bthe pump.

19. Cylinder walls can be lubricated either by the oil mist thout from between the bearings, or by jets directed to the walls.jets directed under the crowns. Other lightly loaded compolubricated by low-pressure oil, which is routed from the high-reducing valve.

20. The hot oil drains into the sump and is returned to thscavenge pump. The capacity of the scavenge pump exceeds ththere is no danger of oil accumulating in the engine sump.

21. The advantages of the dry sump system are summarised b

(a) The system is suitable for inverted flight, as discus

(b) The volume of oil carried is not limited by the size

(c) Higher engine rpm may be maintained without dacleaned (filtered) and cooled following each journ

(d) Oil is not permitted to accumulate in the sump, an

(e) The component parts of a dry sump system are no



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h prevent a surge of oil during the engine by the firewall, and is

id the engine pump. An air space isermitted in the tank. This air spaceg due to aeration. It also allows for

rs and the increased volume of oillt from oil that has drained into the

nown as a hot well, or hot pot. Therculates through the engine on start- viscosity is rapidly achieved.

ers the oil outlet to the engine. Onosing a ring of small diameter portsosity cold oil in the main tank.t well. The circulated oil is returnedil is returned, the hot well acts as aank oil until a constant temperature

ing of feed ports to the hot well aree an oil reserve for feathering in theline, or excessive oil consumption in



(f) The oil tank normally contains baffles, whicmanoeuvres. The tank is normally separated frommounted as high as possible so that gravity can aalways present above the maximum level of oil pallows for thermal expansion of the oil and frothindisplacement of oil from variable pitch propelledelivered to the tank on start up, which may resusump whilst the engine was shut down.

(g) Most dry sump oil tanks incorporate a chamber khot well ensures that only part of the oil supply ciup, so that optimum oil temperature and therefore

(h) The hot well consists of a metal cylinder that covstart up the level of oil in the hot well drops, expthat offer a high resistance to the high viscConsequently very little cold oil passes into the hoto the hot well and is re-circulated. As the hot oheat exchanger and progressively heats the main tis achieved throughout the tank.

(i) When feathering propellers are fitted, the lower rpositioned above the bottom of the tank to providevent of main tank drainage due to a ruptured oil the engine.



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d over a de-aerator plate to the hotThe air gap in the oil tank is ventedphere. A typical dry sump oil tank is

FIGUDry SuTank



22. The oil that is scavenged from the engine sump is routewell. This plate separates the air from the oil to reduce frothing. to the sump, and the sump is vented, via an oil breather to atmosillustrated at Figure 12-3.

RE 12-3mp Oil



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r purpose being to absorb foreignem to contain at least two filters, a

and is made of fairly coarse mesh toery pump itself.

les of foreign matter before the oil is, or perhaps felt, reinforced with ailter is deeply corrugated to increase

e filter and out to the engine by way exceeds a pre-set value the pressure the oil pump. Clogging of the filterhe by-pass valve. This will cause the the engine. This arrangement is



Oil Filters. Filters are incorporated in all oil systems, theimatter in the oil such as dirt and carbon. It is normal for a systsuction filter and a pressure filter.

23. The suction filter is positioned upstream of the oil pump remove large particles of foreign matter. It thus protects the deliv

24. The pressure filter is designed to remove very small particpassed to the engine bearings. It is made of very fine mesh wirewire gauze cover that prevents the element from collapsing. The fthe filtration area, and is positioned over a filter spring.

25. Oil from the oil pump passes through the fine mesh of thof a check (non-return) valve. If the supply pressure to the enginerelief valve opens, releasing excess pressure to the suction side ofincreases the pressure differential across it, and therefore across tby-pass valve to open, maintaining lubricating oil supply toillustrated at Figure 12-4.



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FIGUOil Filt

ring use. The first, oxidation, is duebustion. The second is the chemicalls after engine shut down. No filter destroy the lubricating qualities of

enewed regularly, as specified by the



RE 12-4er

26. Two important chemical changes take place in the oil duto contamination from corrosive lead salts produced during comeffect of water vapour condensing inside the engine as the oil coosystem will reverse these chemical changes, which will eventuallythe oil. It is therefore essential that the engine oil is drained and rengine manufacturer.



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ngine and are driven by the engine both the delivery (pressure) pump geared type of oil pump (spur gear

FIGUGearePump



Pressure Oil Pump. Oil pumps are mounted on or in the eitself, hence the name engine driven pumps (EDP’s). Frequentlyand the scavenge pump are driven by a common drive shaft. Apump) is shown at Figure 12-5.

RE 12-5d Type Oil



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which is driven by the drive shaft, around either side of the casing ine pump will depend on the rpm of viscosity of the oil. The capacity ofe required minimum oil pressure isximum permitted value at high rpm as shown at Figure 12-2.

e scavenge pump will be larger thanger capacity, in order to avoid an

several other types of valves are

amage to the oil cooler on start up.ingly high viscosity oil which hasd forced, in a surge of high pressure, oil pressure exceeds a safe value it

lso located between scavenge pump than pressure, and again routes the

shown at Figure 12-2. It preventso the engine sump after engine shut



A geared oil pump consists of two deep-toothed gears, one of which are encased in a close fitting pump housing. Oil is carriedthe space between the teeth. The pressure of oil delivered by thengine and pump, on the diameter of the pump outlet and on thethe pump must ensure that, even at low rpm and with hot oil, thmaintained. In order to prevent the oil pressure exceeding the maand low oil temperature a pressure regulating relief valve is fitted

27. Scavenge Pumps are normally also of the geared type. Thits associated delivery pump, since it is required to have a laraccumulation of oil in the sump.

28. Apart from the pressure relief valve already mentionedcommonly found in a dry sump system.

29. An anti-surge valve is fitted in order to prevent possible dThis would occur when starting from cold with correspondaccumulated in the sump being cleared by the scavenge pump anthrough the oil cooler. When the anti-surge valve senses that thesimply bypasses the oil cooler until the pressure has stabilised.

30. A thermostatic valve will serve a similar function and is aand oil cooler. The thermostatic valve senses temperature ratheroil through or around the cooler as required.

31. A check valve is fitted between oil tank and pump, asseepage of hot oil through the stationary delivery pump and intdown.



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car radiator. A matrix is used with volume of unit. The oil cooler unitpending on the level of protectionf the cooler.

lled by the pilot and which governystem it is necessary to be aware ofll result from rapid cooling of the oil the cooler matrix and to cause arise in oil temperature, with not. Despite this high oil temperaturee congealed oil.

is controlled automatically by a Figure 12-6.



Oil Cooler. The oil cooler works on the same principle as athe intention of creating the maximum surface area for minimummay or may not have its own pressure relief by-pass valve, deoffered by anti-surge and/or thermostatic relief valves upstream o

32. Some oil coolers have radiator shutters which are controthe amount of air passing over the cooler matrix. With such a sthe possibility of coring occurring due to mishandling. Coring wiin the cooler and causes the oil so cooled to congeal withinblockage. The cooler will then be by-passed and a rapid corresponding increase in cylinder head temperatures, will resulthe correct action is now to close the shutters, in order to heat th

33. In most cases the flow of oil through the cooler thermostatically operated temperature control valve as shown at



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FIGUThermCooleTempeContro

and oil is directed around the core rises the temperature control valveooler where it transfers heat to the



RE 12-6ostatic Oil r rature l

34. When the oil is cold the temperature control valve is openof the cooler, so that it rapidly warms up. As the oil temperatureprogressively closes, directing more oil through the core of the ccooling air passing over the core matrix.



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p. The problems of feeding oil toaft rotation takes it through the oile piston walls. It is still necessary to

Figure 12-7.



The Wet Sump System35. In a wet sump system the oil is stored in the engine summany of the engine components are overcome, since the crankshreservoir (the sump), and this causes the oil to splash lubricate thuse an oil pump to force oil into other parts of the engine.

36. A wet sump system for an opposed engine is illustrated at



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FIGUWet SSystem



RE 12-7ump Oil



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rincipal disadvantages. The first ofhin film of oil is required to preventch thicker film of oil is deposited inil.

the cylinders from the sump is hot

of the engine is filtered, and cooled

d in it would be turned into useful 30% of the heat energy released by40% passes out with exhaust gases,ipate the heat that is passed into theilure, will occur.

controlled, but also local hot spots does not occur. It is relatively easyooling medium. Pistons, valves and only be cooled indirectly. Pistonsquently they are made of a highlyure, and so pose no great problem.ry high temperatures. Sodium filled



37. The wet sump system of lubrication suffers from two pthese is that over-oiling tends to occur at high rpm. Only a very tmetal to metal contact. With splash lubrication at high rpm a muthe cylinder walls, and energy is required to remove the surplus o

38. The second disadvantage is that the oil splashed up intoand unfiltered.

39. Pressure oil supply to the remainder of the working partsby passage through an oil cooler.

Engine Cooling40. If an engine were perfectly efficient all the heat producework. This is clearly not the case, and in fact only approximatelycombustion is converted into mechanical energy. Approximately whilst the remainder heats the engine itself. It is necessary to dissengine block, otherwise mechanical damage, and perhaps total fa

41. Not only must the mean temperatures of the engine be must be avoided, so that distortion due to differential expansionto deal with the cylinders, which are in direct contact with the cspark plugs require special attention however, since these candissipate their heat by conduction to the cylinder walls, conseconductive metal. Inlet valves are cooled by the incoming mixtExhaust valves are exposed to the burnt gases and operate at veexhaust valves help to dissipate the heat as previously described.



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em. This gives reasonably efficientich is in any event prone to leakage

pation is assisted by increasing theeep fins. In order to avoid uneven

ther than along it. In the latter eventr of the engine. With in-line engines

rected across the cylinders by meansowling above the engine and bafflesthrough the cylinder cooling fins to

re opened manually by the pilot innd, to encourage maximum coolingnt to ensure adequate cooling flowr-cooling.

is to avoid excessive cylinder heads engine damage. Monitoring of theead temperature (CHT) gauge.

red by a thermocouple instrument, to operate a circular-scale indicator.source of heat. They rely upon whatr metal are joined at both ends (as temperature difference between ther the emf.



42. The majority of aircraft engines use an air cooling systcooling, and avoids a large and heavy liquid cooling system, whand other malfunctions.

43. To improve the efficiency of the air cooling, heat dissiexternal surface area of the cylinder heads and barrels using dcooling of the cylinders the airflow must be across the engine, rathe cooling effect would be progressively reduced towards the reathe cooling airflow is introduced to one side of the engine and diof baffle plates. In opposed layouts the air is introduced into the cbetween the cylinders ensure that the airflow passes downward the lower part of the cowling.

44. Cowl flaps are often fitted in the lower cowling. These alow airflow situations, such as when the aircraft is on the grouflow over the cylinders. In flight the ram air pressure is sufficiewithout the aid of cowl flaps, and they are closed to prevent ove

45. One of the principal reasons for cooling the cylinders temperatures, which are a prime cause of detonation and serioueffectiveness of air-cooling is therefore by means of the cylinder h

46. Piston engine cylinder head temperature is usually measuwhich produces an electrical output proportional to temperatureThermocouples convert heat energy into electrical energy at the is known, as the Seebeck effect where, if two wires of dissimilashown, in Figure 12-8) an emf will be produced when there is atwo junctions. The greater the temperature difference, the greate



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FIGUThermPrincip

nnected to a point on the cylinder, which measures current flow. Theand therefore the greater the current



RE 12-8ocouple le

47. The hot junction represents the thermocouple probe, cohead. The cold junction is formed at the temperature indicatorgreater the temperature difference, the greater the emf produced flow to deflect the CHT gauge pointer.





Piston Engine Ignition and Starting Systems

Starters for Piston Engines



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d Starting

ce a series of sparks of sufficienttial that the spark should occur atropriate sequence.

sion stroke a high voltage electricalted in the upper part of the cylinder.he sparking plug electrodes to ignitecycle. This process must be repeated

essure environment of the cylinderlly in excess of 20,000 volts. This isrmer.

point in the cycle of operations a the sparking plugs at the required a four-cylinder opposed engine thisgine manufacturer. In a six-cylinder



13Piston Engine Ignition anSystems

1. The purpose of the engine ignition system is to produintensity and duration to ignite the fuel/air mixture. It is essenprecisely the right time and that the cylinders are fired in the app

2. As the piston approaches top dead centre on the comprescurrent is applied to the central electrode of a sparking plug situaThe high voltage causes a spark to jump across the gap between tthe fuel-air mixture at precisely the right time in the four-stroke in each cylinder at the same point in the cycle.

3. To provide a spark of sufficient intensity in the high-prduring the compression stroke a high voltage is necessary, typicaachieved by a system of electro-magnetic induction and a transfo

4. To ensure that the spark occurs at precisely the right mechanically operated device distributes the electrical supply toinstant and in the correct firing order of the engine cylinders. Infiring order is either 1-3-2-4 or 1-4-2-3, depending upon the enopposed engine it is 1-4-5-2-3-6.



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ery-operated coil ignition system, as

FIGUBatterSystem



Battery Ignition System5. Automobiles and a few aircraft piston engines use a battillustrated at Figure 13-1.

RE 13-1y Ignition



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flow through the ignition switch toalled a contact breaker. The cam isrent flow through the low-resistances it and the secondary coil. Thes than the primary, typically a ratio

s, thus breaking the primary circuituses the magnetic field surrounding

h emf in the closely-wound turns ofgh the primary coil also induces ats. Given the ratio of the primary/nduced in the secondary coil will be

ibutor, which is mechanically timedlectrical connection to the sparking the compression stroke.

it at the instant the contact breakere contact breaker points. To absorb parallel with the contact breaker.



6. The battery supplies a low voltage, high current electricala primary coil and thence to earth via a cam-operated switch cdriven through gearing from the engine crankshaft. The high curprimary coil creates a strong magnetic field, which surroundsecondary coil is wound from thin wire and has many more turnof at least 100 to 1.

7. As the cam rotates it separates the contact breaker pointand instantaneously terminating the primary current flow. This cathe primary coil to rapidly collapse, inducing as it does so a higthe secondary coil. The same rapid collapse of the field throumomentary high voltage there also, typically of about 240 volsecondary transformer it follows that the instantaneous voltage iof the order of 24,000 volts.

8. The secondary coil is connected to an engine-driven distrto ensure that, at the instant of induced high voltage, there is eplug in the cylinder where the piston is just approaching TDC on

9. The relatively high voltage induced in the primary circupoints separate would be liable to cause undesirable arcing at ththis voltage and prevent arcing a capacitive condenser is fitted in



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ston engine ignition systems, exceptator called a magneto, rather than arimary and secondary transformer.

licated. That is to say the completeses connecting the magneto to itslf of the system will work entirelyndent of the other aircraft electrical

vantages, but also improves engine the compressed fuel/air mixture at

horough combustion of the mixture,

own at Figure 13-2.



Magneto Ignition System10. This principle of operation is common to most aircraft pithat the source of primary current is an engine-driven a.c. generbattery. As well as a generator the magneto contains the pwindings, the contact breaker and distributor, and the condenser

11. The ignition system of an aero engine is completely dupsystem comprises two magnetos, two ignition wiring harnessparking plugs, and two sparking plugs per cylinder. Each haindependently of the other and both will function entirely indepesystems.

12. Duplication of the ignition system has obvious safety adefficiency. This is achieved because the two sparking plugs ignitetwo separate points in the cylinder, giving more even, rapid and traising the power developed in the engine.

13. The basic construction of a rotating magnet magneto is sh



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FIGUBasic RMagne

ncillary drive and rotates within ther is the primary coil, formed fromhe rotor magnets cut through theetic induction. If the primary coil isf will cause a relatively high current

ery in the ignition system previouslyat it requires no external source ofctrical system failure.



RE 13-2otating

t Magneto

14. The four-pole permanent magnet is engine-driven via an ajaws of the soft iron ‘U’ shaped stator. Wound upon the statorelatively thick, low-resistance copper wire. As the fields of tprimary coil windings an emf is induced in them by electro-magnnow connected to a suitably low-resistance circuit the induced emto flow through the circuit.

15. Thus, the rotating magnet magneto has replaced the battdescribed. The advantage of the magneto for aircraft use is thpower and is therefore reliable even in the event of battery or ele



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ings (not shown at Figure 13-2). Asire than the primary windings, ande function of this transformer is toimary circuit into high voltage/low

The current induced in the primarye secondary coil. With the ignition earth for the current in the primary closed.



16. Surrounding the primary windings are the secondary windbefore, the secondary windings contain many more turns of wtherefore the combination of the two acts as a transformer. Thconvert the low voltage/high amperage current flow in the pramperage current in the secondary circuit.

17. A magneto ignition system is illustrated at Figure 13-3. coil creates a magnetic field that saturates the conductors of thswitch ON (that is to say open, see Figure 13-3) the only path tocircuit is through the points of the contact breaker when they are



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FIGUMagneSystem



RE 13-3to Ignition



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pened by an engine-driven rotatings open the primary circuit is brokene primary magnetic field collapses,

ndary winding induces a very highsupplied, by way of the distributor

ss the primary winding, where itr) at the contact breaker points. Toer) is connected across the contactit voltage, preventing the potential arcing. The capacitor discharges itsd. As this condenser ages, it is likely it may be unable to absorb the highll become pitted or eroded.

stem, or rather half of it, since thessed. Note at Figure 13-3 that the primary winding with the ignitionit on shut down, the magneto willurned during ground handling. If a engine will now quite happily startn the propeller.



18. The contact breaker is a mechanically operated switch, ocam and closed by spring force. When the contact breaker pointand primary current immediately falls to zero. Consequently, thvirtually instantaneously, through the secondary winding.

19. This rapid movement of a magnetic flux across the secovoltage in the secondary circuit. It is this high voltage which is rotor/stator contacts, to the appropriate spark plug.

20. The collapsing primary magnetic field also passes acroinduces a significant voltage, sufficient to cause arcing (flashoveprevent this undesirable arcing, a primary capacitor (condensbreaker. The capacitor stores the temporary high primary circudifference across the points from becoming high enough to causestored voltage once the contact breaker points have fully separatethat its capacitance will drop below its rated value. Consequentlyvoltage, arcing will occur, and the working faces of the points wi

21. Figure 13-3 shows a simplified piston engine ignition sysystem shown would be totally duplicated as previously discumeans of switching off a magneto is to earth the live side of theswitch. Should the ignition switch not earth the primary circuoblige with a perfectly healthy spark should the propeller be tcombustible charge has been left (or drawn into) the cylinder, therunning, with disastrous consequences for whoever was leaning o



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cam-driven contact breaker and the arm are both driven through spur

per contact point is supplied withn contact with the lower point by apleting the primary circuit to earthinst the rotating cam by a spring so

point arm and separate the points,ated by an oil-soaked felt strip, to from the mounting bracket by and wear-resistant material, such ast the contact breaker points open ath the primary coil windings, is at a



22. Contained within the same casing as the magneto is the distributor. The contact breaker cam and the distributor rotorgearing from the magneto rotor shaft.

23. A contact breaker is illustrated at Figure 13-4. The upprimary current through a conducting flat metal arm. It is held ileaf spring when the cam is in the dwell position as shown, comthrough the mounting bracket. A cam follower is held lightly agathat the cam lobes will force it up against the upper contact breaking the primary circuit. The cam and follower are lubricminimise cam wear. The upper contact point arm is separatedinsulator. The contact breaker points are made of heat- anplatinum-iridium alloy. The cam drive gearing is arranged so thathe time when the magnetic flux from the rotor magnets, througmaximum.



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FIGUConta

from the magneto rotor shaft at a around the stator, or casing, of therking plug. At the end of the rotore secondary coil of the magnetot the gap between the two is smallo sparking plug. A distributor for a



RE 13-4ct Breaker

24. The distributor consists of a rotating arm, gear-driven speed that is half engine crankshaft speed. Spaced equidistantlydistributor are a number of electrical contacts, one for each spaarm is an electrical contact, supplied with voltage from thtransformer. As the end of the rotor arm passes a stator contacenough to allow high-voltage current flow from secondary coil tfour-cylinder engine is illustrated at Figure 13-5.



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FIGUDistribFour-CEngine

butor is arranged to ensure that the This, and the timing of the contacte magneto. Ignition timing, that isine operating cycle, is achieved by

in order to ensure the maximumtiming is illustrated graphically at



RE 13-5uter for a ylinder

25. The gear drive between the magneto rotor and the districontacts are coincident when secondary voltage is at peak value.breaker point opening, is known as the internal timing of thensuring that the spark occurs at the right instant in the engadjustment of the geared drive between crankshaft and magneto.

26. The correct internal timing of the magneto is essentialpossible induced voltage in the secondary coil. The internal Figure 13-6.



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FIGUGraphMagne



RE 13-6ical Internal to Timing



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ux concentrated in the soft iron core decreases to zero before increasingagnetic flux field induces an emf inaximum value, the contact breakerto earth. The current flow producesduced by the rotating magnet, thushe resultant flux.

this period when the high core flux cessation of current flow, shown inux from high of one polarity to high Figure 13-6.

ondary coil induces the high voltagere 13-6.

point when the flux in the core duem, and the opening of the breaker

rised with air to prevent arcing, ors dense and this allows high voltageode more easily than in denser air.nd are usually recognisable because

ring no external electrical supply to aircraft piston engines. It suffers,he engine.



27. As the rotor magnet of the magneto turns the magnetic flof the primary coil increases to a maximum of one polarity, thenagain to a maximum, but of the opposite polarity. This moving mthe primary coil. Whilst the magnetic flux is decreasing from mpoints are closed and consequently a primary current is flowing a magnetic field that opposes the changing flux of the field prosustaining high flux strength in the primary coil core, known as t

28. It is arranged that the contact breaker points open duringstrength is sustained by the primary current flow. The immediatethe upper graph at Figure 13-6, then ensures a rapid change of flof the opposite polarity. This is illustrated in the second graph at

29. The rapid change of flux through the windings of the secneeded for the ignition spark, as shown in the third graph at Figu

30. The number of degrees of magneto rotation between the to the rotating magnet (the static flux) is at a steady maximupoints, is known as the E-gap.

31. Magnetos required to operate at high altitude are pressuflashover within the distributor. At altitude atmospheric air is leselectricity to jump across the gap between rotor arm and electrPressurised magnetos are used on many turbo-charged engines athey are painted grey or dark blue (as opposed to black).

32. The self-sufficiency of the magneto ignition system, requikeep the engine operating, is the main reason for its use withhowever, from a serious disadvantage when it comes to starting t



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crankshaft speed. At the very lowhe magneto is turning too slowly tothe primary flux field is very weak. this weak field cannot induce at the sparking plugs. With no spark necessary to introduce some device starts and magneto rotary speed is

dle and the magneto shaft. As theutch is wound up and then released rotor is briefly turning very rapidlye. Once the engine has started, the is direct from engine drive shaft to

been separated to show the shell, orconnected to the hub, or cam, which



33. The magneto of a four-cylinder engine rotates at half crankshaft rpm when the starting motor is turning the engine, tgenerate sufficient current in the primary circuit, consequently Thus, when the contact breaker points open the collapse ofsufficiently high voltage in the secondary coil to cause a spark athere is no ignition and so the engine will not start. Clearly it isthat will boost the magneto output until such time as the engineadequate.

The Impulse Coupling34. A spring-loaded clutch is located between the drive spinengine is turned slowly by the starter motor the spring-loaded clonce every half-revolution. During the release phase the magnetoand releasing a sufficiently energetic spark to ignite the mixturspring-loaded clutch is disengaged by centrifugal force and drivemagneto rotor shaft.

35. Figure 13-7 is a diagram of an impulse coupling that has body, which is driven from the engine drive shaft. This is in turn is connected to the magneto rotor, by a strong spring.



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FIGUImpuls

on the hub of the impulse coupling.ce of the starter motor. The body of held stationary by a stop pin on the two is being wound up, storing



RE 13-7e Coupling

36. Figure 13-8 shows the action of the flyweights mounted In Figure 13-8a the engine has started to rotate under the influenthe coupling is turning, but the hub and therefore the magneto iscasing of the magneto. Consequently the spring between themechanical energy.



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FIGUImpulsOpera



RE 13-8e Magneto tion



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lug attached to it contacts a triggerte about its pivot point, disengagingng hub, and magneto rotor at high at the plug. This procedure will be

ifugal action on the heavy flyweighty are clear of the stop pins and drive illustrated at Figure 13-8c.

crankshaft turns through a numberse coupling. This has the effect of retarding the ignition. This has theill immediately begin turning in theo the magneto is uninterrupted the

pulse coupling.

ator instead of an impulse couplinghe device supplies pulsating direct closed. A diagram of an induction



37. In Figure 13-8b the body has rotated to a point where aramp on one of the flyweights. This causes the flyweight to rotait from the stop pin. The wound spring now rotates the couplienough speed to generate sufficient secondary voltage for a sparkrepeated until the engine fires and accelerates.

38. Once the engine is running under its own power the centrtails will rotate the flyweights about their pivot points so that thefrom the impulse coupling body to the hub is unimpeded. This is

39. It will be appreciated that, during engine start, the engineof degrees whilst the magneto is held stationary by the impuldelaying the spark beyond its normal point in the cycle and thusdesirable effect of ensuring that when the engine does fire it wcorrect direction of rotation. Once the engine starts and drive tignition timing returns to its normal advanced position.

40. Usually only one of the two magnetos is fitted with an im

Induction Vibrator41. Some magneto ignition systems employ an induction vibrto ensure a sufficiently high secondary voltage for starting. Tcurrent to the magneto primary coil whilst the starter switch isvibrator circuit is shown at Figure 13-9.



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FIGUInductCircuit

lied to the vibrator coil VC throughs to earth. The retard breaker pointsenergised the electro-magnetic fieldrgised. The vibrator contact points

imes per second, sending a pulsating



RE 13-9ion Vibrator

42. When the starter switch is closed battery current is suppthe vibrator contact points and through the retard breaker pointare located in one of the two magnetos. As the vibrator coil is opens the vibrator contact points and the vibrator coil is de-eneclose again under spring force and the process is repeated many tcurrent through the primary winding of the magneto.



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ition system, inducing high voltagegs. The retard breaker points when

ary coil. They are timed to open ation during starting. Once the enginehe induction vibrator.

ll that screws into the cylinder head, is supplied and a ceramic insulatore one or more electrodes set a smallhe spark jumps.

l around the upper part of the plug and which would otherwise cause

-10.



43. The system thus behaves in the same way as a battery ignimpulses in the magneto secondary circuit for the sparking pluclosed short circuit the supply from vibrator coil to magneto prima point later than normal in the engine cycle, thus retarding ignithas started the starter switch will be opened, thus de-energising t

Sparking Plugs44. A sparking plug comprises three main parts, the metal shea central electrode to which the magneto secondary high voltageseparating the two. Attached to the lower end of the steel shell ardistance from the central electrode, forming the gap over which t

45. Aircraft sparking plugs usually incorporate a metal barreto provide a shield for the electro-magnetic field they produce,serious radio interference.

46. A typical aircraft sparking plug is illustrated at Figure 13



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FIGUSparki



RE 13-10ng Plug



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on system is checked by means of a system functions correctly when its function when its ignition switch ispm recommended in the operating in rpm with the cylinders firing on of the individual magneto ignition drop observed on a magneto drop

rive (direct-cranking) electric motor.s with a large diameter starter geartor is only engaged with the engine

tarter motor it disengages from the

ce battery power is applied to theircraft it may be engaged manually.lly disengaged to prevent the engine motor windings.

r.



Magneto Drop Check47. The correct functioning of the duplicated magneto ignitimagneto drop check. This is a check that each magneto ignitionignition switch is open (in the ON position) and that it ceases toclosed (in the OFF position). With the engine running at the rmanual, the magnetos are individually switched off and the dropone plug only (each) is noted. This is a check both of the healthsystems and that the ignition switches are not earthed. The rpmcheck should typically be between 50 and 125 rpm.

Starters for Piston Engines48. Aircraft piston engine starters normally employ a direct dThis drives the engine through a starter pinion gear that engagering attached to one end of the engine crankshaft. The starter moduring starting. When battery power is disconnected from the sengine.

49. Engagement of the starter drive is usually automatic onmotor (by operating the starter switch), although in some light aAs soon as the engine has started the starter drive is automaticadriving the starter motor, which would lead to overheating of the

50. Figure 13-11 shows a typical direct-cranking starter moto



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FIGUDirectStarter

engagement is the Bendix drive, splined to a flexible drive, which isnion is threaded onto helical splinesthe starter gear ring by the anti-drift



RE 13-11-Cranking Motor

51. Probably the commonest method of automatic starterillustrated at Figure 13-12. The motor armature shaft is keyed orin turn keyed to the drive shaft of the engaging gear. The drive pion the drive shaft, and is normally held out of engagement with spring.



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FIGUBendixStarter

e armature rotates the drive shaft.ly and so it moves axially along theith the starter gear ring and turn the



RE 13-12 Dive

52. When battery power is applied to the starter motor thBecause of its inertia, the drive pinion tends to rotate more slowhelical splines on the drive shaft (like a nut on a bolt) to engage wengine.



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pinion faster than the starter driveout of engagement with the starter off, by releasing the starter switch, gear ring.

stationary gear ring, each time the. This, in turn, can lead to failure ofer warning lamp remains lit on thee to the starter motor.



53. When the engine starts, the starter gear ring turns the shaft. This forces the pinion back along the helical splines and ring, assisted by the anti-drift spring. When battery power is cutthe anti-drift spring holds the pinion disengaged from the starter

54. In service the engagement of the rotating pinion with theengine is started, leads to wear and damage of the gear ring teeththe pinion to disengage on start-up. When this occurs an ambcontrol panel and the engine must be shut down to avoid damag






Introduction

The Carburettor

Slow Running Control

Carburettor Icing

Fuel Injection Systems



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ly

so that there is sufficient oxygen toximum release of heat energy. The

uired for the engine power or rpmetering of the fuel supply are the

ixture to the engine at the correcte power condition. This process is

carburettor employs a venturi tube,enturi tube is placed in the inlet aire, or throat, in the air passage as



14Piston Engine Fuel Supp

Introduction1. The fuel supplied to the engine must be mixed with air ensure efficient and complete combustion of the fuel for the maquantity (mass) of fuel supplied must also be exactly that reqdemanded by the pilot. The devices that achieve this exact mcarburettor or the fuel injection system.

The Carburettor2. The function of the carburettor is to meter a fuel/air mmixture ratio and in the quantity required for the specific enginknown as carburetion.

3. In order to achieve the correct fuel/air mixture ratio the the operation of which is based upon Bernoulli’s theorem. The vpassage to the engine and forms a convergent/divergent nozzlillustrated at Figure 14-1.



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FIGUVentur

me as that flowing out of it since thelocity of the air as it passes through. Bernoulli’s theorem states that the

ncrease in velocity must be matched

onal to the air mass flow through itportionate mass of fuel to mix withigure 14-2.



RE 14-1i Principle

4. The quantity or mass of air flowing into the tube is the saentry and exit passages are the same diameter. Therefore, the vethe venturi must increase in order to maintain constant mass flowtotal energy of a fluid in motion is constant at all points, so an iby a decrease in pressure.

5. The pressure drop at the throat of the venturi is proportiand this pressure drop is used in the carburettor to meter a prothe air at the correct fuel/air ratio. The principle is illustrated at F



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FIGUCarbuVentur

irflow at the venturi throat, wherechamber is subject to atmosphericchamber and nozzle. This will forceirflow in the venturi to mix with the to the pressure difference between

f airflow through the venturi.



RE 14-2rretor i

6. A fuel chamber is connected to a nozzle placed in the athere is a pressure reduction, or depression. The fuel in the pressure and thus there is a positive pressure difference between fuel to flow from the chamber, through the nozzle and into the aair. The quantity or mass of fuel that flows will be in proportionchamber and nozzle, which is in turn proportional to the mass o



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o the engine cylinders by the actionbed. The amount of air drawn in is cylinders, in the inlet manifold. The14-3.



7. Air is induced to flow through the carburettor venturi intof the pistons during their induction stroke, as previously descricontrolled by a throttle valve placed between the venturi and thetype of valve used is usually a butterfly valve as shown at Figure



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FIGUCarbuButter



RE 14-3rretor fly Valve



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t manifold and pivoted on a centraliphery of the disc is shaped to form to an angle of about 70° to the axis disc is rotated to a position parallel

w, because of its thinness. This is the

chamber contains a float-operatedional to fuel flow out through thetterfly valve. The fuel chamber iss placed between fuel chamber andhe nozzle will be proportional to the seen is proportional to the airflowor and is the basis of many aircraft



8. The butterfly valve is an oval metal disc placed in the inlespindle that is perpendicular to the axis of the manifold. The pera close fit with the walls of the manifold when the disc is rotatedof the manifold. This is the closed position of the valve. When theto the axis of the manifold bore it offers little restriction to airflofully open position of the valve.

9. Figure 14-4 shows a basic carburettor system. The fuelinlet valve, which controls fuel flow into the chamber proportnozzle. Airflow to the engine is controlled by a throttle buconnected to atmospheric pressure and a fixed orifice, or jet, inozzle to limit the maximum fuel flow rate. The fuel flow from tpressure difference across the jet orifice, which we have alreadythrough the venturi. This is known as a float-chamber carburettengine carburettors.



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FIGUFloat CCarbu



RE 14-4hamber

rretor



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from the nozzle due to changes inntroduced to the float chamber via a the air intake will equally affect theto intake pressure fluctuations. A

gure 14-5.

FIGUPressuDuct

itations. The main problem is the ratio under all conditions of enginerflow through the venturi. This willenturi, and will cause more fuel to



10. In order to prevent changes in the rate of fuel dischargeatmospheric pressure at the air intake, air pressure is frequently iduct from the intake itself. In this way any change in pressure atfloat chamber, and the mixture ratio should not change due carburettor fitted with a pressure balance duct is illustrated at Fi

RE 14-5re Balance

11. The system described is simple, but has some serious liminability of the system to deliver a fuel/air mixture at the requiredspeed. An increase in the engine rpm will cause an increase in aiincrease the pressure differential between float chamber and venter the airflow as described in the foregoing paragraphs.



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irectly proportional to the pressurep. Consequently as airflow throughp across the jet orifice increases ande drop. Hence the fuel/air mixturely lean at low rpm.

h the air mass flow, maintaining aome form of fuel flow adjustment isat-chamber carburettors.

of the float-operated fuel inlet valvedischarging from the fuel nozzle thelve is closed. This is known as thetial created across the jet orifice, orvel in the float chamber will fall, asdle valve. When the needle valve is out the float chamber fuel level wille was stationary.

erence across the main jet and thethe float must fall before the needleamber. From this it can be seen that

main fuel metering jet as shown at



12. However, the flow of a fluid through an orifice is not ddrop across the orifice, but to the square root of the pressure drothe venturi increases with increased engine rpm the pressure drofuel flow to the engine increases as the square of that pressurwould become unacceptably rich at high rpm and/or unacceptab

Mixture Ratio Control13. To ensure that the fuel mass flow increases linearly witconstant fuel/air ratio over the whole speed range of the engine, snecessary. One such device is the diffuser tube, fitted in many flo

14. The diffuser tube relies for its operation upon the effect in the float chamber. When the engine is stopped and no fuel is fuel level in the float chamber will rise until the inlet needle vastatic fuel level. When the engine is running the pressure differenmain jet, will ensure that fuel flows from the nozzle. The fuel lefuel is drawn off, and the float will fall with it, opening the neeopen just sufficiently so that the fuel flow in equals the fuel flowstabilise, but at a lower level in the chamber than when the engin

15. The higher the engine rpm the greater the pressure diffgreater the fuel flow from the float chamber. Hence, the further valve is open sufficiently to stabilise the fuel level in the float chthe float chamber fuel level is proportional to engine speed.

16. The diffuser tube is fitted immediately downstream of theFigure 14-6.



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FIGUDiffuse



RE 14-6r Tube



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fuser, which is basically a perforated the static fuel level, so that they areel in the diffuser drops, uncoveringf air at intake pressure to enter the

ression pressure. Thus, the pressurefuel flow through the jet slightly.

level falls, uncovering more of thesure into the tube. The increase ofcreasing rpm.

efully graduated to ensure that therough it, is matched to the airflow

nge of the engine. To put it anothere main jet as airflow through the

t proportion to the airflow increase,

atomisation of the fuel. Because thell break up into tiny droplets - aness dense than liquid fuel, it can beAlso, the emulsion will have a larger

leed into the fuel feed from main jetted at Figure 14-7.



17. Air from the pressure balance duct is bled through the diftube within the main fuel passage. The perforations occur belowcompletely covered at low rpm. As the rpm increases the fuel levthe upper perforations in the diffuser tube. This allows a bleed otube and raise the pressure to a value slightly above venturi depdifferential across the main jet is decreased slightly, reducing the

18. The higher the engine rpm the more float chamber fueldiffuser tube perforations and allowing more air at intake prespressure differential across the main jet is thereby limited with in

19. The sizes of the perforations in the diffuser tube are carpressure drop across the main jet, and therefore the fuel flow ththrough the venturi at a constant ratio over the whole speed raway, the diffuser tube adjusts the pressure difference across thventuri increases, to ensure that the fuel flow increase is in directhus maintaining a constant fuel/air ratio at all engine speeds.

20. The diffuser has a second advantage in that it aids in the diffuser, by its action, introduces air into the fuel, the fuel wiemulsion of air and fuel. Since such a mixture or emulsion is ldrawn to the lip of the discharge nozzle that much more readily. surface area than liquid fuel, so it will evaporate more readily.

21. In carburettors where a diffuser tube is not used an air bto nozzle is introduced to create such an emulsion. This is illustra



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FIGUAir BleEmulsi

th running at low rpm. Despite thextures at low rpm, the problem stillrottle butterfly valve almost totally sufficient fuel through the main jet

this situation a slow-running jet is



RE 14-7ed fier

Slow Running Control22. A slow running, or idling jet is necessary to ensure smoobest efforts of the diffuser to overcome the problem of weak miexists to some extent. This is due to the fact that, with the thclosed, the pressure drop through the venturi is too small to drawand the engine is liable to stop or run unevenly. To remedy positioned as shown at Figure 14-8.



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ttle butterfly valve almost closed, alf. Fuel from the slow-running jet isir that is bled from just upstream of

FIGUSlow R



23. The slow running system works because, with the throconsiderable venturi effect occurs around the butterfly valve itsetherefore introduced just downstream of the throttle valve. The athe butterfly valve helps to atomise the fuel.

RE 14-8unning Jet



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ge to the slow running nozzle. Thishe engine, normally by selecting therather than simply switching off thenning is avoided. This would occur to continue igniting even with the

valve is opened should produce aine. In other words, slowly openinge opening produces a rapid increasewould not be the case. Opening theease in airflow and subsequently a to begin accelerating. This wouldceleration. As can be seen, after an

rather than immediate.

mand for more power or rpm it is as well as airflow. This is achieved



24. It will be noted that a valve is positioned in the fuel passais the slow running, or idle cut-out which is used to shut down tidle cut-off position on the mixture control lever. By doing this ignition (earthing the magnetos), the possibility of engine over-ruif any carbon hot spots within the cylinders caused the mixturespark plugs inactive.

Accelerator Pump25. It is clearly desirable that the rate at which the throttlesimilar response, in terms of rate of increase in rpm, from the engthe throttle results in a slow increase in rpm, whilst rapid throttlin engine rpm. In the simple carburettor discussed thus far this throttle valve rapidly would immediately produce a small incrsmall increase in fuel flow, which would cause the engine rpmincrease the airflow, and thus the fuel flow, further increasing acinitial delay (called a flat spot) the process would be progressive

26. In order to ensure immediate response to a throttle denecessary to ensure that there is an immediate increase in fuel flowby the accelerator pump, which is illustrated at Figure 14-9.



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FIGUAccele

ersed in the float chamber of themechanism. When the throttle lever is forced down within its cylinder,positioned adjacent to the main fuelo the pilot’s demand.



RE 14-9rator Pump

27. The accelerator pump piston operates in a cylinder immcarburettor and is linked directly to the throttle valve operating is moved to open the throttle valve the accelerator pump pistondischarging fuel through a delivery tube to an accelerator nozzle nozzle, enabling the engine to accelerate in immediate response t



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ses the spring of a delayed actionerator pump piston the compresseder a further controlled discharge ofuring acceleration is thus directlyent.

essary to be able to vary the fuel/aircreases with altitude its density alsone receives less mass of air at eachore rich. Consequently, to maintainurettor must be reduced as altitude density increases, whether due toat constant pressure, in order to

to avoid excessive cylinder headhe fuel burn temperature can beixture.

n be attained at low power settings, by selecting a slightly lean mixture.



28. The downward movement of the piston also compresplunger. As the engine responds to the initial action of the accelspring expands, forcing the delayed action plunger down to delivfuel to maintain acceleration. The additional fuel supplied dproportional to both the rate and extent of throttle lever movem

Mixture Control29. Because of the changes in air density with altitude it is necmixture ratio for an aircraft piston engine. As the air pressure dedecreases and so the normally aspirated (un-supercharged) engiinduction stroke and the mixture ratio becomes progressively mthe correct mixture ratio the amount of fuel supplied by the carbincreases. Similarly, the fuel supplied must be increased as airincreasing pressure in the descent or decreasing temperature maintain the correct fuel/air ratio.

30. When operating at high power settings it is essentialtemperatures and the accompanying danger of detonation. Tconstrained by selecting a higher-than-normal (i.e. rich) fuel/air m

31. Mixture control is also required so that fuel economy cawhere high cylinder temperature and detonation is not a hazard,



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ity of fuel passing through the main by altering the pressure differentialure 14-10 is a diagram of a mixtureed by altering the pressure above the

FIGUManuaContro

pressure drop across the main jet isake pressure. This is the fully rich through the venturi. Moving thevely reduce the pressure differenceme intake pressure into the diffuserith the mixture control valve fully



32. In all cases mixture control is by adjustment of the quantjet whilst the airflow remains unchanged. This may be achievedacross the main jet or by restricting the fuel flow to the jet. Figcontrol system in which the fuel flow through the main jet is varijet and thus the pressure difference across it.

RE 14-10l Mixture l

33. If the mixture control valve in Figure 14-10 is closed thethe difference between venturi depression pressure and air intsetting where fuel flow will be maximum for a given airflowmixture control valve toward the open position will progressiacross the main jet, reducing fuel flow through it, by leaking sotube above the main jet. The fully lean condition is achieved wopen.



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valve that directly controls the sizewhich is controlled from the pilots’.

FIGUNeedleMixtur



34. An alternative type of mixture control valve uses a needleof the main jet orifice. The needle valve is actuated by a cam, mixture lever. Such an arrangement is illustrated at Figure 14-11

RE 14-11 Valve e Control



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control device that adjusts the fuel/, alleviating the pilot of the need toillustrated at Figure 14-12. A bleedd an aneroid capsule is connected to

FIGUAutomMixtur



35. Some sophisticated carburettors use an automatic mixtureair ratio automatically as ambient atmospheric pressure changesadjust mixture ratio during climb or descent. Such a system is from the venturi is connected to the carburettor float chamber ana valve situated in the float chamber vent to atmosphere.

RE 14-12atic e Control



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the atmospheric vent valve is fullyfor any given airflow through the atmospheric pressure decreases the

. Meanwhile the suction bleed fromin the chamber, which the restrictedjet, and the fuel flow through it, is

d to deliberately enrich the mixture power range. The valve, or valves, is advanced beyond a pre-set point,eliver take-off power. Because thesee, they are sometimes referred to as

directly into the venturi, adding toe throttle setting approaches the fulle rotates sufficiently at full throttleection to the economiser nozzle. As lower piston and forces that down,erential pressure across the jet thenly to the engine is supplemented, or



36. At sea level the aneroid capsule is fully compressed andopen, so pressure difference across the main jet is maximum venturi. This is the rich mixture setting. As altitude increases andaneroid capsule expands, moving the vent valve towards closedthe venturi to the float chamber is causing a pressure reduction vent cannot equalise. The pressure difference across the main reduced thus reducing (leaning) the fuel/air mixture ratio.

Economiser37. Economisers, or power enrichment systems are often fittewhilst the engine is operating towards the top of its permissibleare cam-operated and may come into operation when the throttleor when the manifold air pressure control mechanism is set to dsystems ensure that enrichment does not occur in the cruise regimeconomiser systems.

38. The enrichment system shown at Figure 14-13 feeds fuelthe fuel already being discharged from the main nozzle, when ththrottle position. A cam attached to the throttle operating linkagselection to force down the upper piston, closing off an air connthe upper piston moves down further it contacts the shaft of theuncovering a jet orifice connected to the float chamber. The diffforces fuel to the economiser nozzle. Thus the normal nozzle suppenriched, at high power settings.



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FIGUEcono

as high as +30°C and in clear air,mperature. The most critical free airoccur at low power settings, such asnd creating a significant pressure/



RE 14-13miser

Carburettor Icing39. Carburettor ice can form with outside air temperaturesgiven a relative humidity of 30% or more, depending upon air tetemperature is thought to be +13°C or so. Icing is more likely to descent power, when the throttle butterfly is almost closed atemperature drop.



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ation by means of some form ofpecially in the region of the throttleient to alter the shape of the venturiom the fuel nozzle. In this event the is melted.

air accelerates through the venturisure and cooling as the evaporating

rburettor.

rnative air source, which is routedurce of this heating air to be within

these facts will help to preclude theuring use at temperatures of 0°C ore air intake ducts and/or filters.

m with fixed pitch propeller systems propeller systems.

the density of the air intake and ambient temperatures of 0°C andf carburettor icing, by raising the



40. It is necessary to protect the carburettor from ice formcarburettor heat control. Ice will form within the carburettor, esbutterfly valve and the venturi tube, and will eventually be sufficto such an extent that the zone of low pressure is moved away frengine stops and probably will not restart until the ice formation

41. The causes of ice formation are adiabatic cooling as theand around the butterfly valve, with the consequent drop in presfuel absorbs latent heat from both the air and the body of the ca

42. Carburettor heating normally takes the form of an altethrough a muffler around the exhaust pipe. It is normal for the sothe engine compartment and for this air to be unfiltered. Both risk of impact icing affecting the alternate (heated) air supply dbelow. Impact icing is simply airframe icing forming on the engin

43. Carburettor and/or intake icing symptoms are reduced rpor reduced manifold pressure with constant speed, variable pitch

44. It should be appreciated that use of hot air will lowertherefore reduce engine power. Furthermore, use of hot air atbelow may in fact compound rather than cure the problem otemperature of the incoming air to a critical level of around +13°



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FIGUCarbuGraph



RE 14-14rretor Icing



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133/1992 (Pink 68), indicates the induction system icing for a typicale much greater risk of serious icingdew point readings the greater the

are shown at Figure 14-15.

FIGUCarbuInduct



45. The chart shown at Figure 14-14, reproduced from AICwide range of ambient conditions conducive to the formation oflight aircraft piston engine. Particular note should be taken of thwith descent power. The closer together the temperature and relative humidity.

Alternate Air46. The principal elements of a carburettor induction system

RE 14-15rretor ion System



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prevent dust, sand or large foreigne air filter may be of the dry paper made of pleated layers of porous units. Polyurethane foam filters areement. Wetted mesh filters compriseintake air must pass.

d directly from the intake scoop tod or in dusty/sandy conditions anarburettor receives an uninterrupted

is able to select an alternate, heated4-15, to shut off the main air duct,e alternate air valve is often sprung

ot to select alternate heated air. This/mass and therefore limit maximum

ng the engine exhaust pipe. Air fromernate air valve is set to HOT, flowsuct to the main air duct.

e mass of the induced charge will be. The more the intake is blocked theether.



47. An air filter is situated near the air scoop in order to objects entering the engine by way of the induction system. Thtype, polyurethane foam or wetted mesh. Dry paper filters arepaper and are similar to those found on most automotive powera relatively new innovation, using wetted foam as the filtration ela mat of wire mesh elements wetted with oil, through which the

48. In some engine induction systems unfiltered ram air is lethe carburettor. When the aircraft is operating on the grounalternate, filtered air supply is selected by the pilot. In flight the cair supply.

49. For flight in known or forecast icing conditions the pilot air supply by operating the alternate air valve, shown in Figure 1opening the air supply from the exhaust-heated air muff pipe. Thto the ‘normal’ position, thus requiring positive action by the pilis because the use of heated intake air will reduce charge densityavailable power.

50. The exhaust heater muff is an open-ended pipe surroundiinside the engine compartment enters this pipe and, when the altaround the hot exhaust before passing through the alternate air d

Air Intake Blockage51. Clearly, if the air supply to the carburettor is restricted threduced and the engine will be incapable of achieving full powerless power available until eventually the engine would stop altog



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will be to reduce the manifold airill depend upon the power plant

peller is not usually fitted with associated with the partial blockagedication that the pilot has.

a constant speed propeller must betion to the pilot of intake systempeller rpm will not change since thent rpm.

gine will almost certainly prevent it carburettor to initiate fuel flow.

dvantages, which were appreciatedain, these disadvantages affect the:

rettor is subject to gravity and iner-strength during manoeuvres and an

to icing in certain conditions. Thisuces engine power by reducing inlet



52. The effect of a partial blockage of the air intake systempressure. The indication to the pilot that this has occurred warrangement of the aircraft.

53. A normally aspirated engine driving a fixed pitch promanifold air pressure (MAP) gauge. The loss of engine power awill result in a reduction of rpm and this is likely to be the first in

54. A normally aspirated or a supercharged engine driving equipped with a MAP gauge. This is because the first indicablockage will be a reduction of manifold air pressure. Engine/propropeller governor will adjust propeller pitch to maintain consta

55. Blockage of the induction system of a stationary piston enfrom starting, since there will be insufficient airflow through the

Carburettor Disadvantages56. The venturi, or choke carburettor has a number of disaearly in the development of aircraft piston engines. In the mmaximum power available from the engine. In summary they are

• The fuel in the float chamber of the carbutia, resulting in variations of the mixture inability to function in inverted flight.

• The choke, or venturi tube, has a tendencymust be countered with heating, which redair density.



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from the carburettor jets impede thee reduce the volumetric efficiency of

een largely superseded, particularlynder pressure directly into the inletpercharger impeller, but more oftenject it into the induction system. Inchieved.

or, in which the choke is retained ass by means of a fuel injection pumpole speed/power range of the engine

turi in the inlet manifold. However, airflow. It is not used to suck fueler. The basic principle of operation



• The choke and the protruding fuel nozzles flow of air through the intake and thereforthe engine.

57. Consequently, the float chamber type of carburettor has bin high-powered engines, by systems in which fuel is injected umanifold. In some cases the fuel is directed into the eye of the suindividual nozzles positioned just upstream of each inlet valve inthis latter system an even distribution of fuel to each cylinder is a

58. Fuel injection may be by means of an injection carburettthe method of measuring airflow. More commonly however, it iin which the maintenance of the correct air/fuel ratio over the whis achieved without the use of a venturi.

Fuel Injection Systems

Injection Carburettor59. An injection carburettor retains a choke passage, or venthe depression it creates is used purely as a means of measuringfrom a float chamber through jets, in fact there is no float chambof an injection carburettor regulator is shown at Figure 14-16.



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FIGUInjectioCarbu



RE 14-16n

rretor



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c pressure and venturi depression Fuel is supplied to the valve underr nozzles positioned at the cylinder

through the venturi to increase,ferential across the diaphragm isring valve. The response to throttler nozzles in direct proportion to the

which spray fuel directly into theifold is by means of a conventional more complex than the simplifiedpplied, and the mixture strength, isemperature (air density) and power

eedle valve, which controls a portn pressure chamber. Adjustment ofnd therefore the opening of the fuel-re differential, leaning the fuel/air



60. The differential pressure between ambient atmospheripressure flexes a diaphragm to position the fuel metering valve.pressure from an engine-driven pump, and delivered to injectoinlet valves.

61. Opening the throttle butterfly valve cause the airflowdecreasing the venturi depression pressure. The pressure diftherefore increased, flexing the diaphragm to open the fuel meteopening is thus the immediate supply of more fuel to the injectoincreased airflow.

62. From the metering valve fuel flows to injector nozzles,manifold at the inlet valves. Control of airflow in the inlet manthrottle butterfly valve. These injection carburettors are muchdiagram at Figure 14-16 would suggest. The quantity of fuel suregulated to match manifold air pressure (MAP), altitude and tsettings.

63. Mixture control is manually adjusted by means of a nbetween the air inlet pressure chamber and the choke depressiothe valve adjusts the pressure differential across the diaphragm ametering valve. Opening the needle valve reduces the pressumixture, closing it has the opposite effect.



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and a carburettor system (whetherrflow is not measured by a venturi.p and the quantity is metered to suit pump that will meter fuel to match



Fuel Injection Pumps 64. The fundamental difference between a fuel injector systemfloat-chamber or fuel injection carburettor) is that the intake aiInstead, fuel is supplied to the cylinders by an engine-driven pumengine requirements. Figure 14-17 is a diagram of a fuel injectionMAP and engine speed conditions.



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FIGUFuel InPump



RE 14-17jection



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eans of a throttle butterfly valve in

mber, pressure in this chamber isf flyweights, mounted on the samee centrifugal force of rotation. In soerned fuel chamber and thence, via

rates the governed and metered fuelbers acts in opposition to the actionfuel flow to engine rpm. The thrustare of the governor rpm and theerence across the diaphragm is also across a jet is proportional to ther needle valve is matched directly to

main metering needle, the positionssure.

ressed, causing the main meteringsed MAP. The profile of the needlemaintained over the whole range of



65. Engine air supply is controlled in the usual manner, by mthe intake. The fuel injector pump unit controls fuel supply.

66. A vane type fuel pump pressurises the governor chacontrolled by the pressure relief valve. The governor consists oshaft as the engine-driven pump, which move outwards under thdoing they open a needle valve allowing fuel to flow into the govvariable jets, to the metered fuel chamber and the engine.

67. A diaphragm attached to the governor needle valve sepachambers. Thus, the differential pressure between the two chamof the centrifugal governor and this has the effect of matching exerted by a centrifugal governor is proportional to the squdiaphragm balances that thrust. Consequently the pressure diffproportional to the square of the rpm. The pressure differencesquare of the fuel flow through it, so fuel flow past the governoengine rpm.

68. The flow of fuel through the main jet is controlled by theof which compensates for changes in MAP and exhaust back pre

69. If MAP increases, the evacuated capsules will be compneedle valve to open and increase fuel flow to match the increavalve is shaped to ensure that the mixture strength is correctly manifold air pressures from idling to maximum power.



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haust back pressure decreases andfficiency increases. This means thatlinders, which tends to weaken the the venturi compensates for this byfuel injector.

ched to the evacuated MAP capsule, increases the back pressure capsulence between MAP and atmosphericedle valve, maintaining the mixture

arge, which results from an increase is used to control the position of ale chamber is connected to the

manifold air temperature causes the valve to reduce fuel flow, matching

bove, fuel injection pumps normallye pilot to lean the fuel/air ratio fore power settings. In many cases thishe size of the main metering orifice,eans of a rotary valve in the orifice.offering a second fuel path to the



70. As ambient air pressure decreases with altitude the exscavenging of the cylinders improves. Consequently volumetric emore of the air available in the manifold is drawn into the cymixture. In a float-chamber carburettor the extra airflow throughdrawing more fuel through the jets, but there is no choke with a

71. To overcome this problem the back pressure capsule, attais connected internally to atmospheric pressure. Thus, as altitudeis progressively compressed (by the progressively greater differepressure) causing it to progressively open the main metering nestrength.

72. To compensate for the reduction in density of the inlet chin charge temperature, a thermometer bulb in the inlet manifoldsecond capsule-controlled metering needle valve. The capsuthermometer bulb by a liquid-filled capillary tube. An increased liquid to expand, compressing the capsule and closing the needlethe reduced density of the charge air.

Manual Mixture Control73. In addition to the automatic mixture control described aincorporate some form of manual mixture control to enable theconomic cruising and to enrich it for fuel cooling at high enginmanual mixture adjustment takes the form of an adjustment to teither by adjustment of the main metering valve position or by mFuel enrichment may also be by means of an enrichment jet metered fuel chamber.



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ad such that the fuel discharge is A typical fuel discharge nozzle ismp, via a fuel distribution manifold,e. At the upper end of the centralpply of external air through radialix with the fuel in an outlet chamberxing within the nozzle assists in ther better combustion



Injector Nozzle74. The fuel injector nozzles are located in the cylinder hedirected into the intake port just upstream of the inlet valve.illustrated at Figure 14-18. Fuel is supplied from the injection puto a central passage in the nozzle, through a calibrated orificpassage the fuel sprays into a counterbore, which induces a suholes. Air is carried along the central passage, with the fuel, to mbefore it is sprayed into the cylinder inlet port. This air-fuel miatomisation of the fuel, which in turn assists with vaporisation fo



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FIGUInjecto



RE 14-18r Nozzle



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. If number one piston is on its

rapidly with all other indicationsction should be taken to correct it?

linders?

dications appear and what action is

rade 20?

cation system?





The firing order of a four cylinder piston engine is 1-3-4-2compression stroke which stroke is number 3 piston?

QUESTION 2.

During cruising flight the engine oil temperature starts to riseremaining normal. What is the most probable cause and what a

QUESTION 3.

What is the purpose of the finning on air cooled piston engine cy

QUESTION 4.

After engine start up how quickly should normal oil pressure inrequired if this is not achieved?

QUESTION 5.

Is an oil of SAE grade 50 more or less viscous than one of SAE g

QUESTION 6.

What is the difference between a wet sump and a dry sump lubri



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ston engine lubrication system?

a dry sump lubrication system?

exhaust of a piston engine?

e in a piston engine?

ication system and why are they of



QUESTION 7.

Where are oil temperature and pressure usually measured in a pi

QUESTION 8.

What is the purpose of the air space at the top of the oil tank in

QUESTION 9.

What is the most probable cause of blue smoke issuing from the

QUESTION 10.

What is the most probable cause of wildly fluctuating oil pressur

QUESTION 11.

What is the purpose of the scavenge pumps in a dry sump lubrlarger capacity than the pressure pumps?

QUESTION 12.

Define the term coring as applied to aero-engine oil coolers?

QUESTION 13.

Describe the form of an opposed piston engine?



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the compression stroke of a piston

engine?

es?

engines?

ine?

nce volume, swept volume and total



QUESTION 14.

What is the primary reason for compressing the mixture duringengine?

QUESTION 15.

What is the equation for the brake horsepower (BHP) of a piston

QUESTION 16.

Define the term Mechanical Efficiency as applied to piston engin

QUESTION 17.

What is the relationship between crank throw and piston stroke?

QUESTION 18.

Define the term volumetric efficiency as used in relation to piston

QUESTION 19.

What causes the inlet and exhaust valves to close in a piston eng

QUESTION 20.

Define the compression ratio of a piston engine in terms of clearavolume?



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lumetric efficiency?

ston engine cylinder vary during the

a piston engine?

on system de-activated after engine

engine?



QUESTION 21.

How and why will increasing cylinder head temperature affect vo

QUESTION 22.

In what ways do the pressure and temperature of the gas in a pilatter half of the power stroke?

QUESTION 23.

How does increasing engine RPM affect volumetric efficiency in

QUESTION 24.

Define the term valve overlap as used in a piston engine?

QUESTION 25.

How is the impulse coupling in a piston engine magneto ignitistarting?

QUESTION 26.

How is the degree of valve overlap in a piston engine expressed?

QUESTION 27.

Where would you expect to find the oil control rings in a piston



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ixture in the cylinder change during

500cc, what will the swept volume

engine?

engine produces maximum power

mon cause of it?

nd disengaged in the Bendix starter



QUESTION 28.

How will the temperature, pressure, volume and mass of the mthe compression stroke?

QUESTION 29.

An engine has compression ratio of 5:1. If the total volume is 1be?

QUESTION 30.

Define the term specific fuel consumption as applied to a piston

QUESTION 31.

What term is used to refer to the fuel:air ratio at which a pistonfor a given RPM?

QUESTION 32.

What is hydraulicing in a piston engine and what is the most com

QUESTION 33.

How is the drive from the starter motor to the engine engaged asystem?



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degrees before TDC said to be more

f valve overlap?

e inlet valves before TDC and closeclose them after TDC?

e lead and lag without significant

four stroke piston engine working

the rotation of the engine at whicher?



QUESTION 34.

In a piston engine ignition system is a spark which occurs at 20 or less advanced than one occurring at 5 degrees before TDC?

QUESTION 35.

What effect does engine RPM have upon the timing and degree o

QUESTION 36.

In the practical four stroke cycle why is it necessary to open ththem after BDC and to open the exhaust valves before BDC and

QUESTION 37.

What aspect of the motion of the piston engine permits valvbackflow of inlet and exhaust gasses?

QUESTION 38.

What is the principal difference between the two stroke and cycles?

QUESTION 39.

What effect does increasing engine RPM have on the point in maximum cylinder pressure occurs in a four stroke engine cylind



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on aero-engine and what source of

ion system in a piston aero-engine,this problem overcome?

check?

is switched off during the magneto

ic ignition system?

l mixture control in float chamber



QUESTION 40.

How is cylinder head temperature measured in a modern pistelectrical power is used?

QUESTION 41.

What source of energy is employed to power the magneto ignitwhat problem does this cause during engine starting and how is

QUESTION 42.

Why does engine RPM reduce when conducting a magneto drop

QUESTION 43.

What is indicated if the engine stops when one ignition system drop check?

QUESTION 44.

What is the purpose of the induction vibrator circuit in a magnet

QUESTION 45.

Describe two methods that may be employed to provide manuacarburettors?

QUESTION 46.

Why are some magnetos pressurised and how is this indicated?



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when the engine is running at idle

action to take is to close the cooler

contact with air flow in order to

f start-up. The engine must be shut



QUESTION 47.

Where in the carburettor should the lowest air pressure occur RPM? What problem does this cause and how is this overcome?

ANSWERS:ANSWER 1.

Number 3 piston is on its induction stoke

ANSWER 2.

The most probable cause is coring of the oil cooler. The correctcowls for a short period before reopening them slowly.

ANSWER 3.

The purpose of the finning is to increase the surface area inmaximise cooling .

ANSWER 4.

Normal pressure indications should appear within 30 seconds odown immediately if this is not so.

ANSWER 5.

An SAE grade 50 oils more viscous than an SAE 20 oil



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p system it is stored in a separate oil

er. Oil pressure is usually measureds to be lubricated.

the oil.

e most common causes are leaking

a piston engine is insufficient oil the result of the pump periodically

o the oil tank. They are of greaterme due to air bubbles within the



ANSWER 6.

In wet sump systems the oil is stored in the sump. In the dry sumtank leaving the sump substantially dry.

ANSWER 7.

Oil temperature is usually measured downstream of the oil cooldownstream of the oil pressure filter immediately before the area

ANSWER 8.

The purpose of the air space is to allow for thermal expansion of

ANSWER 9.

Blue smoke indicates that oil is being burned in the engine. Thpiston rings or valve oil seals.

ANSWER 10.

The most probable cause of wildly fluctuating oil pressure incontents. Under such circumstances the pressure fluctuations arerunning dry because of air bubbles circulating in the system.

ANSWER 11.

Scavenge pumps draw oil from the sump and transfer it back tcapacity than pressure pumps to allow for the greater voluscavenged oil.



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auses a highly viscous layer of oil toil flow through the cooler causing a

osite sides of a common crankshaft.

er

Brake Horsepower to its Indicated



ANSWER 12.

The term coring refers to a condition in which excessive cooling cform on the inner walls of the oil cooler matrix. This reduces orapid increase in oil temperature.

ANSWER 13.

An opposed piston engine is one in which the cylinders are at opp

ANSWER 14.

To increase its temperature.

ANSWER 15.

Brake Horsepower = Indicated Horsepower – Friction Horsepow

ANSWER 16.

The Mechanical Efficiency of a piston engine is the ratio of itsHorsepower.

ANSWER 17.

Crank throw is half the piston stroke.



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e mass of mixture drawn into thelume (the swept volume) of mixturen of the induced volume at ambient

olume to its clearance volume.

nce Volume

mixture temperature causing it toe mass drawn into the cylinder and

he piston.



ANSWER 18.

The volumetric efficiency of a piston engine is the ratio of thcylinder during the induction stroke, to the mass of the same voat ambient pressure. Alternatively this may be stated as the ratiopressure, to the swept volume.

ANSWER 19.

The valve springs cause the valves to close.

ANSWER 20.

The compression ratio of a piston engine is the ratio of its total v

Compression Ration = Total Volume/Clearance Volume

Or, because Total Volume = Swept Volume + Clearance Volume;

Compression ratio = (Swept Volume + Clearance Volume)/Cleara

ANSWER 21.

Increasing cylinder head temperature will increase incoming expand. This will reduce the mixture density thereby reducing thhence volumetric efficiency.

ANSWER 22.

Pressure and temperature both reduce as energy is extracted by t



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creasing friction losses in the inlet

r part of the exhaust stroke and firstnlet valves are open simultaneously.

e weights in the impulse coupling to casing. The impulse coupling then

s and the compression rings around

ll increase, its volume will decrease,ke.



ANSWER 23.

Increasing engine RPM reduce volumetric efficiency due to inmanifold.

ANSWER 24.

Valve overlap is the period of crankshaft rotation during the lattepart of the induction stroke during which both the exhaust and i

ANSWER 25.

As engine RPM increases after starting centrifugal force causes thmove such that they no longer contact the stop pins in the outeracts as a normal drive shaft.

ANSWER 26.

Valve overlap is expressed in degrees of crankshaft rotation.

ANSWER 27.

The oil control rings will be located between the oil scarper ringthe skirt of the piston.

ANSWER 28.

The temperature and pressure of the mixture in the cylinder wiand its mass will remain unchanged during the compression stro



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e clearance volume (CV). The totalhis case total volume is 1500cc.

e weight of fuel required to produce specified in 1bf/Brake Horsepower/

wer at a given RPM is known as the

re of the cylinders, prevents enginerings in inverted or radial engines.



ANSWER 29.

The compression ration (CR) = total volume (TV) divided by thvolume (SV) is the swept volume plus the clearance volume. In t

This means that CR = 1500/CV = 5, so CV = 1500/5 = 300cc

But TV = (SV + CV) = 1500cc, and CV = 300

So SV = 1500 – 300 = 1200cc

ANSWER 30.

The specific fuel consumption of a piston engine is defined as tha unit of power per unit of time. In imperial units this would beHour.

ANSWER 31.

The fuel:air ratio at which a piston engine produces the most pobest power mixture.

ANSWER 32.

Hydraulicing is the situation when liquid trapped in one or morotation. The most common cause is oil leakage past the piston



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g a helical spline to engage with the pinion moves back along the spline

ed than one that occurs at 5 degrees

verlap angle because these are fixed

low caused by the valves and thee valves opened and closed at TDCtends the period of induction andut.

motion is very small over a finitelittle or no tendency for backflow inlve lead and lag is well within the



ANSWER 33.

Initial rotation of the starter causes the drive pinion to move alonengine starter ring. When engine speed exceeds starter speed theout of engagement, assisted by a spring.

ANSWER 34.

A spark which occurs at 20 degrees before TDC is more advancbefore TDC.

ANSWER 35.

Engine RPM has no effect on the timing of the degree of valve ofunctions of engine design and construction.

ANSWER 36.

Because of the effects of gas inertia and the restriction to fmanifolds, insufficient gas would by induced and exhausted if thand BDC. Opening the valves early and closing them late exexhaust thereby improving volumetric efficiency and power outp

ANSWER 37.

Because of the phenomenon of ineffective crank angle piston angular range close to TDC and BDC. This means that there is the inlet and exhaust manifolds provided the degree of the vaeffective crank angle range.



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exhaust processes for each cylinder the four system the same processes

do not normally employ automatiche same point in rotation regardlessuired to complete combustion is

that the timing of combustion and RPM does however mean that the combustion completion and hence

normally measured using surfaceby the seebeck effect on the hot and

hanical energy provided through thee magneto drive shaft is too low to

at the spark plugs. This problem is couplings.



ANSWER 38.

In the two stroke system the induction, compression, power andare achieved in a single revolution (two strokes) of the engine. Inare carried out over two revolutions (four strokes).

ANSWER 39.

Because of their relatively low rotation rates piston aero-enginesignition advance and retard and hence ignition is commenced at tof engine RPM. Also the flame rate and hence time reqapproximately constant at all speeds. The above factors meanhence maximum pressure does not vary with RPM. Increasingpiston will have moved further down the cylinder by the timemaximum cylinder pressure occurs.

ANSWER 40.

Cylinder head temperature in a modern piston aero-engine ismounted thermocouples. The electrical power used is produced cold junctions.

ANSWER 41.

The power source in a piston aero-engine magneto system is mecengine drive shaft. During engine starting the rotation rate of thprovide sufficient energy to generate the high voltages required usually overcome by means of HT or LT booster coils or impulse



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ombustion is initiated at two pointsimately half the time that would be

eck each ignition system is tested insingle point. Because of the longerder pressures are achieved reducing

is indicates that the other system issparks to initiate combustion.

boosted low tension current to thed primary current is too low due to

r tube the pressure drop across thecontrolling the fuel flow to the main

high voltages to flashover or jumprevent this are painted blue or grey.



ANSWER 42.

During normal operation with both ignition systems operating cand conducted on two fronts. It is therefore completed in approxtaken with a single ignition point. During the magneto drop chisolation from the other, thereby causing ignition to occur at a time required for combustion on a single flame front, lower cylinpower output and RPM.

ANSWER 43.

If the engine stops when one ignition system is switched off thdefective in that it is producing no spark or insufficiently strong

ANSWER 44.

The purpose of the indication vibrator circuit is to provide a magneto primary circuit during engine starting when the inducethe low magneto rotation rate.

ANSWER 45.

By varying the air supply from the float chamber to the diffusemain fuel jet can be modified to adjust fuel mixture. By directly jet effectively altering the size of the jet.

ANSWER 46.

At high altitude the low air pressures mean that it is easier foracross gaps in circuits. Magnetos that have been pressurised to p



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he carburettor will occur where thelve. Under these circumstances the

adequate supply of fuel through themally overcome by positioning ano maintain fuel flow at low RPM.



ANSWER 47.

With the engine running at idle RPM the lowest air pressure in tair is caused to accelerate around the closed throttle butterfly vapressure drop in the choke tube will be insufficient to draw an main jets to keeps the engine running. This problem is noradditional slow running or idling jet close to the butterfly valve t





Piston Engine Power Augmentation and Performance

Introduction

Super Charging

Turbochargers or External Superchargers

Internal Superchargers



ance

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mentation

say, it converts heat energy createdmeans of the moving pistons andhe work done by these moving parts

nal to the quantity of fuel burned inhe correct ratio, which for efficienteight. This mixture of fuel and air,

arge. Consequently, the statement atthe power output of the engine is

heric pressure to enter the cylinder,tmospheric pressure. This reduced during the induction stroke. Hence,is always at a pressure lower thanottle valve wide open.

Piston Engine Power Augmentation and Perform


15Piston Engine Power Augand Performance

Introduction1. The internal combustion engine is a heat engine. That is toby burning fuel in its cylinders into mechanical energy by crankshaft. The power delivered by the engine is the function of tover a period of time.

2. Thus, the power output of the engine is directly proportiothe cylinders. The fuel will only burn if it is mixed with air in tcombustion must be 15 parts of air to every one part of fuel by wwhen supplied to the engine cylinders is known as the fuel/air chthe beginning of this paragraph can be expanded to say that directly proportional to the weight of the fuel/air charge.

Charge Induction3. In a normally aspirated engine, for air at ambient atmospthe pressure in the cylinder must be less than the ambient apressure is created in the cylinder by the movement of the pistonthe charge in the cylinder at the end of the induction stroke ambient atmospheric pressure, even when operating with the thr



ance

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ospheric pressure, then the chargedensity is defined as mass per unitess than the same volume of air at

directly proportional to its density,e, developing the conclusion of thegine is directly proportional to thee in the induction system during the the inlet air manifold and so the

ifold Air Pressure (MAP).

ut of a piston engine is directlyduction system between the throttle

twin-engine aircraft is shown atute pressure, which means that thehen the engine is stopped the gaugegure 15-1 is calibrated in inches of



4. It follows that, if the charge pressure is less than atmdensity must be less than ambient atmospheric density. Since volume, the mass (or weight) of the induced charge must be lambient atmospheric pressure.

5. Let us now examine this. The weight of the charge is which is in turn directly proportional to its pressure. Thereforsecond paragraph on this page, the power output of a piston encharge pressure. The charge pressure is, by definition, the pressurinduction stroke. The induction system is otherwise known asinduction system pressure is more commonly referred to as Man

6. From this it can be concluded that the power outpproportional to the MAP. MAP is measured at a point in the invalve and the cylinder inlet valves.

7. A typical manifold pressure gauge presentation for aFigure 15-1. Strictly speaking, MAP stands for manifold absolgauge is calibrated to read pressure above absolute zero. Thus wshows the ambient atmospheric pressure. The gauge scale in Fimercury (ins. Hg), which is usually the case.



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FIGUManifoGaugePresen



RE 15-1ld Pressure tation



ance

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ted chamber sealed by a diaphragm,pressure increases the diaphragm or

in supercharged engines. We haveower, since it has direct effect uponectly affects the cylinder operatingand detonation.

ut of the engine is indicated by thespeed (variable-pitch) propeller thees according to the propeller pitchnifold pressure, and thus the powerxcessive torque.

obstruction to the air passage to thele value but, as explained above, in

ric pressure.

it offers a restriction to airflow toion, and must create greater suctionrefore less weight of charge. Closing



8. The manifold pressure gauge typically contains an evacuaor containing a bellows unit connected to the inlet manifold. As bellows flexes proportionately to operate the gauge pointer.

9. The manifold pressure gauge is of particular importancealready established that MAP is directly proportional to engine pthe mean effective pressure in the cylinders. This, in turn, dirtemperature and high cylinder temperature leads to pre-ignition

10. With a fixed-pitch propeller installation the power outpengine rpm. However, when an engine is driving a constant engine rpm reamains constant and power increases or decreassetting. It is therefore important that the pilot is aware of the maoutput of the engine, to ensure that the engine is not producing e

Factors Affecting Induction11. When the throttle valve is wide open it offers virtually no cylinders. Under these conditions the MAP is at its highest possiba normally aspirated engine it is still less than ambient atmosphe

12. If the throttle valve is set at anything less than wide openthe cylinders. The pistons are trying to draw air past this restrictin order to do so. Increased suction equals reduced MAP and thethe throttle reduces engine power!



ance

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rnal surfaces the flow of air to thehan it could be and therefore chargeower.

ity and reduced air density meansating in conditions of high ambienttting is reduced. This includes fullase to the induction air temperatureoutput.

is decreased proportionally. When assure, as at altitude, the decreased

the higher the operating altitude the

ctors which affect the weight of theficiency of the engine. Volumetrictmospheric pressure to the volume

ced charge pressure is always at lessuced-pressure gas were at ambientvolume displaced by the piston.ating with a volumetric efficiency of



13. If the induction passages have sharp bends or rough intecylinders will be impeded. The result will be that MAP is lower tweight is less than it could be. Poor intake design results in lost p

14. The higher the air temperature, the lower the air densreduced charge weight, which means reduced power. When operair temperature the power available for any given throttle sethrottle, so the maximum power available is reduced. Any increwill decrease the charge weight and, therefore, the engine power

15. If the pressure of a volume of gas is decreased its density piston engine is operated in conditions of low atmospheric predensity of the induced air results in lower charge weight. Thus, lower the power output of the engine.

Volumetric Efficiency16. The astute student will by now have spotted that all the fainduced charge are also those which affect the volumetric efefficiency is the ratio of the volume of the induced charge at adisplaced by the piston. In a normally aspirated engine the induthan ambient atmospheric pressure. Hence, if that mass of redatmospheric pressure its volume would be less than the Consequently, a normally aspirated piston engine is always operless than 100%.



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0% is operating at full throttle andossible to increase the volumetric

en the power output of the engineP.

by normal aspiration is known asstem with some sort of air pump, orir forced in, the higher the MAP andmeans it is possible to increase theavailable at altitude as at sea level,ational ceiling of the aircraft. This islable at sea level can be increased,d-boosted engine.

in atmospheric air and discharges itensity and consequently the mass of (mass) of the air-fuel charge at ae power constant. Because of the

exhaust gases improves. A greatergine need do less pumping work in maintained in the climb, there is a altitude gained.



17. Suppose now an engine with a volumetric efficiency of 8let us also suppose that it is producing 80 BHP. If it were pefficiency, by increasing the weight of the charge, to 100% thwould increase in direct proportion (in theory at least) to 100 BH

Super Charging18. Increasing the weight of the charge beyond that possiblesuper charging. It is achieved by forcing air into the induction sycompressor, instead of sucking it in with the pistons. The more atherefore the greater the power output of the engine. By this power developed by a piston engine so that the same power is significantly improving both the climb performance and the operknown as an altitude-boosted engine. Further, the power avaigiving improved take-off performance. This is known as a groun

19. The supercharger is a blower, or compressor, that draws to the engine induction system thereby increasing the pressure, dair in the induction system. Maintaining the MAP and densityconstant value during the climb will, of itself, maintain engindecrease in ambient atmospheric pressure, scavenging of the proportion of the charge is able to enter the cylinders and the enexpelling the exhaust gas. The result is that, if constant MAP ispower increase at the rate of about one per cent per 1000 feet of



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een less than fully open on take-off, pressure decreases, the throttle ismanually or automatically), andr when the throttle butterfly valve is obstruction and suffers a reductiones climbing it will eventually reachAP constant, the throttle valve will

e for that MAP/power. Any furtherss the supercharger can be made to

m continuous power for the engine)ated altitude. Rated altitude is theed power. In a turbo-charged engineut more of that later.

ides a means of achieving a greaterhe weight and size of the power unit

charger is that the former is drivenving the fuel pump, and the latter isriven, or internal-type supercharger



20. With an altitude-boosted system, the throttle will have bin order to avoid over-boosting. In the climb, as ambient airgradually opened to maintain the required power (either consequently throttling losses are reduced. Throttling losses occuother than fully open, since the air has to accelerate around thisin pressure (and density) in consequence. As the aircraft continuan altitude at which, in order for the supercharger to maintain Mhave to be wide open. This is known as the full throttle altitudincrease in altitude must result in a fall in MAP and power unlepump more air (by increasing its speed).

21. When an aircraft is climbing at rated power (the maximuand it reaches full throttle altitude this is also known as the rgreatest altitude at which a supercharged engine can maintain ratthis same condition is usually referred to as the critical altitude, b

22. Supercharging or turbocharging of a piston engine provpower output from a given engine without an undue increase in tconcerned.

23. The basic difference between a supercharger and a turbodirectly by the engine, with an ancillary drive much like that dridriven by the exhaust gases. The basic layout of a mechanically dsystem is shown at Figure 15-2.



ance

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FIGUInternaSuperc

ternal-type supercharger is shown at



RE 15-2l harger

24. The basic layout of an exhaust driven turbocharger, or exFigure 15-3.



ance

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FIGUExternSuperc(Turbo



RE 15-3al harger charger)



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t upon the weight of the inducedt follows that the power output of ahown in curve (A) at Figure 15-3. Areater power at any given altitudef power with altitude in the samee to maintain sea level full power tog altitude as shown at curve (C). A but also of maintaining that power



25. Since the power output of a piston engine is dependencharge, and this is in turn dependent upon the density of the air, inormally aspirated engine must fall as altitude increases. This is sground-boosted, internally supercharged, engine will produce gthan a similar unsupercharged engine, but will exhibit loss omanner, as shown in curve (B). An altitude-boosted engine is abla given altitude, but thereafter power will decline with increasinturocharger is capable not only of increasing power at sea level,to a significant altitude as shown in curve (D) of Figure 15-3.



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FIGUCurveMaximvs. Alt



RE 15-4s of um Power itude



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hargers

aust gas flow, rather than by directan external-type supercharger.

essor outlet pressure to compressorunit of given dimensions, this ratio

ant in the climb, since its speed iscavenging improves as the exhaustne increases with altitude up to the



Turbochargers or External Superc

The Turbocharger26. When the impeller of the supercharger is driven by exhdrive from the engine, it is known either as a turbocharger or as

27. As with an internal-type supercharger the ratio of comprinlet pressure is the supercharger/turbocharger ratio and, for a will depend upon the speed of the compressor.

28. The output of a turbocharger remains essentially constincreased to match decreasing atmospheric pressure. Hence, sback pressure decreases, so the volumetric efficiency of the engicritical altitude (discussed shortly).

29. A turbocharger arrangement is illustrated at Figure 15-5.



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FIGUTurbocFlow

arger its speed of rotation, and theas would the back pressure in the

ion by directing the exhaust gasesdiverts the exhaust gases to the

mosphere through a valve known as



RE 15-5harger gas

30. If all of the exhaust gases were used to drive the turbochconsequent compressor output pressure, would be too high, exhaust manifold.

31. It is therefore necessary to control the speed of rotatthrough a Y junction, whereby one arm of the plumbing turbocharger turbine, whilst the other arm directs the gases to ata waste gate. This principle is illustrated at Figure 15-6.



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FIGUTurbocContro



Turbocharger ControlRE 15-6harger l



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waste gate governs the amount oftherefore governs the manifold air of waste gate movement is limited,prevent stalling.

, within which engine oil pressurelied to the actuator through a fixedriable orifice valve, operated by the

capsule connected to the variableted to the induction manifold on thesensitive to compressor discharge

intain a constant absolute pressurerottle valve. This is known as the

tween throttle valve and engine inlet

ncreasing the throttle butterfly valve inlet manifold, MAP will increase.essor discharge pressure to fall andoller to expand, partially closing the



32. From Figure 15-6 it can be seen that the position of theexhaust gas available to drive the turbocharger turbine, and pressure downstream of the turbocharger compressor. The rangeto ensure that some gas is always routed through the turbine, to

33. The waste gate position is set by a hydraulic actuatoropposes a spring to operate the waste gate linkage. Oil is supporifice, and is allowed to return to the engine sump through a vaabsolute pressure controller.

34. The absolute pressure controller consists of an aneroidorifice needle valve. The capsule is mounted in a chamber connecoutlet side of the turbocharger compressor, and is therefore pressure.

35. The function of the absolute pressure controller is to ma(i.e. pressure above absolute zero) between compressor and thcompressor discharge pressure. Manifold air pressure (MAP), bevalves, is controlled by the throttle valve setting.

36. Let us assume that the throttle lever has been advanced, iopening. As the valve opens, admitting more charge air to theHowever, the greater demand on the compressor causes comprthis will cause the aneroid capsule in the absolute pressure contrvariable orifice (oil outlet) needle valve.



ance

Chapte

ceed outflow through the variabler piston will increase. Overcoming piston to the right (in Figure 15-6),

, increasing compressor dischargel throttle advance. As compressor

e will be progressively compressed,ches oil inflow through the fixedd, with the throttle butterfly valveer to maintain constant compressoring.

tle valve or underspeeding of the.

e acting on the waste gate actuatorll open position. The main exhaustck pressure during starting.

ld air pressure will be low and thepressed. Consequently, the variablebuild up in the waste gate actuator,w rpm, the oil pressure will also be

ste gate will probably not reach thegas will be directed to the turbine,ressor.



37. Oil flow through the fixed inlet restriction will now exorifice needle valve, and oil pressure on the waste gate actuatospring pressure, the increased oil pressure will move the actuatorclosing the waste gate to direct more gas through the turbine.

38. Turbine, and therefore compressor speed will increasepressure to maintain the higher MAP required by the originadischarge pressure rises the absolute pressure controller capsulopening the variable orifice valve until oil outflow just matrestriction. At this point a new steady-state condition is reacheopen wider, waste gate further closed, turbocharger rotating fastdischarge pressure and MAP higher to match the new power sett

39. A decrease in MAP, whether due to closing the throtturbocharger, will have the reverse effect to that described above

40. When the engine is stopped there is no engine oil pressurpiston, so the spring will move the actuator to the waste gate fugas path is therefore direct to atmosphere to minimise exhaust ba

41. Once the engine is idling, or at low power, the manifoabsolute pressure controller capsule will only be partially comorifice needle valve will be almost closed and oil pressure will moving the waste gate towards the closed position. Because, at lolow the actuator spring will not be fully compressed and the wafully closed position. However, most of the available exhaust maintaining an adequate flow of induction air through the comp



ance

Chapte

re is an altitude at which the wasterequired for that power. Above thatincrease speed to match the fallingected (the throttle is fully open), the



42. For every possible power setting selected by the pilot thegate will have to be fully closed in order to maintain the MAP altitude, power will fall since the turbocharger can no longer atmospheric pressure. When a higher power setting cannot be selturbocharger's critical altitude has been reached.



ance

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FIGU



RE 15-7



ance

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ller, to ensure that MAP does not controller may be fitted, whereby

n some older engines the waste gateo progressively close the waste gately open it in the descent to prevent

ressor, which is necessary at altitude charge going to the cylinders. Thisoosting the MAP slightly) and also

rapping. This is a hunting conditionine and compressor). The assemblyhaust gas flow through the turbine,

e gate is opened to increase exhaustcreases, which causes an increase inreases exhaust flow further.......and

essure controller actuates the wasteelay (due to turbo lag) the turbine

h decreases MAP, which decreasesroblem is not usually as bad as ite fitted. A dual-unit turbocharger



Manifold Air Pressure (MAP) Control43. Turbochargers may be fitted with a fixed datum controexceed a pre-set maximum value. Alternatively a variable datumany MAP exceeding a value set by the pilot will be prevented. Omay be controlled manually, in which case the pilot is required tin the climb to maintain the boost pressure, and to progressiveover-boosting.

44. The increased speed of rotation of the turbocharger compin order to maintain MAP, will increase the temperature of thewill result in a small loss of power (which can be overcome by ban increase in cylinder head temperature.

45. Turbochargers are prone to a condition known as bootstthat arises because of the inertia of the rotating assembly (turbtakes time to accelerate or decelerate following an alteration in exand this is known as ‘turbo lag’. Suppose for example, the wastgas flow. After a delay due to turbo lag the turbocharger speed inMAP, which increases the power output of the engine, which incso on. In an attempt to overcome this situation the absolute prgate, which decreases exhaust gas flow to the turbine. After a dbegins to decelerate, which decreases compressor speed, whicpower, which decreases exhaust gas flow.......and so on. The psounds, but to overcome it additional controller units may bcontrol system is illustrated schematically at Figure 15-8.



ance

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FIGUDual USupercContro



RE 15-8nit harger l System



ance

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Figure 15-6 is designed to maintainnifold air pressure (MAP) required.tle valve will be only partly open,ent way of operating, since it meanse manifold air pressure required.

re 15-8), which is sensitive to theis adjusted to reduce turbochargerDifferential Pressure Controller iss out any tendency towards hunting

ivided by a diaphragm which has air pressure plus a spring acting onop across the valve is least and thetroller is held closed by the spring.

al pressure across it, and thereforee differential between superchargerl sufficient to overcome the springthe bleed valve. This will reduce theen and reduce turbocharger rpm to

harge pressure from exceeding itssure only when the throttle valve isitical altitude the waste gate will bend is consequently sensitive to air



46. The single-unit turbocharger control system illustrated ata constant supercharger discharge pressure, regardless of the maHence, at low power, where a low MAP is required the throtcreating a large pressure drop across the valve. This is an inefficithat the turbocharger is being driven faster than necessary for th

47. By introducing a second controller (as shown at Figudifferential pressure across the throttle valve, the waste gate discharge pressure at partial throttle openings. Because the sensitive to the pressure drop across the throttle valve it smoothof the manifold pressure, or bootstrapping.

48. A chamber in the Differential Pressure Controller is dsupercharger discharge pressure acting on one side and manifoldthe other. When the throttle valve is fully open the pressure drdiaphragm operated bleed valve in the Differential Pressure ConThe more the throttle valve is closed the greater the differentiacross the diaphragm. Hence, as the throttle valve is closed thdischarge pressure and manifold air pressure will increase untiforce. Beyond this the diaphragm will flex to progressively open waste gate actuator servo pressure, causing the waste gate to opmatch the reduced power selected.

49. The Density Controller prevents the supercharger disclimiting value. It controls the waste gate actuator servo oil presfully open and up to the turbocharger's critical altitude (above crfully closed). The closed capsule is filled with dry nitrogen atemperature changes as well as pressure changes.



ance

Chapte

alue the capsule will be compressed, pressure. The actuator spring will and thereby reducing supercharger

, slightly closing the bleed valve ande waste gate towards CLOSE andcrease turbocharger rpm, and thusharge air.

trol system in which two controllersde and a third takes over from theseat Figure 15-9. The waste gate, simplicity. Their layout is as shown



50. If supercharger discharge pressure rises above a pre-set vopening the bleed valve and reducing waste gate actuator servomove the waste gate towards OPEN, reducing turbocharger rpmdischarge pressure.

51. If the discharge temperature rises the capsule will expandincreasing the waste gate actuator servo pressure, moving thdiverting more exhaust gas to drive the turbine. This will insupercharger discharge pressure, to maintain the density of the c

52. Some ground-boosted turbochargers use a triple-unit concontrol the waste gate actuator servo pressure up to critical altituabove critical altitude. Such a control system is illustrated turbocharger and throttle valve are omitted from the diagram forat Figure 15-8.



ance

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FIGUTriple TurbocContro

er discharge pressure below criticaller at Figure 15-6.



RE 15-9Unit harger l System

Absolute Pressure Controller. This controls the superchargaltitude in exactly the same way as the absolute pressure control



ance

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ich supercharger discharge pressureh might cause interim overboosting

e bleed valve diaphragm, but supplyrcharger discharge pressure will fallssure across the diaphragm, closing

gate towards CLOSE until pressure

rger discharge pressure. However, ift pressure rises more on one side of opening the bleed valve and moving increase of supercharger discharge

qualised, the spring will return thehe Absolute Pressure Controller.

and ambient atmospheric pressurere air in order to maintain constantnal work done by the supercharger engine. This could, if continued to



Rate Controller. The rate controller controls the rate at whcan increase, in order to prevent a surge of boost pressure, whicof the engine when the throttle is opened.

53. Supercharger discharge pressure is led to both sides of thto one side is restricted. Thus, if the throttle valve is opened supeand the restricted orifice will ensure a temporary differential prethe bleed valve, increasing servo pressure and moving the wasteon either side of the diaphragm has equalised.

54. This will increase turbocharger speed to restore superchadischarge pressure rises too rapidly the restrictor will ensure thathe diaphragm than the other. The diaphragm will therefore flex,the waste gate temporarily towards OPEN to limit the rate ofpressure.

55. Once pressures on either side of the diaphragm have ebleed valve to the closed position, returning primary control to t

Pressure Ratio Controller. As aircraft altitude increases, decreases, the supercharger has to rotate faster and compress momanifold air pressure, and therefore engine power. The additiocompressor increases the temperature of the air delivered to thesufficient altitude, lead to detonation.



ance

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um manifold air pressure above aer automatically limits superchargeropening a bleed valve and reducingg turbocharger speed. This results insure.

ontaining an aneroid capsule. At theto contact the bleed valve stem. Ifng the bleed valve to reduce servogas to atmosphere and less to drivereduce the waste gate actuator servoe pressure and atmospheric pressure

ppet valve, is normally fitted to thevent of failure of the boost pressure

to serious structural damage to the

mand considerable power to drivef the 40's and 50's, it is not so with

eral aviation aircraft. Hence, theocharger.



56. In order to prevent this, a limit is placed upon maximspecified altitude (often 16,000 feet). The pressure ratio controlldischarge pressure above the specified altitude by progressively waste gate actuator servo pressure, thereby progressively reducina controlled progressive reduction of supercharger discharge pres

57. Ambient atmospheric pressure is supplied to a chamber cspecified altitude the capsule will have expanded sufficiently altitude is increased further the capsule expands further, openipressure and adjust the waste gate so as to allow more exhaust the turbocharger turbine. The Pressure Ratio Controller is set to pressure so as to maintain a ratio between supercharger dischargof, typically, 2.2: 1.

58. Finally, a pressure relief device, in the form of a simple poinlet manifold to bleed off excess pressure (overboosting) in the elimiting devices, or failure of the waste gate.

59. Overboosting would cause excess power and could lead engine.

Internal Superchargers60. Mechanically driven, or internal-type superchargers dethem. Whilst this was acceptable with the high powered engines othe relatively low powered piston engines of modern gensupercharger has been almost universally superseded by the turb



ance

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nally-driven superchargers generally to the crankshaft (typically betweener is usually capable of maintainingnd 10,000 feet, depending upon thee past, two-speed superchargers or altitude of the engine.

ft speed, at full power rpm both arer is to maintain sea level MAP atir at sea level it follows that theer, maximum engine rpm must be

closing the throttle valve.

open the throttle valve in order tonecessary for the throttle valve to bee, as previously discussed. For each

so for each power setting there is a full throttle altitude.

of high atmospheric pressure (lowair than necessary, which tends toy not only cause detonation, but isture means decreased density). For

arburettor and throttle valve, where



61. Those few engines still in use which are fitted with interemploy a single impeller compressor driven at a fixed speed ratiosix and twelve times crankshaft speed). This type of superchargsea level manifold pressure up to an altitude of between 5,000 agear ratio, at rated power. In high-powered piston engines of thdouble-impeller compressors have been used to increase the rated

62. Since the supercharger speed is geared to engine cranksharotating at maximum speed. The function of the superchargealtitude, where the air is less dense. In the relatively dense asupercharger impeller is rotating faster than necessary. Howevmaintained for full power, so MAP can only be limited by partly

63. As the aircraft climbs it will be necessary to progressivelymaintain constant MAP. When an altitude is reached where it is fully open to maintain MAP, this is known as full throttle altitudpower setting there is a different engine rpm/impeller speed anddifferent full throttle altitude. The lower the power, the lower the

64. From the foregoing it will be seen that in conditions altitude) the supercharger is doing more work in compressing overheat the induction air. Excess induction air temperature maalso partly defeating the object of supercharging (high temperathis reason superchargers are always fitted downstream of the cevaporation of the fuel in the charge can exert a cooling effect.



ance

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djusting the position of the throttlebut to relieve the pilot of the need tolimb an automatic device is oftenic Boost Controller, is illustrated at



65. Control of manifold air pressure (MAP) is achieved by avalve. The pilot may adjust the throttle valve in the normal way, progressively open the throttle valve for a constant-power cprovided for this purpose. Such a device, known as an AutomatFigure 15-10.



ance

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FIGUSupercAutomContro



RE 15-10harger atic Boost ller



ance

Chapte

n of a piston valve supplied with oile the pilot has set the throttle valve

imbs air intake pressure falls, whilstercharger discharge (manifold air)e aneroid capsule which expands,e underside of a servo piston whilstn moves upwards as a result and,e supercharger, restoring MAP.



66. An aneroid capsule, sensitive to MAP, controls the positiounder pressure (from the engine's lubricating oil system). Supposfor the required MAP at the start of the climb. As the aircraft clsupercharger rpm remains constant. This will cause the suppressure to tend to fall. The drop in MAP is sensed by thpositioning the piston valve so that oil pressure is directed to thconnecting the upper side to a return system. The servo pistothrough linkages, opens the throttle valve to admit more air to th





Piston Engine Performance

Introduction

Effect of Ambient Conditions

Engine/Propeller Relationship

Effects of Fuel-Air Ratio

Other Factors Affecting Performance

Fuel for Piston Engines



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ce

on engine are manifold air pressuresing the MAP increases the power9 ), it was shown that calculation of per minute as one of the factors the number of working strokes per power output from a piston engine

factors are inextricably linked. Ther decreased is by adjustment of thettle, which affects the manifold air

ven airspeed or rate of accelerationarticular engine/propeller rpm. Withcontrol is the throttle and the only



16Piston Engine Performan

Introduction1. The factors affecting the power output of an aircraft pist(MAP) and engine speed (rpm). It has been shown that increaoutput in direct proportion. In the section dealing with power (3.indicated horsepower includes the number of working strokesdirectly related to power. The higher the engine rpm the greaterminute and the greater the power developed. To achieve a givenrequires that a particular MAP and rpm must be selected.

2. With an engine driving a fixed pitch propeller the two only way in which the power of the engine can be increased oengine/propeller rpm. This is achieved by operation of the thropressure, of course, but this is a secondary consideration. A girequires a particular propeller thrust which, in turn, requires a psuch a power plant there is no MAP gauge, the pilot's 'power' 'power' reference is the rpm gauge.



Chapte

n by a normally aspirated or super/ther, not only control the power, butntrolled by the propeller governor,ioning the propeller lever selects atch to maintain that rpm constant.blades to a finer pitch, which offersropeller faster and engine/propeller

ove the propeller into coarser pitch,

control engine power. The more theower developed. If the pilot were to increase. The propeller rpm wouldeller to maintain constant rpm. Theeller blade angle is now incorrect for

gine/propeller rpm will increase but insufficient power for the higher



MAP versus rpm3. Where a constant speed, variable pitch propeller is driveturbo-charged engine the pilot now has two controls which, togealso the performance of the engine. Engine/propeller rpm is cowhich is set by the pilot using the rpm (propeller) lever. Positparticular rpm and the governor adjusts the propeller blade piWhen a higher rpm is selected the governor moves the propeller less resistance to rotation and so the engine is able to rotate the prpm increase. Selection of a lower rpm causes the governor to mwith opposite results.

4. The second pilot control is the throttle, which is used to throttle valve is opened the higher the MAP and the greater the popen the throttle, increasing MAP, engine power output wouldnot change however, since the governor would coarsen the propengine is working harder to maintain the same rpm and the propthe airspeed.

5. Similarly, if the propeller (rpm) lever is advanced the enengine power remains the same. The engine is now deliveringairspeed associated with higher rpm.



Chapte

n engine driving a constant speedximum propeller thrust at optimumlever are never operated in isolation,wer (in the climb for example) for. The manufacturer determines the aircraft type testing. These are then

AP for the selected rpm. This willpeller in coarse pitch. The result isad temperatures with the risk ofigh propeller load.

xpressed in terms of altitude whichsphere (ISA).

mean sea level) in the ISA to whichard conditions of temperature and

g the formula:

erature deviation)



6. In order to obtain the optimum performance from apropeller it is necessary to select a MAP and rpm which gives mapower for a given airspeed. Consequently, throttle and propeller but in conjunction with each other to produce the optimum pooptimum fuel economy, and airspeed for range and enduranceoptimum MAP and rpm settings for each situation during initialpublished in chart or tabular form for the pilot.

7. It is particularly important to avoid setting too high a Mresult in the engine delivering high power at low rpm to a prolikely to be overheating of the engine (excessive cylinder hedetonation) and over-torquing of the propeller shaft due to the h

Effect of Ambient Conditions

Pressure and Density Altitude8. Pressure altitude is defined as an atmospheric pressure ecorresponds to that pressure in the International Standard Atmo

9. Density altitude is defined as that height (above or belowthe actual density at any particular point corresponds. (In standpressure, density altitude is the same as pressure altitude.)

10. Density altitude can be converted to pressure altitude usin

pressure altitude = density altitude (120 temp×±



Chapte

used is plus)

mperature, indicating a decrease inup to about 20,000 ft), is given byvalue 1000 ft higher. In the ISA atuarter at 40,000 ft.

dent upon air density, which in turnfactors must be taken into accounttain the actual power of an enginee sea level performance in terms of achieved with the same settings atcorrected for ambient temperature

wer output (performance) increasesde or temperature) to a point where

to fall off.

also be considered. As air densitydecrease (especially in a normallyesistance to the propeller, which



(if temperature deviation is minus, the sign

11. Pressure reduces in the atmosphere more rapidly than tedensity with height at all levels. The decrease at lower levels (subtracting 3% of the value for any given level to obtain the 20,000 ft density is approximately half its sea level value, and a q

Altitude and Temperature12. At a given engine rpm and MAP its power output is depenis dependent upon pressure, temperature and humidity. These when determining the performance of an engine. In order to obpower charts are produced by the manufacturer which show thBHP for any given MAP and rpm and the altitude performancevarious density altitudes. (Density altitude is pressure altitude conditions).

13. So long as MAP can be maintained for a given rpm, powith altitude. However, when density falls (due to increased altituMAP cannot be maintained, power output (performance) begins

14. The effect of density on propeller performance must decreases engine power available to drive the propeller will aspirated engine). However, the less dense air offers less rconsequently requires less power to maintain constant rpm



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peller power required, in conditions

FIGUEngineVersusPower

temperature or higher than ISAchieved for a given throttle setting.ince it has to be turned in a denserIn conditions of decreased densityverse will apply.



Engine/Propeller Relationship15. The relationship between engine power available and proof changing air density, is illustrated at Figure 16-1.

RE 16-1 Power Propeller

16. When density is higher than Standard (lower than ISApressure) an increased power output from the engine will be aHowever, the power required by the propeller will be greater, senvironment where there is increased resistance to rotation. (higher than ISA temperature or lower than ISA pressure), the re



Chapte

er to increased/decreased propeller there will be an excess of engine

onsequently less power is needed tofor the same throttle setting.

ed at a given throttle setting will beompared with the propeller powerssary to maintain a given power.

ller rpm in the less dense air.

-air mixture ratio and it is necessarytor. These are the mixture for best

ich the engine produces maximumr example.

e least fuel consumption for a given



17. However, the ratio of increased/decreased engine powtorque requirement is not constant. When density is above ISApower available compared with the propeller power required. Cmaintain the same propeller rpm or higher rpm can be achieved

18. When density is lower than ISA the engine power producreduced and there will be insufficient engine power available crequired. Consequently, a higher throttle setting will be neceHowever, less power will be necessary to maintain a given prope

Effects of Fuel-Air Ratio19. The performance of a piston engine is affected by the fuelto consider the two cases of prime interest to the aircraft operapower and the mixture for best economy.

20. The best power mixture is the fuel-air ratio under whpower for a given rpm and is of importance during climb-out, fo

21. The best economy mixture is the fuel-air ratio that gives thpower and is of importance when flying for range.



Chapte

t which the fuel/air mixture ratio isr exhaust cooling. As an aeroplanel toward the LEAN setting. This is

ced into the cylinders decreases withs essentially the same. Consequently,ively richer until the engine ceases to

in order to both obtain smooth to determine the optimum mixture

operating manual and it will vary or fuel injection.

usual to adjust the mixture controled or until engine rough running is to fire. The mixture is then slightly

e mixture is leaned until the engineooth running.



22. At take-off power the mixture setting is FULL RICH, adesigned to permit full power with adequate additional fuel foclimbs it is necessary to progressively adjust the mixture controbecause, with increasing altitude, the mass of the air charge induthe decreasing air density, whilst the mass of fuel induced remainif no adjustment were made, the mixture would become progressrun smoothly and fuel consumption is unnecessarily high.

23. With the engine at cruise power the mixture is leanedoperation and to economise in the use of fuel. The procedure usedsetting is usually described in the pilot’s handbook or aircraftaccording to whether the engine is fitted with a float carburettor

24. With a float carburettor and a fixed-pitch propeller it istoward lean until the maximum engine rpm or airspeed is obtainencountered. This roughness is due to the leanest cylinder ceasingenriched.

25. With a float carburettor and a constant speed propeller thjust begins to run roughly and then enriched slightly to regain sm



Chapte

st gas temperature (EGT) gauge it ishe optimum mixture setting. At thent of unburned fuel, which cools thening the mixture reduces the excessch that there is no excess fuel andture is leaned beyond this point thethe excess air in the mixture. Any principle is illustrated graphically at

ure at cruise power is to adjust the is obtained. At this setting the best

t time) is achieved and the aircraftimum cruise power is required thetil the EGT is about 100°F (55°C)



26. With a fuel injected engine that is equipped with an exhauthe usual practice to use the EGT gauge reading to determine tFULL RICH setting and cruise power there is a significant amouexhaust gas and results in a low reading on the EGT gauge. Leafuel and causes the EGT to rise. When the mixture setting is suthere is complete combustion a peak EGT is reached. If the mixEGT will begin to fall because there is now cooling due to significant leaning beyond peak EGT will result in misfiring. TheFigure 16-2.

27. The procedure generally followed when setting the mixtmixture control from FULL RICH toward LEAN until peak EGTspecific fuel consumption (fuel used per BHP produced per unirange is increased by approximately 15%. Alternatively, if maxmixture is leaned to the peak EGT point, and then enriched unbelow peak value.



Chapte

FIGU



RE 16-2



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ncetherefore mass of air entering theensity and therefore the lower thettor or manifold air temperature atso be remembered that too high an

decided effect upon the volumetricpressure the lower the volumetric

ve ambient atmospheric pressure isore convoluted the passage of the

the higher the back pressure in the

mance is gained from the ram airr intake scoop. This has a slightd hence power developed, is greater

is a fuel with a high calorific value with air at ordinary atmospheric engine fuel are the calorific value of



Other Factors Affecting Performa28. Carburettor Air Temperature affects the density, and induction system. The higher the temperature the lower the dpower developed. Performance decreases with increased carburethe rate of approximately one HP for each 6°C rise. It must alinduction temperature will inevitably lead to detonation.

29. Exhaust Back-Pressure, as we have seen before, has a efficiency of a piston engine. The higher the exhaust back-efficiency and the worse the engine performance.

30. The pressure in the exhaust system (back-pressure) abolargely dependent upon the design of the system itself. The mexhaust gases, through the exhaust ports, silencers and pipes, system and the more adverse its effect upon performance.

Ram Air Pressure. Some benefit in terms of engine perforpressure, due to forward velocity of the aircraft, at the aisupercharging effect so that the engine volumetric efficiency, anthan under static conditions.

Fuel for Piston Engines31. The aeroplane piston engine burns gasoline because thisand which will evaporate readily when brought into contacttemperatures. The important characteristics of an aviation pistonthe fuel, its volatility and its anti-knock qualities.



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rgy available in it. Those fuels withnergy per unit volume.

tile fuel makes for easy starting, butan lead to the formation of vapourfore be sufficiently volatile to make will vaporise in the fuel system.

ance to detonation. The anti-knockel-air mixture can be raised before the factors which determines the

l combustion engines are gasoline,umber of disadvantages that rendercompatible with many of the seal

rapidly. The expense of using sealof gasoline/alcohol mixtures, even

nd atomic structure is identical with either from coal or oil. They areest known of the aromatic fuels is. Its disadvantages are its ability to

lusion in aviation gasoline is limited



32. The calorific value of the fuel is a measure of the heat enea high hydrogen content will produce the most calories of heat e

33. The volatility of a fuel is its readiness to evaporate. A volafuel evaporation can increase the risk of carburettor icing, and clocks in the fuel supply system. An ideal aviation fuel must thereengine starting easy, but must not evaporate at such a rate that it

34. The anti-knock value of a fuel is a measure of its resistquality of the fuel will determine the pressure to which the fuignition, without detonation occurring. It is therefore one ofmaximum power output of the engine.

Types of Fuel35. The types of fuel ideal for use in spark ignition internaalcohol and the aromatic fuels. The latter two, however, have a nthem less than ideal for use in aircraft engines. Alcohol is inmaterials used in aircraft engines, causing them to deterioratematerials resistant to alcohol attack does not justify the use though these may be cheaper than pure gasoline.

36. The aromatic fuels are so called because their molecular athat of perfumes. They are hydrocarbon compounds obtainedpresent to a limited extent in most gasoline distillations. The bBenzol, which has the advantage of being resistant to detonationdissolve rubber and its very slow burning rate. As a result its incto 5 per cent by volume.



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ne. All are rubber solvents, thought useful because of its low freezingtion gasoline up to 15 per cent by

carbon and 15% hydrogen. The airroximately 78% nitrogen and 21% carbon dioxide and water vapour,herwise be an explosion rather thanteady burn of the mixture, giving aon the flame rate, the rate at which to 80 feet per second.

explodes, then detonation is said toe, and causes physical damage to the

s normally, the temperature of theontaneously, giving a flame rate of

ulting from detonation, and so the temperature, and the formation ofn, which in due course will collapseround the valves, and can result inns.



37. Other important aromatics are toluene, xylene and cumeto a lesser degree than Benzol. Of the three, toluene is the mospoint and good volatility and it is suitable for blending in aviavolume.

Avgas38. Avgas (aviation gasoline) consists of approximately 85%with which it is mixed to form a combustible fluid consists of appoxygen. During the burn the hydrogen, carbon and oxygen formand the nitrogen, which is inert, acts as a buffer in what would ota controlled combustion. The result, hopefully, is a rapid but suniform rise of pressure on the piston. In this controlled situatithe flame spreads through the mixture, will be in the order of 60

39. If the burn is not controlled, and the mixture effectively occur. This is evidenced by a characteristic knocking of the enginunit in very short order.

40. Detonation occurs because, although combustion beginunburned part of the mixture becomes so high that it ignites sp1000 ft/sec or so.

41. The piston cannot absorb the rapid rise in pressure resenergy is wasted as heat, which causes an overall rise in enginelocal hot spots. Prolonged detonation will burn the piston crowaltogether. Detonation will also rapidly ruin the gas tight seals adistortion of the cylinder head, the connecting rods and the pisto



Chapte

essure of the charge during thewill burn instantaneously (explode).result in an increase in temperature.perature, and the ratio of fuel to airy or all of the following may cause

ted.

te use of carburettor heat, or ofcessive supercharging.

nstant speed propeller.

though detonation may follow as aed, or if the combustion chamber ise combustion chamber, such as the

head, may rise so much that it glowshe mixture before the spark occursy continue to fire with the ignition



42. Detonation is caused when the temperature and prcompression stroke reaches a level at which the fuel-air mixture Because of adiabatic heating an increase in pressure will always The density of the mixture will influence both pressure and temwill also be an important factor. In practical terms, therefore, andetonation:

(a) Too weak (lean) a mixture.

(b) The ignition too far advanced.

(c) The engine (torque) loading too high for rpm selec

(d) A high charge temperature, due to inappropriaalternate hot air in a fuel-injected system, or of ex

(e) The selection of high power at low rpm, with a co

(f) Too low a fuel octane rating.

43. Detonation should not be confused with pre-ignition, alresult of pre-ignition. If an engine is allowed to become overheatexcessively carbonised, the temperature of some projection in thsparking plug points or a piece of carbon deposit on the cylinder red hot. This is known as a ‘hot spot’ and is liable to ignite tduring the compression stroke. In this condition the engine masystem switched off, although probably not on all cylinders.



Chapte

and further overheating. The higheretonation almost always diminishesect fuel and by operation within the

d in the octane rating. The octanepure hydrocarbon spirits, normallyure at which detonation will occur. detonation, and heptane, which hasne rating is one which shows similare and 15% heptane.

sents a problem to engine designers. Clearly in an engine with a high

t be high. This is especially so if thefuel-air mixture will already be at aions exist where detonation is likely to the fuel as an anti-knock agent.a value greater than 100, at which

el/air mixture ratio it used to be thes, these being the ratings for a lean

/145 and 80/87 octane fuels.



44. With pre-ignition there is a loss of power, rough running,the engine speed the worse pre-ignition will become, whereas dwith an increase in rpm. Pre-ignition is avoided by use of the corrcorrect limits of MAP and cylinder head temperature.

Octane Rating45. The anti-knock qualities of a fuel are broadly expressesystem of classifying fuel is based on the ratio at which two found in gasoline, are mixed together to give a certain temperatThe two spirits are isooctane, which has a very high resistance toa very low resistance to detonation. Thus, a fuel with an 85 octaanti-knock characteristics to a mixture comprising 85% isooctan

46. The current trend is towards low lead fuels, and this presince it is the lead in petrol which tends to prevent detonationcompression ratio, the anti-knock qualities of the fuel used musengine is turbo-charged or supercharged, since in either case the relatively high pressure on reaching the inlet manifold. If conditto present a problem then tetra-ethyl-lead (TEL) may be addedThe addition of TEL can raise the octane rating of the fuel to point it becomes correctly referred to as a performance number.

47. Since the anti-knock qualities of a fuel depend upon the fupractice to assign two performance numbers to aviation gasolineand a rich mixture. Examples of this practice were 100/130, 115



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ide, which prevents the lead fromustion chamber.

ne grade, known as Avgas 100 LLmately 0.72, and must comply withD) specification 2485. 100 LL iss a maximum of 2 millilitres of TEL

ilable. Avgas 100 is coloured greenns more tetra-ethyl-lead (TEL) thanndly. It is the nearest equivalent fueles approximately to the old octane

o main reasons. The first of these isn all but minimal quantities. Such temperatures water in the fuel willnd obstruct the flow of fuel to the



48. The addition of TEL is accompanied by ethylene dibromoxidising onto the various harmful components within the comb

Fuel Grades49. Aviation gasoline (Avgas) is only generally available in o(low lead). It is coloured blue, has a specific gravity of approxiDirectorate of Engineering Research and Development (DERapproximately equivalent to the old grade of 100/130. It containper gallon.

50. There are two other categories of Avgas that may be avaand has similar anti-knock characteristics to 100 LL, but contai100 LL and is therefore considered to be environmentally unfrieto the old 115/145 rating. Avgas 80 is coloured red and equatrating of 80/87.

Water Content51. Water in aviation gasoline presents a serious hazard for twthat it will block the small jets in the carburettor if present iblockage will inevitably result in engine failure. Secondly, at lowprecipitate and form ice crystals which will block fuel filters aengines.



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ater pollution in aviation gasolinem in partly filled aircraft fuel tankslace. To prevent this it is important

ft parked overnight. The maximum0 parts per million.

ne, has created a problem for the to use this grade of fuel. In manyermitted limited use of automotive fuel is known as Mogas.

use in the UK implies limitations volatile than Avgas and thus more high temperatures and/or altitudes.ad to pre-ignition and detonation inbe harmful to aircraft fuel system cause deterioration of the seals iny.



52. Whilst great efforts are made to ensure freedom from wstorage and distribution facilities, condensation will readily forduring times when atmospheric temperature changes are taking pto ensure that aircraft tanks are kept filled, especially in aircrapermissible water content in aviation gasoline is approximately 3

Alternative Fuels53. The virtual non-availability of Avgas 80, or 80/87 octaowners and operators of aircraft with piston engines designedEuropean countries, including the UK, the authorities have pgasoline, provided that it conforms to certain specifications. This

54. Automobile gasoline (Mogas) may be available, but itsspecified in CAA Airworthiness Notices 98 and 98A. It is moreprone to cause vapour lock in fuel lines and pumps, especially atIt does not possess the anti-knock qualities of Avgas and may lesome aircraft piston engines. Detergents used in Mogas may components and alcohol, found in some grades of Mogas, willmost aircraft fuel systems, with the obvious hazard to flight safet



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rove the use of Mogas and using itrthiness Notice Number 98 permitsulates detailed conditions of fuel

s. In particular, in order to minimisenks must be less than 20°C and ther volatility of the fuel (and possible

al and operated, since the risk of

AA Airworthiness Notice Number

ture ratios, varying between aroundarts of air to 1 of fuel (a very lean

ombustion, and this ratio is termede carbon and hydrogen in the fuel water vapour.

temperature is too high and, apartm a process known as dissociation.ustion split momentarily into their the combustion pressure. Although the lost heat is regained, this comes



55. Most major aircraft manufacturers have refused to appmay invalidate the engine manufacturer's guarantee. CAA Airwoits use in certain light aircraft/engine combinations, but stipspecification, source of supply, storage and operating precautionthe risk of vapour lock, the temperature of fuel in the aircraft taaircraft must not be flown above 6000 feet. Because of the highewater content) carburettor hot air systems must be functioncarburettor icing is greater than with Avgas.

56. Any pilot considering using Mogas should first consult C98 and CAA General Aviation Safety Sense Leaflet Number 4.

The Mixture57. A fuel-air mixture will be combustible over a range of mix8 parts of air to 1 of fuel (a very rich mixture), to around 20 pmixture).

58. A mixture ratio of 15:1 is most likely to give complete cthe chemically correct mixture. It is at this ratio that all of thcombines with the oxygen of the air to form carbon dioxide and

59. Unfortunately, at the chemically correct ratio, the ignitionfrom the risk of detonation occurring, a loss of power results froThis dissociation occurs because the various products of combseparate elements. This process absorbs heat, and this decreasesthe various elements recombine later on in the power stroke, andtoo late in the process to be of any real value.



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rect mixture described above wererised to make detonation extremely

here is a problem in supplying eachcraft is in the cruise it is therefore unit which is 15% richer than theeceives a mixture that is on the rich

uously at high rpm, such as in theny tendency towards detonation ashe degree of enrichment may be asing on the particular engine that is

ed above, acts as a coolant but doessary to increase the mass of fuel-airre. The unburned fuel will leave the form the deadly carbon monoxide

y with mixtures that are richer thanker. This occurs because the rate off the high levels of nitrogen that aret combustion is still occurring afterful lives of the valves.



60. If a mixture that is 10% richer than the chemically corused, the excess fuel would absorb sufficient latent heat as it vapounlikely and to prevent dissociation.

61. Because of the difference in inertia between fuel and air tengine cylinder with mixture at the same ratio. Whilst the airnormal to produce a mixture at the carburettor or fuel injectionchemically correct value. This ensures that the leanest cylinder rside of the chemically correct value.

62. During the periods where the engine is working continclimb, it is normal to further enrich the mixture to counteract athe working temperatures increase under the heavy workload. Tmuch as 20% or 30% over the chemically correct value, dependemployed.

63. The excess fuel that is supplied to the cylinders, as discussnot produce extra power. In order to increase power it is necesmixture, not simply to increase the mass of fuel within that mixtuexhaust manifold as fuel vapour, or will combine with oxygen togas.

64. Strangely enough, a cooler burn can be achieved not onlthe chemically correct value, but also with mixtures that are weaburn (the flame rate) is slower with a weaker mixture, because opresent. With very weak mixtures the burn may take so long thathe end of the ignition stroke. This will obviously shorten the use



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decrease in power output, but also amixture having approximately 80%ine’s best specific fuel consumptionount of fuel used for each unit of

most economical flight conditions.

lier, the exhaust valve remains openxhaust gas velocity is low, and therento the cylinder, thereby weakeningproblem it is therefore necessary to

ing a cold engine, a quantity of neat fuel/air mixture is drawn into thetype of carburettor the fuel may be

rburettor tickler. This is a manuallyrce down the carburettor float. This

ber to rise, and eventually fuel willuction manifold.

ttle-operated accelerator pump, thelt in fuel being sprayed into the



65. The slow burn associated with weak mixture results in a decrease in fuel consumption measured in gallons per hour. If a of the strength of the chemically correct value is used, the eng(SFC) is likely to be achieved. The SFC is a measure of the amdistance flown in still air. The lowest SFC will therefore give the

66. The final problem occurs at idling rpm. As discussed earduring the initial part of the induction stroke. At idling rpm the eis a tendency for some of the exhaust gases to be sucked back ithe mixture for the subsequent cycle. In order to overcome this progressively enrich the mixture as idling rpm is approached.

Engine Priming Systems67. In order to avoid unnecessary engine cranking when startfuel may be supplied to the induction manifold, so that a richcylinders when the engine starts to rotate. Depending upon the supplied by one of the following methods.

(i) Some float carburettors are fitted with a caoperated plunger which may be used to foaction allows the fuel level in the float chamflow from the discharge nozzle into the ind

(ii) Where carburettors are fitted with a throaction of opening the throttle will resuinduction manifold.



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d, a priming pump (manually or a fuel tank and discharge it to

via a system of priming pipes andp it is usually necessary to twist therings outward, opening the supplyresult in the engine drawing excessrt-up. The engine will stall after aing).

ally accomplished by operating the period with the mixture control setr nozzles.

in is fitted to the lowest point in theay have collected will pass through

rawn from each of the aircraft tank considered unfit for aircraft use if



(iii) Where a seperate fuel primer is installeelectrically operated) will draw fuel fromvarious points in the induction manifold nozzles. To operate a manual priming pumplunger in order to release it. It then spvalve. Failure to re-lock the plunger will fuel through the open supply valve on stashort period due to over-enrichment (flood

(iv) With fuel injected engines priming is usufuel booster, or auxillary, pump for a briefso as to permit a flow of fuel to the injecto

68. In order to avoid flooding the engine with neat fuel, a drainduction or supercharger casing so that any surplus fuel that mthis drain.

Fuel Sampling69. Before flight each day a small quantity of fuel should be ddrain valves and inspected in a glass container. Fuel should bevisual inspection shows:

(i) More than a trace of sediment

(ii) Globules of water

(iii) Cloudiness



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or a paper or chemical detector.

aircraft fuels:

T A-1 is undyed and can vary in.

ar as droplets on the sides of thettom.

ll cause the fuel to appear hazed or

ay be suspended in the fuel or settle

right. Clear refers to the absence ofo the sparkling appearance of fuel

o obvious separation of fluids, testty the sample onto a smooth surfaceld be all water !



(iv) A positive reaction to water finding paste,

The following should serve as a guide to the visual assessment of

(i) Colour. AVGAS is dyed blue, AVTUR JEappearance from colourless to straw yellow

(ii) Undissolved water (free water) will appesampling vessel, or as bulk water in the bo

(iii) Suspended water (water in suspension) wicloudy.

(iv) Solid matter (rust, sand, dust, scale, etc) mon the bottom of the sampling vessel.

(v) The fuel sample should appear clear and bsediment or emulsion and bright refers twhich is free from cloudiness or haze.

70. Because of its lower SG, fuel floats on water. If there is nthe sample by smell, colour and finally by evaporation test. Empand see how quickly it evaporates. The fluid in the container cou





Piston Engine Power Transmissions and Propellers

Introduction

Propeller Reduction Gearing

Torquemeters

Fixed Pitch Propeller

Variable Pitch Propellers

Synchroniser and Synchrophaser Systems

Propeller Checks

Aircraft and Engine Protection

Forces Acting on a Rotating Propeller



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nsmissions

er, or torque delivered by the enginehrough the air. In the case of all butry between the engine and propeller.

ngine ideally needs to operate at aeds.

aft speed to the speed suitable fortion gears are always used on radial-line engines. Any of the above may

while a ball or roller thrust bearingarings and a thrust washer performicating oil system. Three types of

Piston Engine Power Transmissions and Propell


17Piston Engine Power Traand Propellers

Introduction1. The purpose of the propeller is to convert the rotary powcrankshaft into propulsive power, or thrust to drive the aircraft tlight single engined aircraft one form of rpm reduction is necessa

Propeller Reduction Gearing2. In order to achieve optimum power output, a piston ecrankshaft speed which is well in excess of efficient propeller spe

3. The propeller reduction gear reduces the engine crankshefficient operation of the propeller. Epicyclic or planetary reducengines whilst spur gear reduction gearing is generally used on inbe used on horizontally opposed engines.

4. Roller bearings will normally support the propeller shaft,transfers propeller thrust to the airframe. Occasionally plain bethese functions. Lubrication is normally from the engine lubrreduction gear are illustrated at Figure 17-1.



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FIGU

ns which exert a thrust on the fixedction to the crankshaft fear, and theut. In order to measure the thrust

loat, being attached to the structure Engine oil pressure, boosted by atted with a bleed back to the engine



RE 17-1

Torquemeters5. As the crankshaft gear rotates it drives the propeller piniogear teeth. This fixed gear will tend to rotate in the opposite direthrust produced will be directly proportional to the power outpwhich is applied to the fixed gear, the fixed gear is allowed to fvia pistons and oil-filled cylinders, as shown at Figure 17-2. torquemeter pump, is fed to the cylinders and each cylinder is fioil system.



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um. The bleed ports are fully open torquemeter pressure gauge readingwill increase and the pistons will beartially covered and the oil pressureed gear. The torquemeter pressureressure).

FIGU



6. At low engine power, thrust on the fixed gear is at a minimwith a resulting low oil pressure in the system. Consequently thewill be low. As power is increased, the thrust on the fixed gear forced further in to their cylinders. Now the bleed ports will be pon the pistons will be increased to match the thrust on the fixgauge is usually calibrated to read BMEP (brake mean effective p

RE 17-2



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ounted on a central hub, which is propellers have a maximum of four-3, with a cross-section through the will be seen that the blade has an aerofoil creates lift in the same waydicular to the aircraft’s longitudinal

FIGUPropelNome

ade between the chord line of thes the geometric pitch angle or, more



7. The aircraft propeller consists of two or more blades mattached to the engine crankshaft. As a general rule piston engineblades. A typical two-bladed propeller is illustrated at Figure 17blade indicating the standard propeller blade nomenclature. Itaerofoil section similar to that of a wing. As the blade rotates theas a wing, but since the plane of rotation of the blades is perpenaxis this lift acts longitudinally and is referred to as thrust.

RE 17-3ler Blade nclature

Fixed Pitch Propeller8. The pitch of a propeller is determined by the angle mpropeller blade and the plane of rotation. This angle is known acommonly, the blade angle and is illustrated at Figure 17-4.



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FIGUBasic PGeom



RE 17-4ropeller

etry



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fixed and cannot be changed whilstlades will depend upon the speed ofrom Figure 17-5 it will be seen thate factors. These are the blade angle, the aircraft.



9. A fixed pitch propeller is one in which the blade angle is the propeller is rotating. The thrust developed by the propeller brotation of the propeller and the angle of attack of the blades. Fthe angle of attack of the propeller blade is dependent upon threthe speed of rotation of the propeller and the forward velocity of



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FIGUFixed-PPropelAttack



RE 17-5itch

ler Angle of



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rward speed of the aircraft with theive angle of attack of the blades toade angle to increase. There is onlyamely about 4°.

e is only one airspeed at which theack it will produce less thrust, butine. At greater angles of attack theut 15°, but blade drag will increasesively greater.

e designed to operate at maximumffective angle of attack of the bladesy. The option to increase rotationalecause of the extra torque required.propeller must clearly be capable ofe blades must not exceed 15°.

a fixed pitch propeller is thereforedes. Since the useful rpm range of a

ble flight speeds the only remainingce the development of the variable-



10. The blade angle is, by definition, fixed so increasing the fopropeller rotating at constant speed (rpm) will cause the effectdecrease. Conversely, increasing propeller rpm will cause the blone angle of attack at which any aerofoil is at its most efficient, n

11. It follows therefore that, at a given propeller rpm, therpropeller will be at maximum efficiency. At lesser angles of attblade drag will be lower so less torque is necessary from the engblades will produce more thrust, up to the stalling angle of aboand the engine torque needed to maintain rpm would be progres

12. For reasons of economy, most fixed pitch propellers arefficiency when the aircraft is in the cruise. At lower speeds the eis greater and the propeller is operating at less than full efficiencspeed (rpm), and reduce the effective angle of attack is limited bFurthermore, at the low forward speed of take-off or climb the producing thrust and therefore the effective angle of attack of th

13. The available range of flight speeds for an aircraft withclearly limited by the range of effective angles of attack of the blapiston engine is also very limited, to increase the range of possivariable is the blade angle, or pitch of the propeller blades. Henpitch propeller.



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ut the aim in all cases is to achievepeeds as possible. By this means thegreatest efficiency over a range of

propeller the blades can be rotatednly be done on the ground with theossible blade angles, by which the

It is a very early attempt at variable

e angle can be changed by the pilotlow speed flight (e.g. take-off), or ace the propeller has two settings at

flight speeds over which it achievese danger of selecting coarse pitch forspeed before it runs out of runway.ellers, of which the most popular for



Variable Pitch Propellers14. There are a number of types of variable pitch propeller, bmaximum propeller efficiency over as wide a range of aircraft sconversion of engine torque into thrust is achieved with the airspeeds instead of only at one specific airspeed.

Ground-Adjustable Propeller. In this type of variable pitchin the hub, using special tools, to alter the blade angle. This can oengine shut down. The system offers a very limited range of peconomical cruising speed of the aircraft can be varied slightly. pitch propellers and is rarely found nowadays.

Two-Position Propeller. With this type of propeller the bladin flight to give either a low blade angle (low or fine pitch) for high blade angle (high or coarse pitch) for high speed flight. Henwhich it achieves maximum efficiency and a reasonable range ofacceptable efficiency. One of the worst aspects of the system is thtake-off, whereupon the aircraft is incapable of achieving flight Two-pitch propellers were superseded by controllable-pitch propuse with piston engines is the constant speed propeller.



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tilises a governor, situated in the engine throttle setting or airspeed., to maintain the effective angle ofh-change mechanism. The setting ofs of a propeller lever in the pilot’ster the propeller speed will result ind.

ward speed increases. We know thatdes (see Figure 17-5). The reduced

to rotate faster because the torquemediately the governor senses the

ases their angle of attack and bladefective angle of attack of the blades

ard speed would decrease and theler rpm would tend to fall, so thelade torque and maintain constantdes returns to the optimum value

ng engine power, the engine wouldsing the initial rpm increase, would

d therefore increase, preventing anyd constant value. Furthermore, theditional torque, and airspeed would angle for maximum efficiency.



Constant Speed Propeller. The constant speed propeller upropeller hub, to maintain constant propeller rpm regardless ofThe governor achieves this by altering the propeller blade angleattack of the blades constant, through a hydraulic or electric pitcthe governor-controlled propeller speed is adjustable by meancockpit. Having selected a desired rpm, anything that tends to althe governor adjusting the blade angle to maintain constant spee

15. Suppose for example the aircraft is put into a dive and forthis will reduce the effective angle of attack of the propeller blablade torque, or resistance to rotation, will cause the propellerbeing applied to it by the engine has not changed. However, imincreased rpm it drives the blades to a coarser angle, which incretorque. Consequently propeller speed is held constant and the efis returned to its optimum value.

16. Conversely if the aircraft were put into a climb the forweffective angle of attack of the blades would increase. Propelgovernor would drive the blades to a finer pitch to reduce bpropeller rpm. Once again the effective angle of attack of the bla

17. Similarly, if the pilot were to open the throttle, increasitend to speed up and drive the propeller faster. The governor, sendrive the propeller blades into coarser pitch. Blade torque woulfurther rise in propeller speed and restoring rpm to the selectecoarser blade angle would enable the propeller to absorb the adincrease until the blade angle of attack is restored to the optimum



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blade angle, to maintain constant

eds of a two-pitch propeller and a

FIGUPropelEfficien

rate various propeller blade ranges

windmilling and therefore creatingturboprop aircraft are fitted withered the leading edge of the blade isother words the angle of attack is 0°



18. A reduction in engine power has the reverse effect onpropeller rpm and optimum angle of attack.

19. Figure 17-6 compares the efficiency with varying airspevariable pitch propeller.

RE 17-6ler cy

20. With a variable pitch propeller it is possible to incorpowhich fall outside the normal fine and course blade angle limits:

(a) In order to prevent the propeller of a failed enginea great deal of drag, many piston engines or feathering propellers. When the propeller is feathmore or less edge on to the oncoming airflow. In and the pitch angle in the region of 85°.



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g and any flight condition wouldhe aircraft taxiing (hopefully) verylade angle be made even finer, toe possible to fit a device which willanoeuvring range of blade angles.

ft is on the ground. Were a ground is probable that the angle of attackorne reverse thrust. For this reasonuarded by an undercarriage squat

is sometimes known as the BETA include the negative blade angles

erse thrust. Now the blade angle iss negative. The airflow from thehe aircraft.

erse thrust is selected and certainlyt to confuse the noise level with theircraft’s kinetic energy will still be

aircraft it is not a bad idea to dwell ensure that you do not have an create directional control problems



(b) The limit of fine pitch used for take-off, landinnormally be in the order of 10° to 12°. With tslowly on the ground it is desirable that the bmaintain a sensible angle of attack. It is thereforenable the propeller to be shifted into a ground mObviously this is desirable only when the aircramanoeuvring blade angle to be selected in flight itwould be come negatives, and you’ve invented airbthe ground manoeuvring selector is normally gswitch.

The range of ground manoeuvring blade angles range, and the BETA range is often taken toinvolved in producing reverse thrust.

(c) Braking propellers provide what is effectively revmade negative and the angle of attack becomepropellers now opposes the direction of travel of t

Braking propellers make a lot of noise when revhelp to decelerate the aircraft. It is important nolevel of efficiency. Something like 80% of the adissipated through the wheel braking system.

When selecting reverse thrust on a multi-engined for a second or two at 20% reverse power toasymmetric reverse situation, which will obviously



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ground manoeuvring selector, thecarriage squat switch.

can provide reverse thrust) with aname suggests, prevents a feathered

ed above.



when 100% reverse power is selected. Like thereverse selector is normally protected by an under

Do not confuse a braking propeller (one that propeller brake. This is a device which, as the propeller from rotating in the airflow.

The diagram at Figure 17-7 shows the blade angle ranges describ



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FIGU

orce the blade towards a low pitcheing feathered, the propeller wouldhigh drag condition.



RE 17-7

21. The centrifugal turning moment of the blade tends to fangle. In the event of an engine failing and the propeller not btherefore tend to go to the fully fine position, with a consequent



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ble, it is a common feature of manytive devices are used to ensure thatse blade angle, thereby minimising

s illustrated at Figure 17-8. This israulic power to the propeller pitch-arse pitch, as necessary. A slightly



22. Since the situation described above is obviously undesiraconstant speed propeller systems that counterweights or alternathe uncontrolled/unfeathered propeller maintains the fully coardrag.

Constant Speed Units23. A type of hydraulic constant speed propeller governor iknown as a double-acting unit since the governor supplies hydchange mechanism to drive the blades toward either fine or cosimpler, single-acting system will be explained subsequently.



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FIGUDoublPitch C



RE 17-8e-Acting hange Unit



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ned axially by either the governores in a ported cylinder to which thehe ports in the rotating cylinder are

oil system, to a return line to thehanism. The piston valve is carefullynd fine pitch sides of a piston in the

ller pitch-change piston against the is required and this is achieved byine lubricating oil system and gear-

s, known as bellcranks, which are speed of the engine, the faster therce acting upon them. Under theis of rotation and the bellcranks liftpring.

rcomes the centrifugal force of the

he governor spring force. Reductionpressing the spring until centrifugalsing the spring force will force thealance the two forces once more.



24. A piston valve is attached to a spindle, which is positioflyweights or the pilot’s propeller rpm lever. The piston valve slidflyweights are attached and which rotates, driven by the engine. Tconnected to a pressure oil supply from the engine lubricatingengine sump and to either side of the propeller pitch control mecprofiled so that it controls the oil supply to and from the coarse aPitch Change Unit (PCU).

25. Because considerable force is needed to move the propeforces acting on a rotating propeller, a high-pressure oil supplymeans of a booster pump. This pump is supplied from the engdriven from the crankshaft.

26. The governor flyweights are attached to L-shaped armpivoted to lugs at the upper end of the cylinder. The higher thegovernor flyweights rotate and the greater the centrifugal foinfluence of this force the flyweights move outward from the axthe piston valve in its cylinder, against the force of the governor s

27. If engine speed is reduced the governor spring force oveflyweights and the piston valve is moved down in its cylinder.

28. Movement of the pilot’s propeller rpm lever may adjust tof the spring force will allow the piston valve to move up, comforce matches spring force. Compressing the spring and increapiston valve down until the spring has expanded sufficiently to b



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governor spring. The tension of thisrpm datum, when the spring and

d Unit (CSU) in the three principalhen the propeller is operating at theen the propeller rpm is less than that

is higher than that selected.

17-8. In this situation the propellergal force exerted by the governor piston valve is held in mid-position,nder. Oil cannot flow to or from thehe propeller blades at their presentand ensuring that the propeller is

two reasons. The first is simply that rpm than was previously set. Thelimb without adjusting the throttlepeller, and the rpm would decay.



29. It will be noted that a smaller spring is located above the spring can be adjusted with the engine stationary to set the flyweight forces are in equilibrium.

30. Let us now consider the operation of the Constant Speeoperating situations. These are firstly the on-speed condition, wrpm selected by the pilot, secondly the under-speed condition whselected and thirdly the over-speed condition, when propeller rpm

The on-speed condition. This is the condition illustrated at Figureis rotating at the rpm selected by the pilot and the centrifuflyweights balances the force exerted by the governor spring. Thewhere its "lands" cover the oil supply and return ports in the cyliPCU, and so the PCU piston is hydraulically locked, holding tblade angle, which is achieving the optimum angle of attack operating at maximum efficiency.

The under-speed condition. This situation will occur for one of the pilot has moved the propeller rpm lever, selecting a highersecond situation might occur when the aircraft is put into a csetting. Initially the propeller would behave like a fixed pitch pro



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or spring force would override there move downwards, uncovering the fine pitch line, and allow oil that is

e coarse pitch line to the inlet side ofpropeller will now move to a finer

aches the desired value the governor

FIGUUnderPropel



31. The result of this lower rpm would be that the governreduced flyweight centrifugal force. The piston valve will therefooil supply and return ports. This will feed high-pressure oil to thedisplaced from the pitch change cylinder to flow back through ththe booster pump. This situation is shown at Figure 17-9. The pitch, which will result in the desired increase in rpm. As rpm reflyweights return the piston valve to its mid, or on-speed position

RE 17-9-Speeding ler



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that described in the under-speedottle spring. The piston valve movese PCU piston. The fine pitch side ofd the pressure difference across the coarser pitch. The propeller rpmed and the on-speed condition is



The over-speed condition. This situation is the reverse ofcase. Now the governor flyweights exert more force than the thrupwards and oil pressure is supplied to the coarse pitch side of ththe piston is connected to the inlet side of the booster pump anPCU piston moves it to change the propeller blade angle to aconsequently falls until flyweight/spring equilibrium is restorregained. This is illustrated at Figure 17-10.



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FIGUOver-SPropel



RE 17-10peeding ler



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raft in flight, the airflow over the rotate in the normal direction off the propeller blades, the higher thele situation, since the windmillinger driving a failed engine may well

possible with many constant speedis permits the propeller blades of aa position where their leading edges so that they do not rotate.

pm lever is moved through the lowlve, completely overriding the rpmo drive the PCU piston to the full

wn the engine. As propeller/enginells, so an alternative feathering oilpeller is fully feathered. This supply

1.



Propeller Feathering32. When the engine is not driving the propeller of an aircaerofoil section blades creates a lifting force causing them torotation, exactly like a child’s toy windmill. The finer the pitch owindmilling rpm of the propeller. This is a highly undesirabpropeller creates a great deal of drag. Furthermore, the propelllead to greater damage, overheating and even fire.

33. To avoid these problems a further operating condition ispropeller control units, known as the feathering condition. Thfailed or stationary engine to be moved through coarse pitch to face into the air stream and the airflow over them creates no lift,

The feathering position. To feather the propeller the propeller rrpm (fully coarse) position. This physically lifts the piston vaspring and governor weights and allowing the booster pump textent of its travel in the coarse pitch direction.

34. At the same time the engine throttle is closed, to shut dorpm decays the output from the engine-driven booster pump fapressure supply is provided to maintain oil pressure until the promay be from an electrically driven pump or an accumulator.

35. The CSU with feather selected is illustrated at Figure 17-1



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FIGUFeathePropel

going text is a double-acting unit, initch. Many aircraft employ single-

he blades toward fine pitch, but thece of a strong spring.

ed to many light and medium sized



RE 17-11red ler

36. The constant speed propeller system described in the forewhich oil pressure is used to drive the PCU to coarse or fine pacting propellers in which oil pressure from a governor drives tblades automatically move toward coarse pitch under the influen

37. A single acting propeller control unit mechanism (as fittaircraft) is shown at Figure 17-12.



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FIGUSingle-Chang



RE 17-12Acting Pitch e Unit



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since oil is responsible for changinger contains a piston, and oil underce of the piston. The piston and rod the angle of the propeller blade. As the propeller blades move to a finerased from the PCU cylinder and thee. This system has the advantage ofse or feathered position in the event

ingle-acting propeller is simpler in only has to control the supply andted at Figure 17-13.



38. The PCU and its associated CSU are termed single actingthe propeller pitch angle in one direction only. The PCU cylindpressure is fed through the centre of the piston rod to the front famove axially under the influence of oil under pressure to changeoil is introduced, the piston moves to the rear of the cylinder, andpitch angle. In order to achieve a coarser pitch, oil pressure is relespring plus the counterweights force the blade to a coarser anglensuring that the propeller will automatically go to the fully coarof loss of the high-pressure oil supply.

39. The constant speed unit governor mechanism for the sconstruction and operation than the double-acting unit, since itreturn of oil on one side of the PCU piston. Such a unit is illustra



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FIGUCSU GSingle-Propel

angle via the centrifugal governormp which boosts engine oil pressurernor fly weights is opposed by thel spring is set by the position of thelve is determined by the balance of

eights, and the force exerted by theer oil line is closed by the governorton (see Figure 17-12).



RE 17-13overnor for Acting ler

40. The CSU at Figure 17-13 controls the propeller blade flyweights, the governor valve, and the engine-driven CSU oil puto a suitable level. The centrifugal force which moves the govecontrol spring, while the datum setting or loading of the contropilot's propeller control lever. The position of the governor vaforces achieved between the centrifugal force of the governor flywcontrol spring. When the two forces are in balance, the propellvalve, and oil pressure is trapped in the cylinder ahead of the pis



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tions to be considered.

be set to maximum or full increasenor valve will be fully down (seell be directed under pressure to the of the feathering spring, forcing thees to the fully fine setting.

ing on the weights will progressivelye the governor valve is blocking thetating at maximum rpm, and any

rther, allowing oil to flow from theston forward in the cylinder, and theg any further increase in propeller

r system, in this condition the forcee control spring, the governor valveintained.

he position of the propeller controlme that a lower rpm is selected (the

control spring will be reduced, andil will flow out of the pitch control

feathering spring, and the rpm will control spring is re-established.



41. During normal operation, there are several different situa

On the ground. Here the pilot's propeller control lever willrpm, while the throttle will be set to low power. The goverFigure 17-13), and therefore the oil from the CSU oil pump wiforward side of the piston. This pressure overcomes the pressurepiston into its rearward position, and therefore the propeller blad

42. As the power and rpm are increased, centrifugal force actlift the governor valve until, eventually, a point is reached wherflow of oil into the propeller oil line. The propeller is now roincrease in this rpm will cause the governor valve to lift still fupitch control cylinder. The feathering spring will now push the pipropeller will move to a coarser blade angle, thereby preventinrpm.

The on-speed condition. As with the double-acting propelleexerted by the governor fly weights is balanced by the force of thblocks the inlet to the propeller oil line, and the status quo is ma

43. Should the pilot change the desired rpm setting (move tlever), the system will respond to re-establish the status quo. Assupropeller control lever is moved rearwards). The loading on thethis will allow the governor weights to lift the governor valve. Ocylinder, the blade angle will coarsen under the influence of thedecrease, until the balance between the governor weights and the



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ases because of say, a decrease indescend to admit oil under pressureus increase the rpm to the original

he under-speed condition describedf oil pressure from the PCU cylindercoarser pitch, reducing rpm to the

fully rearward, the governor valve governor weights and the controloves forward under the influence ofcoarse pitch to the fully feathered

ndesirable in a multi-engine aircraft,twin-engine aircraft this asymmetricen impossible to control in certain multi-engine installations, but are



The under-speed condition. If the propeller loading increairspeed, the rpm will initially decrease. The governor valve will into the PCU cylinder to drive the blades to a finer pitch and thsetting.

The over-speed condition. This is the exact opposite of tabove. The governor valve will rise to allow a controlled escape oand the feathering spring will move the propeller blades to a desired value.

The feathering position. When the propeller lever is movedis lifted fully up, in effect negating the imbalance between thespring. Oil flows out of the pitch control cylinder as the piston mthe feathering spring, and the propeller blades travel through position.

44. The high drag of a windmilling propeller is particularly uwhere it will inevitably give rise to asymmetric yawing forces. In yaw may be so severe as to render the aircraft difficult or evsituations. Consequently, feathering propellers are essential inrarely found in single-engine aircraft.



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rward direction) whenever it is at ahen the aircraft is stationary on the

ved to a negative blade angle. Thisn a rearward direction, enabling the turbine powered propeller aircraftft. Apart from one or two highlyund in single-engine aircraft.

have the capability to use, on thesuperfine. With the blades in such auseful whilst taxying, particularly

n start-up or to increase drag during

tted, they may only be used on thet fine position, and flight fine pitchthe air is known as the alpha range.s the beta range. The beta range isgate in the power lever.



Reverse Thrust45. A propeller will produce reverse thrust (i.e. thrust in a reanegative angle of attack. Perhaps the simplest example of this is wground (no forward velocity), and the propeller blades are moproduces a negative angle of attack and thrust will be produced iaircraft to reverse taxi. This capability is a feature of many gas(turbo-props), but is rare nowadays in piston engined aircraspecialised stunt examples, it is not a feature that is commonly fo

46. Some aircraft, whilst not having a reverse capability, doground, a very fine blade angle, often known as ground fine or position the thrust produced will be minimal. This may be downhill, and it may also be used to minimise propeller torque othe landing run.

47. With either a superfine only or full reverse capability figround. In the air the propeller may only be driven to the flighstops are fitted in the hub. The range of movement available in On the ground, any blade angle less than flight fine is known ausually selected via a gate in the propeller rpm lever or a reverse



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ystemspropellers of a twin-engine aircraftare continuously subject to smallems were acting independently, theed. To overcome this problem, someh an automatic synchroniser systemm. In such a system one propeller is

s necessary to ensure that the twoans ensuring that the relative bladed or behind, remains the same. This



Synchroniser and Synchrophaser S48. To ensure balanced thrust it is clearly desirable that the rotate at the same speed. Since constant speed propellers adjustments from their CSU’s, if the two propeller control systtwo propellers would rarely be operating at exactly the same spetwin-engine aircraft with constant speed propellers are fitted witto ensure that both propellers are always operating at the same rpselected as Master and the other is slaved to it.

49. To avoid vibration due to out-of-phase frequencies it ipropellers are rotating in phase with each other. This usually mepositions of the two rotating propellers, when viewed from aheais illustrated at Figure 17-14.



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FIGUSynchrPropel

ol system that accomplishes bothintaining the relative phase angle ofight deck noise.



RE 17-14ophased lers

50. Many twin-engine aircraft employ a combined contrsynchronising (matching propeller rpm) and synchrophasing (mathe propeller blades) which also minimises propeller and hence fl



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damage in the form of indentations, or cracking that could cause stress

nes and other hard objects, and soscores, breaks in the surface finish,p damage must be removed and ane is removed by sanding to a new

similar manner. Small longitudinalre of sawdust and glue as filler and

ll cause stress concentrations, whichinspected for nicks, dents, cuts and carried out using a smooth file andlt should be a smooth and shallow

ad to cracks should be blended outntil the propeller is removed.

have severe surface damage must bef a propeller blade, or cracks on anyemoval of the propeller. Damage onle damage on the outer third of the



Propeller Checks51. Aircraft propellers must be regularly checked for signs ofor gouges that could lead to out-of-balance forces and vibrationfailure.

52. Wooden Fixed Pitch Propellers are easily damaged by stoshould be inspected frequently for nicks, cracks, delamination, and for security of the leading edge sheath. Deep cuts and deeinsertion repair carried out. If inner trailing edge or tip damagprofile, then the other propeller blade should be reshaped in acracks and minor indentations may be repaired by using a mixtusanding smooth.

53. Indentations and scores in metal Fixed Pitch Propellers wiin the end will cause failure. This type of propeller needs to be corrosion or for any other surface damage. Minor repairs may beemery cloth to remove metal from the damaged area. The resudepression.

54. In Variable Pitch Propellers, cuts or gouges which may leas soon as noticed, though dents and minor erosion may be left u

55. Metal Propeller blades that are twisted, cracked, bent or considered unserviceable. Any damage found on the inner third opart of the propeller, must be considered as immediate cause for rthe middle third of the propeller blade may be blended out, whipropeller blade may be removed by using a hacksaw.



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track, in the same plane. Blades thatmbalance causes propeller inducedd statically by manually rotating the from it does not vary by more thanariable pitch blades into fine pitch

e engine running and using a strobee tip of each blade.

ced in terms of weight distributionith the propeller removed from the be checked dynamically with thee of vibration when the propeller is

ded propeller is slightly heavier thanof balance force that gives rise to alocation, sensitive only to radiald a strobe light or magnetic pick-up vibration frequency it is possible to



Checking Propeller Track 56. It is important that the blades of a propeller all rotate, or are out of track do not produce equal thrust and the thrust ivibration. On an installed propeller blade tracking can be checkeblade tips past a fixed reference point and checking that clearancea limited amount, typically about 1½ mm. It is usual to put vwhilst carrying out the check.

57. A dynamic check of blade tracking can be made with thlight to illuminate strips of reflective tape placed accurately on th

Checking Propeller Balance58. It is also important that the propeller blades are all balanto avoid out-of-balance vibration. This is checked statically waircraft and placed on a balance stand. Propeller balance canpropeller installed by devices that sense the amplitude and phasrotating at various speeds.

59. To take a simple example, suppose one blade of a two-blathe other. When the propeller is rotating this will cause an out-once-per-revolution vibration. An accelerometer in a fixed movement, is used to measure the amplitude of the vibration anis used to measure blade phase angle. By matching phase angle todetermine which blade is causing the vibration.



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d propellers to protect the enginee allowed to become too fine, andme too coarse. These pitch-limiting

that the propeller of a failed engineered by the pilot. For example, well automatically drive the propellere failure.

d, the pressure in the cylinder will the propeller will slowly reach thering. Since this could well causert, a centrifugal latch is fitted in theeathered position, as illustrated at



Aircraft and Engine Protection60. A number of devices are associated with constant speeagainst overspeeding, which could occur if propeller pitch werovertorquing, which could occur if pitch were allowed to becodevices are grouped together under the heading of pitch locks.

61. Similarly, aircraft handling must be protected by ensuringeither feathers automatically, or can be quickly and easily feathhave seen how the feathering spring of a single-acting PCU wiblades to the feather position in the event of governor oil pressur

62. However, when the engine is shut down on the grounslowly bleed away as a result of leakage through the CSU. Thusfeathered position under the influence of the feathering spunacceptably high loads on the engine during the subsequent stapitch control cylinder to prevent the piston moving to the fFigure 17-15



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FIGUPropelCentri

ound idle (the beta range), the latch that the piston can only be movedh. With increasing rpm, centrifugalw the piston full range of movement



RE 17-15ler fugal Latch

63. When the engine is shut down, or when the engine is at grweight tension springs position the latch weights in such a wayforward a short distance, holding the blades at ground fine pitcforce will disengage the latch weights from the latch stops to allofrom fully fine to feather pitch.



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s potential hazards in flight. In the example, seizure due to loss of oil),e circumstances it is imperative that

he fail to do so before the rpm fallsit will be impossible to feather the problem that, if it were to occur atrable.

ry low power settings. Consider apeller lever) is set to 2,300 rpm andsed the engine produces less power. overcome this. Eventually the PCUd the blade angle will now remainh unit, the low power setting must

h that it ensures the propeller bladelt with the aircraft in flight. In this

ble, for operating efficiently at loweeded. The flight-fine pitch stop is when the aircraft weight is on the



64. It should be noted however that this system has seriouevent of an engine failure caused by a major mechanical fault (forthe rate at which the engine decelerates may be rapid. Under thesthe pilot takes immediate action to feather the propeller. Shouldinto the latch engagement regime (typically about 1000 rpm) propeller. The pilot is now faced with a serious asymmetric dragthe low airspeed associated with climb-out, could prove irrecove

Pitch Stops65. The constant speed range will not be maintained at vepiston engined aircraft, and assume that the rpm selector (the prothe throttle to 23" Hg (inches of mercury). As the throttle is cloThe CSU senses a reduction in rpm and fines off the propeller topiston will contact a fine pitch stop located in the cylinder, anconstant. Since the propeller has become effectively a fixed pitcnow result in a decreasing rpm.

66. The position of the fine pitch stop in the PCU may be succannot move to such a fine pitch that reverse thrust would resucase it becomes known as a flight-fine pitch stop.

67. During ground manoeuvring a finer pitch setting is desirataxi speeds and ensuring that maximum taxi speed is not exctherefore withdrawn, either automatically or by pilot selection,wheels.



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pellers it will be necessary for theve the blades into reverse pitch. Thisgh a gate, into the reverse range.ve pitch a reverse pitch stop may be

ion is similarly limited, either to a feather blade angle, typically abouthen the pilot selects ‘feather’.

propeller is for the blades to moveone means of preventing this is itsbe equally effective in the event of ahe high-pressure oil to the CSU.

lves, which close when the pressure a pre-set value. This now effectivelyk ensures that the piston within thegle does not change.

has effectively become a fixed pitchce rather than design. If the failuretch), and the rpm would be high.d great care must be taken to avoid



68. Should the aircraft be equipped with reverse-pitch proground fine pitch stop to be withdrawn to permit the PCU to driaction is often achieved by moving the propeller lever throuSimilarly, if it is necessary to limit the extent of reverse, or negatiincorporated in the PCU.

69. The travel of the PCU piston in the coarse pitch directmaximum coarse pitch to prevent engine overtorquing or to the85°. Where a coarse pitch stop is incorporated, it is withdrawn w

Hydraulic Pitch Lock70. The natural tendency for a windmilling variable pitch towards fine pitch. Having considered the double-acting CSU, ability to achieve hydraulic pitch lock. This is a system that will failure of the engine, or of the oil booster pump which supplies t

71. All that is involved is strategically placed spring loaded vaof oil in the feed lines to the pitch control mechanism falls belowisolates the cylinder of the pitch control unit. The hydraulic loccylinder does not move, and therefore that the propeller blade an

72. Once the hydraulic lock is established then the propeller propeller. The blade angle, however, is a matter of circumstanoccurred in the climb the blade angle would be low (fine piSubsequent control of rpm would be effected with the throttle, anan engine over-speed condition at high power settings.



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nd on larger transport aircraft, forlost through governor failure, is theed above, a propeller pitch lock will

erated valve will close at typicallyblade angle. The valve is then offsete bleed shut-off valve will now closeng the propeller. A hydraulic lock ise will remain locked in this position maximum. Even with the pitch lock

ted to incorporate auto-featheringh, which activates when the engine solenoid valve that enables oil from same time it causes the CSU pistonil to be fed to the coarse pitch feed feathered position.



Enhanced Propeller Pitch Lock73. A slightly more sophisticated system, normally only foupreventing excessive over-speed of a propeller should control be enhanced propeller pitch lock system. As with the system describform automatically, the pilot cannot control it.

74. As engine rpm increases during take-off, a flyweight-op95% of maximum rpm, in order to prevent a rapid decrease in by the propeller governor to permit 100% of maximum rpm. Thif rpm increases by 10% and the governor is no longer controlliformed and any further rpm increase is prevented. The bleed valvuntil the propeller rpm falls to some low figure, typically 50% ofengaged the propeller could be feathered if necessary.

Auto Feathering75. The double acting propeller system can easily be adapdevices. The auto-feather sequence is initiated by a torque switctorque falls below a pre-set threshold. The torque switch opens aa feathering pump to flow into the CSU, (see Figure 17-8). At thevalve to be lifted fully, allowing the feathering pump supply of oline, driving the propeller through the fully coarse position to the



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lleristing and bending tensile forces., air resistance to rotation applies an of rotation and the thrust load on

them, which tend to alter the bladees towards coarse pitch, but a muchds fine pitch.



Forces Acting on a Rotating Prope76. A rotating propeller is acted upon by centrifugal, twCentrifugal force tends to stretch the blades away from the hubtorsional force that tends to bend the blades opposite to directiothe blades tends to bend them forwards.

77. Variable pitch propellers have further forces acting uponangle (pitch). Aerodynamic twisting force tends to turn the bladgreater centrifugal twisting force tends to move the blades towar

78. These forces are illustrated at Figure 17-16.



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FIGUForcesA Prop



RE 17-16 Acting On eller





Piston Engine Operation and Handling

Handling Procedures

Range and Speed Charts for Power Setting

Power Ratings



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and

ation and in no way supersede thees.

that the propeller arc is clear ofg around which may be drawn intoIDLE CUT-OFF before connecting

ocedure should be adopted:

d ensure arc clear.

and starter motor.



18Piston Engine OperationHandling

Handling Procedures1. Please note that the following notes are general informmanufacturer's operating procedures for particular aircraft engin

Pre-Start Inspection. Ensure the aircraft wheels are chocked,equipment and personnel, and that there are no loose objects lyinthe propeller. Check ignition OFF, throttle closed, mixture to ground power. Check parking brake ON.

Starting Procedures (Carburettor Fuel System). The following pr

(a) Master Switch and boost pump ON.

(b) Throttle approximately ½ OPEN.

(c) Mixture FULL RICH.

(d) Propeller (constant-speed) to FULL INCREASE an

(e) Ignition to START - this energises both magnetos



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th magnetos energised, starter de-

seconds of start-up. SHUT DOWN

should be adopted:

lly.

re control is set to FULL RICH for ahe Auxiliary Fuel Pump switch is set

ld be adopted:

nge within 30 seconds and remains

perating range before operating the

e normal operating range. Mixtures to prevent overheating.

wer/rpm settings. Complete a static



(f) Release ignition switch to BOTH position. Boenergised.

(g) Check oil pressure rises to normal range within 30IF IT DOESN'T.

Starting Procedures (Injection Systems). The following procedure

(a) The mixture setting is set to IDLE CUT-OFF initia

(b) After switching on the Fuel Boost Pump the mixtubrief period to prime the engine. In some systems tto PRIME for a few seconds prior to starting.

After Start and Operating Checks. The following procedure shou

(a) Check that the oil pressure rises to the normal rawithin limits.

(b) Allow the oil temperature to rise to the normal oengine at high power.

(c) Allow the cylinder head temperature to rise to thcontrol should be RICH during ground operation

(d) Check that the manifold pressure is normal for poboost check.



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R and in each case check for a droplimits (typically 50-125 rpm).

ontrols, that is to say MAP against

:

smoothly, typically 2 to 3 seconds

s soon as practical, to avoid excess

to avoid cylinder head cracking due

xcessive torque between engine and

vance throttle

ust mixture.



(e) Move IGN switch from BOTH to L and BOTH toin engine rpm which is within the manufacturer's

(f) Check engine response to operation of propeller crpm.

Power Adjustments. The following procedure should be adopted

(a) Throttle movements should be made slowly andfrom full open to closed or vice versa.

(b) After take-off, reduce power setting to CLIMB aCHTs.

(c) In the event of high CHT, reduce power gradually,to over-rapid cooling.

(d) In constant-speed propeller aircraft always;

(e) increase rpm in advance of throttle (MAP)

(f) reduce MAP in advance of rpm in order to avoid epropeller.

The basic procedures are;

(a) to increase power enrich mixture, increase rpm, ad

(b) to decrease power retard throttle, reduce rpm, adj



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uring a long descent, periodically

e cruise, in accordance with thet and landing the mixture control

uld be decreased at high altitudes on

larly likely during reduced power

to fall to normal low power valuesrocedure is not followed conducted,n by setting mixture to IDLE CUT-e cowl flaps OPEN until the engineication systems, wait an appreciableck to the sump.

if care is taken in its operation ande engine is subject to two types of



2. During prolonged low power operation, for example dadvance the throttle for a short period to clear the plugs.

3. Ensure that lean mixtures are only used during thmanufacturer's instructions. At high powers and during descensetting should be FULL RICH or RICH. The mixture setting shonormally aspirated engines.

4. If there is any possibility of carburettor icing, particu(descent) operation, set carburettor heat to ON.

Shut Down Procedure. Allow the CHT and oil temperature(less than 400°C CHT) before shutting down an engine. If this presidual heat may cause damage to bearings and seals. Shut dowOFF and switch OFF ignition after the engine has stopped. Leavhas cooled, check all cockpit switches to OFF. In wet sump lubrtime before checking engine oil contents, to allow oil to drain ba

Engine Life5. The life of a piston engine can be considerably lengthenedhandling to ensure that stresses are kept as low as possible. Thstress, thermal and mechanical.



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res. Cylinder head temperatures areing excessive power (MAP) settings.trols (cowl flap opening). It should

ceptable stress on the engine, as well

ean effective pressure (BMEP). The causes large inertia loadings on the greater these loadings become. Thepower setting, clearly the higher thetroke.

ta it is always beneficial to operate MAP for each required operating



6. Thermal stresses are a function of CHT and oil temperatukept within limits by correct use of the mixture control and avoidOil temperature is controlled by correct use of the oil cooler conbe borne in mind that too low an oil temperature can place unacas too high a temperature.

7. Mechanical stresses are a function of rpm and brake mabrupt reversal of travel of the pistons at the end of each strokepistons, connecting rods and crankshaft. The higher the rpm, thebrake mean effective pressure in the cylinders is dependent upon power setting the higher the cylinder pressure during the power s

8. Within the limitations of the manufacturer’s operating daat the lowest recommended rpm and the highest recommendedcondition.



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r SettingFIGUEffectsSetting(Piston



Range and Speed Charts for PoweRE 18-1 of Power s on Range Engine)



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FIGUEffectsSetting(Piston



RE 18-2 of Power s on TAS Engine)



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charts for a Piper Apache aircraft.eas Figure 18-2 shows the relationthe charts it is possible to determine of the aircraft, given flight altitude,

of 4000 feet. If it is desired to maker (75%) would be chosen. From rpm, the TAS at this engine power gallons per hour, so it is a simple

conomy the minimum cruise powerpproximately 128 mph and fuel

gine is limited by the ability of its accelerations, and velocities. Highication films between moving parts, when running and the ultimate life

this damage. The rate of damage ise is proportional the duration ofgiven engine is therefore inverselys employed.



9. Figure 18-1 and Figure 18-2 show the range and speedFigure 18-1 shows the effects of power settings on range, wherbetween power settings and true airspeed (TAS). By referring to the correct power setting to select for any flight, within the rangedistance and desired flight time.

10. Suppose a flight of 600 miles is to be made at an altitude the flight in the shortest time then the highest cruise poweFigure 18-2 it will be seen that, at Full Throttle setting and 2400will be approximately 171 mph. Fuel consumption will be 18.8matter to calculate the fuel required in still-air conditions.

11. If, alternatively, it is required to make the flight at best ewould be chosen (45%). From Figure 18-2 TAS will be aconsumption, from Figure 18-1, will be 11.1 gallons per hour.

Power Ratings12. The amount of power that can be generated by an encomponents to withstand the resulting temperatures, pressures,temperatures in particular tend to cause a break down in the lubrresulting in rapid wear. All engines sustain damage continuouslyof an individual engine depends upon the cumulative effects of proportional to the power setting and the cumulative damagoperation at any given power setting. The useful life of a proportional to the duration and magnitude of the power setting



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ly throughout a specified lifetime,gs, some of these may be usedt engine will typically be authorised "rated" power output and also atvres such as take-off and climbing.d to ensure that cumulative damageed throughout the authorised life ofuous power are likely to result in

to 5 minutes per flight)

uise flight

ight



13. In order to ensure that engines can be operated safemanufacturers authorise a number of different power ratincontinuously whilst others are subject to time limits. An aircrafto operate continuously at a certain "maximum continuous" orsome higher power settings for limited periods during manoeuThe time limits set for these higher power settings are calculateremains within acceptable limits and catastrophic failure is avoidthe engine. Excessive use of settings above maximum continmaterial failure within the authorised life of the engine.

Typical power ratings are indicated below:

Power Rating Authorised duration

Maximum Take-Off Time limited (typically 1

Maximum Cruise Unlimited throughout cr

Maximum Continuous Unlimited throughout fl



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" and state a simple equation to

aero-engines vary with changes in

a normally aspirated piston aero-

es?

ers in aircraft engines and state the

?




QUESTIONSQUESTION 1.

Define the terms "Density Altitude" and "Pressure Altitudedetermine pressure altitude?

QUESTION 2.

Describe how the performance of normally aspirated piston ambient pressure and temperature?

QUESTION 3.

Explain the effects of aircraft altitude on the power output ofengine?

QUESTION 4.

Define the term "critical altitude" as applied to piston aero-engin

QUESTION 5.

Summarise the reasons for fitting turbochargers and superchargdifference between ground boosting and altitude boosting.

QUESTION 6.

Describe the difference between turbochargers and superchargers



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charged piston engine?

gine, its location, and its operating

raft piston engine?

light from including, before enginewer at low altitude, climb to high

titude of normally aspirated, turbo-?

re gauges (MAP and Boost)?



QUESTION 7.

Describe the principle of operation of the turbocharger?

QUESTION 8.

Explain the function of an intercooler in a turbocharged or super

QUESTION 9.

Describe the purpose of the waste gate in an aircraft piston enprinciple?

QUESTION 10.

List the methods of controlling the waste gate position in an airc

QUESTION 11.

Describe the positions of the waste gate throughout a normal fstart up, engine idle, take-off power, maximum continuous poaltitude, descent to low altitude and low power setting?

QUESTION 12.

Compare and contrast the curves of maximum power versus alcharged, and supercharged engines, identifying significant points

QUESTION 13.

Describe the purpose and operating principles of manifold pressu



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plied to aircraft piston engines?

piston engines?

?

escribe its causes and effects?

ntamination?

, "lean (weak) mixture", and "rich



QUESTION 14.

Define the terms "full throttle height" and "rated altitude" as ap

QUESTION 15.

Define the term "turbo lag" as applied to turbo-charged aircraft

QUESTION 16.

Describe how different fuel grades are identified?

QUESTION 17.

Define the term "octane rating" as applied to piston engine fuels

QUESTION 18.

Define the term "detonation" as applied to piston engines, and d

QUESTION 19.

Identify situations and power settings that promote detonation?

QUESTION 20.

Describe the method of checking piston engine fuels for water co

QUESTION 21.

Define the terms "chemically correct ratio", "best power ratio"mixture" as applied to piston engines?



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tures in aircraft piston engines?

mixture ratio?

setting?

in its disadvantage.

by comparing it with a fixed pitch

cting variable pitch propellers?



QUESTION 22.

Describe the advantages and disadvantages of weak and rich mix

QUESTION 23.

Describe the relationship between specific fuel consumption and

QUESTION 24.

Describe the use of exhaust gas temperature as an aid to mixture

QUESTION 25.

Describe the fixed pitch propeller, its operating modes and expla

QUESTION 26.

Explain why propeller blades are twisted from root to tip?

QUESTION 27.

Describe the variable pitch propeller and explain its advantagespropeller?

QUESTION 28.

Describe the operating principle of the single acting and double a



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tem for both single and multi engine

during flight, including feathering,nd altitude?

ther system?

(centrifugal latch)?

lied to aircraft propellers?

chrophasing systems?

pellers vary with flight speed?



QUESTION 29.

Describe the operating principle of a constant speed propeller sysaircraft?

QUESTION 30.

Describe the operation of the constant speed propeller system unfeathering, and changes in RPM selection and aircraft speed a

QUESTION 31.

Explain the purpose and basic operating principle of an auto-fea

QUESTION 32.

State the purpose and describe the operation of a low pitch stop

QUESTION 33.

Define the terms "Synchronising" and "Synchrophasing" as app

QUESTION 34.

Describe the basic operating principles of synchronising and Syn

QUESTION 35.

Describe how the efficiency of fixed pitch and variable pitch pro



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inciple?

condition before flight?

ring normal flight from engine start

?

er"



QUESTION 36.

State the purpose of propeller reduction gearing?

QUESTION 37.

State the purpose of a torque-meter and describe its operating pr

QUESTION 38.

Explain why it is necessary to check the propeller for its physical

QUESTION 39.

Describe the general procedures for setting the engine controls duup until shut down?

QUESTION 40.

State the possible use of time limits for take-off and climb power

QUESTION 41.

Define the term "Rated Power" or "Maximum Continuous pow



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ich the prevailing density occurs.

n from standard atmosphere)

osphere datum of 1013.25 mbs.

ation from standard atmosphere x

with both decreasing pressure andperature is to cause air to expandugh the engine at a given RPM andhas the same fundamental effect, the of reducing exhaust back pressuree. The overall effect is a gradual

pressure which tends to reduce airut. Below 36000 feet this effect isy offsets the density reduction. The6000 feet followed by a greater lossely 3.5% per thousand feet overall.



ANSWERS:ANSWER 1.

Density altitude is the altitude in the standard atmosphere at wh

Density Altitude = actual altitude + (120 x temperature deviatio

Pressure altitude is the altitude above or below the standard atm

Pressure Altitude = actual altitude + (4 x temperature devialtitude/1000 feet)

ANSWER 2.

The power output of normally aspirated piston engines reducesincreasing temperature. The principal effect of increasing temthereby reducing its density. This reduces the air mass flow throhence reduces its power output. Although reducing air pressure overall consequences are modified somewhat due to the effectswhich tends to improve the volumetric efficiency of the enginreduction in power output for a given RPM

ANSWER 3.

The principal effect of increasing altitude is a reduction in air density, mass flow through the engine and hence power outpmoderated somewhat by a reduction in temperature which partloverall effect of these factors is a gradual loss of power up to 3rate above this altitude. The average loss rate being approximat



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iston engines. As a turbo-chargeddually closes in order to increaset. This process continues until, at a

fully closed and no further increasede will result in a decrease in MAPh the waste gate becomes fully openguish this situation from the "Fullecomes fully open in attempting to

rgers in aircraft piston engines is tohe effects of decreasing air pressure.itude boosting. In ground boostingntained with increasing altitude, thede boosting MAP is not increased at.

ain MAP and hence engine powertly by a shaft from the engine ande power to drive the compressor isgine RPM and compressor RPM istterfly valve which is always locatedo direct link between the engine and



ANSWER 4.

The term "Critical Altitude" applies only to turbo-charged pengine climbs at constant power setting, the waste gate gracompressor output to maintain a constant MAP and power outpucertain altitude depending upon power setting, the waste gate isin compressor output is possible. Any further increase in altituand engine power output. Critical altitude is the altitude at whicin an engine climbing at rated power. The term is used to distinThrottle Height" in a supercharged engine, where the throttle bmaintain MAP.

ANSWER 5.

The principal reason for installing turbo-chargers and superchamaintain engine power output to higher altitudes by offsetting tThe two basic approaches employed are ground boosting and altMAP is increased above ambient pressure at sea level and is maioverall effect being an increase in power at all altitudes. In altitusea level, but sea level MAP is maintained up to a higher altitude

ANSWER 6.

Both superchargers and turbo-chargers are intended to maintoutput with increasing altitude. Superchargers are driven direchence compressor speed is directly related to engine RPM. Thdrawn directly from the engine. Because the ratio between enfixed, compressor output pressure is controlled by the throttle buupstream of the compressor. In turbo-charger systems there is n



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ough a turbine. Compressor speedetermines the proportion of engine

upled to a centrifugal turbine whichrger and hence its output flow andough the turbine. This in turn is parallel with the turbine such thatby-passing the turbine. In this wayf opening of the waste gate. The variations in MAP, air temperature,

or must carry out mechanical workto increase the static pressure of theffects of friction. This increased reducing the beneficial effects of

ing cylinder temperatures might be fuels and rich mixtures. In order to

ed between the compressor and the levels.



the compressor which is driven by engine exhaust gas passing thrand hence output pressure is controlled by a waste gate which dexhaust gas passed through the turbo-charger turbine.

ANSWER 7.

The turbo-charger comprises a centrifugal compressor directly cois driven by engine exhaust gasses. The speed of the turbo-chapressure are determined by the mass flow of exhaust gas thrcontrolled by a waste gate located in an exhaust gas passage inopening the waste gate increases the proportion of exhaust gas turbo-charger output is inversely proportional to the degree oposition of the waste gate is controlled by a servo piston sensingand pressure drop across the throttle butterfly valve.

ANSWER 8.

In order to maintain MAP with increasing altitude the compresson the air passing through it. Although the principal objective is air, its temperature is also increased, mainly through the etemperature tends to cause a reduction in air density therebyconstant MAP. More importantly, at high altitudes the resultsufficiently high to cause detonation even when using high octaneovercome this problem some systems employ intercoolers locatengine in order to reduce the temperature of the air to acceptable



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vide a means of varying compressorient conditions and engine RPM. Itpassing directly to atmosphere. The in parallel with the turbocharger by-passes the turbine to go directlyM of the turbine. Because of thismosphere, turbo-charger RPM and of opening of the waste gate.

o-charger is controlled by a spring- pressure and so the spring holds the varying servo spill off valves using

mploys an aneroid capsule sensing

mploying a nitrogen filled capsulee controller employing a diaphragm control servo spill valves.



ANSWER 9.

The purpose of the waste gate in a turbo-charger system is to prospeed and hence to maintain a constant MAP with changing ambachieves this by varying the proportion of engine exhaust gasses waste gate is located in an exhaust gas passageway connectedturbine. By varying the proportion of engine exhaust gas whichto atmosphere, it controls the gas flow through and hence RPrelationship between waste gate angle and exhaust flow to athence compressor output are inversely proportional to the degree

ANSWER 10.

The position of the waste gate in an aircraft piston engine turbbiased servo piston. With the engine shut down there is no servowaste gate in the open position. Servo pressure is controlled bythe following methods:

(a) In the most simple systems a MAP controller eMAP to control a servo spill valve.

(b) In more advanced systems a density controller esensing air temperature and a differential pressursensing the pressure drop across the throttle, each



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e pressure controller employing anssure, a rate controller employing autput pressure, and a pressure ratiombient pressure, each control servo

eld fully open by the servo pistong servo oil pressure drives the waste causing it to accelerate to the speedtes to take of power the increasing

ncrease MAP. The waste gate thenAP, the overall effect being a slighter the waste gate gradually closes

, at critical altitude the wastegate islosed as MAP and power output fallremains closed until critical altitude,sing ambient pressure. As power isn, becoming fully open after engine



(c) Still more advanced systems comprise an absolutaneroid capsule sensing compressor discharge prediaphragm sensing rate of change of compressor ocontroller employing an aneroid capsule sensing aspill valves.

ANSWER 11.

Prior to start up servo pressure is zero and the waste gate is hspring. As the engine accelerates from start up to idle speed, risingate to the closed position directing all exhaust gas to the turbinenecessary to establish the required MAP. As the engine acceleraexhaust gas output causes the turbine to accelerate tending to imoves to the almost fully open position to maintain constant Mrise in turbo-charger RPM. As the aircraft climbs at rated powincreasing turbo-charger speed to maintain constant MAP untilfully open. Above critical altitude the waste gate remains fully cwith increasing altitude. As the aircraft descends the waste gate below which it gradually opens to maintain MAP against increareduced during and after landing the waste gate continues to opeshut down.



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ltitude. The power output of suchroximately 3.5% per 1000 feet.

nt level between sea level and fullr normally aspirated engines. Alsos in a gradual reduction in exhaustfficiency and hence power output ofger results in a reduction in power

round level, followed by a gradualhich power reduces at a rate of

titude, above which it reduces in theary to drive the turbocharger is notccurs at ground level. However, theefficiency with altitude are largelyfect is a gradual reduction in powers at approximately 3.5% per 1000



ANSWER 12.

In normally aspirated engines MAP decreases with increasing aengines therefore reduces with increasing altitude at a rate of app

In supercharged piston engines MAP is maintained at a constathrottle height, above which it reduces in the same way as that foas ambient pressure reduces with increasing altitude, this resultback pressure. This has the effect of improving the volumetric ethe engine. Finally the power extracted to drive the supercharoutput.

The overall result of these factors is a reduction in power at gincrease in power output up to full throttle height, above wapproximately 3.5% per 1000 feet.

In a turbo-charged engine MAP remains constant up to critical alsame way as in a normally aspirated engine. The power necessdrawn directly from the engine so no significant loss of power oreduction in exhaust back pressure and improved volumetric negated by the restriction caused by the turbine. The overall effrom sea level up to critical altitude above which power reducefeet.



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the automatic boost controller wille altitude the throttle will be fully in MAP and power similar as if the

tle becomes fully open is termed theght for every possible power setting is termed the "rated altitude".

t in this case it is the waste gate thatn. The altitude at which this occurs

bo-charger systems requires a finitet. Each time the throttle position is

ic MAP control system must detectlibrium condition. This can lead toening, followed by an overshoot in may occur before equilibrium is minor adjustments, fine tuning the



ANSWER 13.

When a supercharged engine climbs at constant power setting gradually open the throttle to maintain constant MAP. At somopen and any further increase in altitude will result in a reductionengine were normally aspirated. The altitude at which the throt"full throttle height". There will be a different full throttle heiand for a climb conducted at rated power the full throttle height

In the case of a turbocharged engine a similar process occurs, bubecomes fully closed rather than the throttle becoming fully opein a rated power climb is termed the "critical altitude".

ANSWER 14.

The term "turbo-lag" refers to the phenomenon whereby the turtime to respond to changes in throttle position selected by the pilochanged the various sensors and controllers within the automatand react to the changing conditions before reaching a new equian initial reduction in MAP immediately after rapid throttle opMAP as the system reacts. In extreme cases several cyclesestablished, or else the pilot might be required to make repeatedsystem to the required condition.



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number, and in some cases lettersthat the fuel is aviation gasolineL). Similarly MOGAS 85 indicatesFSII) indicates aviation turbine fueldition to the above markings on fuel an aid to identification. AVGAS

ion resistance of fuels are comparedharacteristics. The first of these isecond, heptane detonates easily. Byrmance is compared with those of indicates the mixture with the samee has the same characteristics as aoline has characteristics 20% better



ANSWER 15.

Different fuel grades are identified by a combination of title, indicating additives. For example AVGAS 100LL indicates (AVGAS), of 100 octane rating (100), with a low lead content (Lmotor gasoline (MOGAS), of 85 octane rating (85). AVTUR ((AVTUR) with a fuel system icing inhibitor additive (FSII). In adtankers and containers coloured dyes are sometimes added as100LL for example is dyed light blue.

ANSWER 16.

The term octane rating refers to the process whereby the detonatwith mixtures of two chemicals of vastly different detonation ciso-octane which is extremely resistant to detonation, whilst the stesting the fuel under standard conditions its detonation perfovarying iso-octane :heptane mixtures. The octane rating numbercharacteristics as the fuel. For example an 85 octane gasolinmixture of 85% iso-octane and 15% heptane. A 120 octane gasthan those of pure iso-octane.



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air mixture explodes spontaneouslyront progresses through the mixture to 1000 feet/second. The resultingigh pressures within the cylinders.nergy of this reaction it is wasted inots. In extreme or prolonged cases

f carburettor heat, alternate air, or

onstant speed propeller system.



ANSWER 17.

The term detonation describes the phenomenon where the fuel rather than igniting and burning. In normal burning the flame fat 60 to 80 feet/second whereas in detonation it progresses at uprapid heating and expansion of the mixture produces very hBecause the pistons cannot react quickly enough to absorb the ethe form of heat causing high temperatures and localised hot spthe engine is likely to suffer structural damage.

Detonation may be caused by any or all of the following:

Excessive induction air temperatures due to inappropriate use oexcessive supercharging.

Weak mixtures or fuel of too low an octane rating.

Ignition too far advanced.

Excessive torque/power selected at too low an engine RPM in a c



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detonation:

ettor heat alternate hot air in fuel

ller.



ANSWER 18.

The following situations and power settings are likely to result in

Too weak a fuel:air mixture.

Ignition too far advanced.

Selected torque too high for RPM.

High charge temperatures due to inappropriate use of carburinjection systems, or excessive supercharging.

Selection of high power at low RPM with a constant speed prope

Too low a fuel octane rating.



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ls for water contamination:

lowest point in the tanks or fromcts at the bottom of the specimen.

smeared onto the end of a dipstickce of water is indicated by a change

nt impregnated with water detectingntity of fuel is then drawn into theater is indicated by a change in the

ombustion occurs causing all of then in the air. This occurs at a ratio of

the most power for a given RPM. Aan 15:1.



ANSWER 19.

The following methods are employed in testing piston engine fue

Visual inspection. A small quantity of fuel is drained from thedrain points and inspected visually for signs of water which colle

Water finding paste. A small quantity of water finding paste iswhich is then dipped to the bottom of the fuel tank. The presenin the colour of the paste.

Water detecting capsules. A capsule incorporating a paper elemechemicals is placed onto the end of a syringe. A specified quasyringe causing it to pass through the capsule. The presence of wcolour of the paper element in the capsule.

ANSWER 20.

The chemically correct air fuel ratio is that at which complete ccarbon and hydrogen in the fuel to combine with all of the oxygeapproximately 15:1.

The best power air fuel ratio is that at which the engine produceslean (weak) mixture is one in which the air fuel ratio is greater th

A rich mixture is one in which the air fuel ratio is less than 15:1.



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ey tend to reduce fuel consumption. The main disadvantage of weakres aid engine starting and providedis case the excess fuel is not burnedecause the excess fuel is not burned

ouling of spark plugs.

required to produce a unit of powernsumption varies with mixture in a

d in order to provide excess fuel toer output and hence specific fuel

mixture is employed with all of theo contribute to power output. Thens.

cause the flame rate is lower in thesures and hence power output. The

, gradually reducing to a minimumery high values as mixture strength



ANSWER 21.

The principal advantage of employing weak mixtures is that thrate and hence increase flight endurance for a given fuel loadmixtures is the increased probability of detonation. Rich mixtuadditional cooling when operating at high power settings. In thbut by absorbing latent heat in evaporating it cools the engine. Bfuel consumption is increased. Rich mixtures can also result in f

ANSWER 22.

The specific fuel consumption of an engine is the quantity of fuelfor a unit of time. It is measured in lbf/HP hour. Specific fuel cocomplex way. At high power settings a rich mixture is requirecool the engine. This excess fuel contributes nothing to powconsumption is high under such conditions.

For economical cruising flight at lower power settings, a weakercarbon and hydrogen in the fuel being combined with oxygen tlowest specific fuel consumption is achieved under these conditio

At still leaner mixtures the engine runs at lower temperatures beweak mixtures. This lower flame rate also reduces cylinder presoverall effect of this is an increase in specific fuel consumption.

Because of the above factors SFC is high at low fuel:air ratiosvalue at a fuel:air ratio of about 0.066, before increasing to vincreases.



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n of the best economy mixture for ratio is reduced from full rich EGTixture. Further leaning beyond thiss selected by carefully adjusting the

tion blades mounted upon a centrale engine crankshaft. The blades arelong their length. Due to its fixed

normal in which it is driven by thethe engine after engine failure or inropeller is that it can be designed toinefficient throughout most of the

constant across the entire propellerus from root to tip. This means that at the tip than at the root, with theilst all others would vary from very. Propeller blades are twisted from length.



ANSWER 23.

Because cylinder temperature varies with fuel:air ratio, selectiocruise flight can be carried out using EGT as a guide. As mixturewill gradually increase attaining a maximum at best economy mpoint will result in a reduction in EGT. Best economy mixture imixture to locate the maximum EGT point.

ANSWER 24.

The fixed pitch propeller comprises of two or more aerofoil sechub. In modern light aircraft it is usually attached directly to thtwisted from root to tip to provide a uniform angle of attack apitch this type of propeller has only two modes of operation, engine, and windmilling in which the airflow causes it to drive flight shut down. The principal disadvantage of the fixed pitch poperate at one on RPM/airspeed combination and is hence operating range.

ANSWER 25.

Although the inflow velocity due to forward speed is more-or-lessdisc, the rotational velocity increases as a direct function of radithe angle of attack of an untwisted blade would be much highereffect that only one part would be at its most efficient angle, whinefficient due to low angle of attack, to approaching the stallroot to tip to provide a uniform angle of attack along their entire



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t can operate efficiently at only onees inefficiently due to either too lowattack at low airspeeds. A furtherdmilling following in-flight engine

tically in flight to achieve optimumght engine failure it can be feathered

s of a piston, usually located in theontrolled by varying the ratio of oilsystem oil pressure acts on only one spring. If the propeller is part of a on the piston is determined by aelow the set value cause blade pitchease. In this way the drag on thed increases reducing angle of attackte. This is sensed by the centrifugalack, RPM, and propeller efficiency.



ANSWER 26.

The principal disadvantage of the fixed pitch propeller is that iRPM/airspeed combination. Under all other conditions it operatan angle of attack at high airspeeds, or too high an angle of disadvantage is that it generates considerable drag when winfailure.

In variable pitch propellers the blade angle can be varied automaefficiency at all power settings and airspeeds. In the case of in-fliin order to prevent windmilling and minimise drag.

ANSWER 27.

In variable pitch propeller systems blade angle is altered by meanspinner. In the case of the double acting system, blade angle is cpressures acting on both sides of the piston. In the single acting side, with the other being subjected to the effects of a powerfulconstant speed system the magnitude of the oil pressure actingcentrifugal governor sensing propeller RPM. Decreases in RPM bto increase and increases above selected RPM cause it to decrpropeller is varied to maintain a constant RPM. Also as airspeereduces drag on the propeller blades tend to make them acceleragovernor which increases the blade angle to restore angle of attThe reverse process occurs in the case of a reduction in airspeed.



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n the spinner. The flow of servo oill governor sensing propeller RPM.ing cause blade pitch to increase orRPM. Also as airspeed increases,eller causing it to accelerate. This isgle to restore angle of attack, RPM,f a reduction in airspeed. Changes

the centrifugal governor in order to

ined further to include a facility forected as master and the speed of theal governors.



ANSWER 28.

Propeller angle is altered by means of a piston, usually located ipressure to and from the piston is determined by a centrifugaVariations in RPM caused by changes in altitude or power settdecrease varying propeller drag forces to maintain a constant reducing propeller blade angle of attack reduces drag on the propsensed by the centrifugal governor which increases the blade anand propeller efficiency. The reverse process occurs in the case oin RPM selection are achieved by adjusting the spring loading onestablish the required datum RPM.

In the case of multi-engined aircraft this system is sometimes refpropeller RPM synchronisation. In this case one propeller is selothers is matched to it by varying spring loading in the centrifug



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n the spinner. The flow of servo oill governor sensing propeller RPM.ing cause blade pitch to increase orRPM. Also as airspeed increases,eller causing it to accelerate. This isgle to restore angle of attack, RPM,f a reduction in airspeed. Changes

the centrifugal governor in order toilure, selection of feather position through coarse pitch to the featherby a torque switch sensing loss of



ANSWER 29.

Propeller angle is altered by means of a piston, usually located ipressure to and from the piston is determined by a centrifugaVariations in RPM caused by changes in altitude or power settdecrease varying propeller drag forces to maintain a constant reducing propeller blade angle of attack reduces drag on the propsensed by the centrifugal governor which increases the blade anand propeller efficiency. The reverse process occurs in the case oin RPM selection are achieved by adjusting the spring loading onestablish the required datum RPM. In the event of engine faoverrides the centrifugal governor causing the piston to be drivenposition. Alternatively in autofeather systems this is initiated engine torque.



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her the propeller in the event of in-initiated by a torque switch sensingdirecting oil pressure from an auto-r valve directing oil pressure to theeather position.

revent single acting variable pitchre reduces to zero after engine shut

ion after shut down this is likely tog subsequent attempts to start thender the action of springs locks thees after start up centrifugal force

ellers of a multi-engined aircraft areharmonic vibrations. The term

s taken a step further whereby inllers are matched.



ANSWER 30.

The purpose of the auto-feather system is to automatically featflight engine failure or shut down. The autofeather sequence is the loss of engine torque. The system then operates a solenoid feather pump to the constant speed unit. This lifts the selectocoarse pitch side of the piston, thereby driving the blades to the f

ANSWER 31.

The purpose of the low pitch stop (centrifugal latch) is to ppropellers from moving to the feather position as servo oil pressudown. If the propeller is permitted to move to the feather positimpose unacceptable torque loading due to drag forces durinengine. The mechanism comprises a centrifugal latch which, ublade angle actuating piston at low RPM. As RPM increasovercomes spring force and disengages the latch.

ANSWER 32.

The term "Synchronising" describes a system in which the propall made to run at the same speed in order to minimise "Synchrophasing" describes a system in which this process iaddition to RPM matching, the angular position of all the prope



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ed using some form of tachometer the other propellers is altered untilse of synchrophasing, one propellerpeller is set as the master for thatf each master blade are sensed. Thes until both RPM and master blade

of a fixed pitch propeller vary withy of such propellers increases fromfore decreasing rapidly with furthercy is very narrow and the efficiency, variable pitch propellers maintainas the effect of achieving maximumnd, and reducing the efficiency lapse

que and RPM. The most efficienth high torque. Conversely pistonue. The purpose of the reduction

e to their most efficient torque/RPM



ANSWER 33.

In the synchronising system the RPM of each propeller is senssystem. One propeller is selected as the master and the speed ofall the tachometer system speed signals are the same. In the cablade is set as the master and an individual blade in each propropeller. The RPM of each propeller and the angular position osystem then alters the speed of the non-master or slave propellerpositions for all propellers are the equal.

ANSWER 34.

Because its blade angle is fixed, the angle of attack of the bladesRPM and aircraft forward speed. At a given RPM the efficienczero when the aircraft is stationary, to some maximum value, beincreases of forward speed. The speed band of maximum efficienlapse rate is very steep above this band. By varying blade pitchoptimum angle of attack over a much wider speed range. This hefficiency at lower airspeeds, maintaining it over a wide speed barate above this band.

ANSWER 35.

The efficiency of the propeller and engine each vary with torcombination for propellers is relatively low RPM coupled witengines are most efficient at high RPM and relatively low torqgearing is to enable both the engine and propeller to operate closranges.



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ot with an indication of the amounte RPM and throttle selections for arying the oil bleed from a number of to engine torque.

aircraft then propeller is liable to bemage can affect the propeller in aefficiency of the propeller. Changesly damage in the load bearing areasatastrophic failure. To prevent theseor any signs of such damage before

hrottle to 1/2 open, mixture to full. After start-up release the ignitiontained within 30 seconds.

ant speed system there is a danger ofg or stalling the engine unless theer is to enrich the mixture, increase

ing power is to retard the throttle,



ANSWER 36.

The torque meter in a propeller system is fitted to provide the pilof torque being applied by the engine. This is used in setting thgiven power output or flight condition. The system works by vacylinders such that the resulting pressure is directly proportional

ANSWER 37.

Because of its high RPM and exposed position at the front of thedamaged by impacts with stones and similar debris. Such danumber of ways. Changes in blade cross-section will reduce the in balance will cause vibration, particularly at high RPM. Finalof the blades is likely to result in fatigue damage and ultimately cproblems it is essential that propellers be thoroughly examined fengine start up and after shut down.

ANSWER 38.

Engine start-up. Set master switch and booster pump to on, trich, propeller RPM to full increase, and ignition switch to startswitch to the both position and check normal oil pressures are at

Power adjustments. When making power adjustments in a constapplying excessive torque to the propeller shaft and overheatincorrect sequences are employed. The sequence for increasing powRPM setting, increase throttle setting. The sequence for reducreduce RPM, then adjust the mixture.



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idle for a short period to allowne is shut down by setting the fuel switching off the ignition systems.

ted by the ability of its components, and velocities. All engines sustainindividual engine depends upon thertional to the power setting and then at any given power setting. The the duration and magnitude of the

t a specified lifetime, manufacturersy be used continuously whilst othershorised to operate continuously at a at some higher power settings for. The time limits set for these higheremains within acceptable limits andhe engine.



Shut-down. Before shut-down the engine must be run at temperatures to stabilise below 400 degrees Celsius. The engimixture to idle cut-off and allowing the engine to stop beforeCowl flaps must be left open until the engine has cooled.

ANSWER 39.

The amount of power that can be generated by an engine is limito withstand the resulting temperatures, pressures, accelerationsdamage continuously when running and the ultimate life of an cumulative effects of this damage. The rate of damage is propocumulative damage is proportional to the duration of operatiouseful life of a given engine is therefore inversely proportional topower settings employed.

In order to ensure that engines can be operated safely throughouauthorise a number of different power settings, some of these maare subject to time limits. An aircraft engine will typically be autcertain "maximum continuous" or "rated" power output, andlimited periods during manoeuvres such as take-off and climbingpower settings are calculated to ensure that cumulative damage rcatastrophic failure is avoided throughout the authorised life of t



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ied life, manufacturers authorise a time limits. The time limits set fors will reach their authorised servicepower setting at which an engine is



ANSWER 40.

In order to ensure that engines remain serviceable for a specifnumber of different power settings, some of which are subject tothese higher power settings are calculated to ensure that enginelife without failure. Rated, or continuous power is the highest authorised to operate without time limits throughout its life.





Gas Turbine Principles of Operation

Introduction

Gas Turbine Working Cycle

Engine Efficiencies

Engine Developments



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f Operation

t propulsion unit, having first flownorld War and only truly beginning

50's.

through the engine and a propellingller (turbo-prop), a large fan (turbo-mples of aircraft using the first twors Viscount. Most present day large

aw of motion which states that forhe amount of forward thrust createde rearward acceleration imparted to

ass of air is given a relatively smallf air is subject to a much greater



19Gas Turbine Principles o

Introduction1. The gas turbine engine is a relatively new form of aircrafsuccessfully as an experimental power plant during the Second Wto replace the piston engine in commercial aircraft in the early 19

2. The gas turbine may be used to accelerate a mass of gas nozzle, in which case it is known as a turbo-jet, to drive a propefan), or to drive the rotor of a helicopter (turbo-shaft). Early exapropulsion methods were the De Havilland Comet and the Vicketransport aircraft use turbo-fan engines.

3. All aircraft propulsion systems employ Newton’s third levery force (or action) there is an equal and opposite reaction. Tis proportional to the product of the mass of air affected and thit. In the case of propeller driven aircraft and turbo-fans a large mrearward acceleration. In turbo-jets a much smaller mass oacceleration.



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producing movement is the rotary jets produces a force (F = ma), orhich the jets are mounted, to rotate.

FIGUReacti(Garde



Jet Propulsion4. One of the best known examples of jet reaction thrustgarden sprinkler. The acceleration of a mass of water from theaction. It is the reaction to this force that causes the arm, upon wThis is illustrated at Figure 19-1.

RE 19-1on Thrust n Sprinkler)



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pulsive thrust to drive the rocket is to the force produced propels the

FIGUReacti(Rocke

ch as missile propulsion) is the ram a moving vehicle, air is forced intoes the total energy of the air and theing nozzle. The force resulting fromust forwards.



5. A rocket is another example of jet propulsion. The proobtained by accelerating a mass of gas rearwards. The reactionrocket forwards. This principle is illustrated at Figure 19-2.

RE 19-2on Thrust t)

6. Another type of jet engine that has aeronautical uses (sujet, illustrated at Figure 19-3. Provided the engine is attached tothe air intake at the front of the engine. Burning fuel then increasexpanding heated gas is accelerated rearwards through a propellthe acceleration of a mass of gas rearwards produces reaction thr



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FIGURamje

ncluding thermal energy in the forma function of velocity. The kineticg in the direction of gas flow. If noese energies remains constant. Thechanged with, for example dynamicn the velocity of the gas is reduced.



RE 19-3t

7. Any mass of gas possesses energy in a number of forms iof its temperature, static pressure, and kinetic energy which is energy of a gas is evident in the form of dynamic pressure actinenergy is added to or removed from the gas then the sum of thdistribution of energy between the three forms can however be pressure being exchanged for static pressure and temperature whe



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ure 19-4) its velocity decreases totion of the duct. This causes thetatic pressure and temperature withssing through a convergent duct to maintain mass flow causing anreductions in static pressure and

s of energy, together with those ofth of which involve the introductiontering the intake passes through a of its dynamic pressure into staticy the combustion of fuel thereby energy gas is then ejected through a the static pressure and temperaturection to the acceleration of the gasrking cycle the engine must first bew. Such engines therefore have no



8. When such a gas passes through a divergent duct (Figmaintain constant mass flow through the increasing cross-secdynamic pressure of the gas to be exchanged for higher levels of sthe overall energy content remaining constant. When pa(Figure 19-5), this process is reversed, with increasing velocityincrease in dynamic pressure coupled with compensating temperature.

9. The operation jet engines is based upon such exchangecombustion and mechanical compression using moving ducts, boof additional energy to the gas. In the ram jet engine, air endivergent duct. The resulting velocity reduction converts somepressure and temperature. Thermal energy is then added bincreasing the temperature and total energy or the gas. The highconvergent duct where the increase in velocity converts some ofinto dynamic pressure. The thrust of the engine being the reathrough the convergent duct. In order to initiate the ram jet woaccelerated to a high velocity to provide the required inlet air flopractical use in current commercial aircraft.



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FIGUFlow tDiverg

FIGUFlow tConve



RE 19-4hrough ent Duct

RE 19-5hrough rgent Duct



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gh a divergent duct and part of it ismbustion increases the total energy.ropelling nozzle) where some of itsy increase or acceleration). The rame it needs initial high velocity before

hrust when the aircraft is stationaryf inducing an air mass flow into thelete the energy conversion/transfer means of an engine driven fan, orthe expanding hot gas is acceleratedowered by the expanding hot gases compressors in the engine. This is



10. In the ram jet engine, air entering the intake passes throuconverted to static pressure energy. Adding heat energy from coThe expanding hot gas passes through a convergent duct (the pheat and static pressure energy are converted to dynamic (velocitjet has little practical application as an aircraft power plant, sincit can develop thrust.

11. A practical aircraft engine must be capable of producing tand at low forward speeds. To achieve this it must be capable ointake, even when the aircraft is stationary, in order to compprocesses described previously. The inflow of air is induced bycompressor. The air is the heated by the combustion of fuel and out of the rear of the engine by a propelling nozzle. Turbines pproduced by combustion are currently employed to power theillustrated at Figure 19-6.



Chapte

FIGUBasic G

s that of a piston engine - induction,e 19-7.



RE 19-6as Turbine

Gas Turbine Working Cycle12. The working cycle of a gas turbine is basically the same acompression, combustion power and exhaust, as shown at Figur



Chapte

FIGUGas TuWorki

n takes place at constant pressure,, and the processes of the cycle aretraction process links places both inrocesses are sequential and power iscesses are continuous and so greater



RE 19-7rbine ng Cycle

13. The major differences are that in jet engines combustiowhereas in the piston engine it takes place at constant volumecontinuous in the gas turbine. Often in jet engines the power exthe turbine and the propelling nozzle. In the piston engine the ponly produced during the power stoke. In the gas turbine the propower can be produced for a given size of engine.



Chapte

mber is not an enclosed space and, air during combustion. Instead, the attached compressor to rotate. Theeplace that expanding through the

occur in the combustion and turbine temperatures coupled with high

rate very close to the limit of theirately limit the maximum operating

tion chamber materials, the greatern, the greater the air mass flow. The the propelling nozzle, and it is thisir mass flow, and the more it is

take place through the gas turbineon, pressure energy is added to the, with the addition of heat energy.he total energy of the gas into useful for propulsive thrust, takes place in



14. It will be seen from Figure 19-7 that the combustion chaconsequently, there is no pressure increase due to heating of thehot air expands through the turbine, causing the turbine and itsrotating compressor draws air in through the air intake to rcombustion chamber.

15. As indicated in Figure 19-7, the highest gas temperatures sections of the engine. In the case of the turbines, these highcentrifugal loading mean that the turbine discs and blades opephysical capabilities. It is therefore these materials that ultimtemperature and power output of the engine.

16. Within the temperature limits of the turbine and combusthe heat, the greater the expansion, the faster the speed of rotatioair mass flow through the engine is given a large acceleration byacceleration that produces reaction thrust. The greater the aaccelerated, the greater the reaction thrust.

17. The changes in pressure, velocity and temperature that engine are shown graphically at Figure 19-7. During compressiintake air. Total energy is further increased during combustionThrough the turbine expansion results in conversion of some of twork, the remainder of the energy conversion, into kinetic energythe propelling nozzle.



Chapte

rgy of the turbine exhaust gas from convergent duct, or nozzle is used,

l speed of sound). Where supersonic jet engine, a convergent/divergent

nversion of energy. Such a nozzle is

FIGUConveDivergPropel

input. One of the most importantrust produced, divided by the fueltion. The major factors that affect

ficiency.

eveloped by the propelling nozzle toof expressing this is:



18. The function of the propelling nozzle is to convert the enepressure and heat energy into velocity (kinetic) energy. For this aso long as the gas velocity does not exceed sonic speed (the locagas velocities are encountered, as is the case in a rocket or ramnozzle, or venturi, is necessary in order to obtain maximum coillustrated at Figure 19-8

RE 19-8rgent / ent ling Nozzle

Engine Efficiencies19. The efficiency of any engine is the ratio of output to measures of the efficiency of a jet engine is the amount of thconsumption. This is known as the thrust specific fuel consumpspecific fuel consumption are propulsive efficiency and overall ef

20. Propulsive efficiency is the ratio of the amount of thrust dthe energy supplied to the nozzle in a usable form. Another way



Chapte

es of gas turbine engine is shown at

FIGUPropulEfficien

ft in the Exhaust Gas-----------------------------------------------



21. A comparison of the propulsive efficiencies of various typFigure 19-9.

RE 19-9sive cies

Propulsive Efficiency =

Propulsive Work Done on AircraPropulsive Work Done in Aircraft + Work Wasted-------------------------------------------------------------------------------------------------------------------------



Chapte

duced by the engine in usable formbination of combustion efficiency,nd so on. It is dependent upon the is to convert the energy in the fuelc energy).

al energy output of the engine to theine inlet temperature increases. The airspeed, due to ram effect at the efficiency of a jet engine is 20% -the thermal efficiency of the piston lower than that of the jet engine at

echanically in a number of forms -known as single-spool, twin-spool,

rious mechanical arrangements aree-driven compressor may be of theers) or of the axial flow type.

ulsive Efficiency



22. Overall efficiency is the ratio of the amount of energy proto the total amount of energy available in the fuel. It is a comthermal efficiency, mechanical efficiency, compressor efficiency aefficiency of each of the working cycle processes, whose functioninto a form that the propelling nozzle can turn into thrust (kineti

23. Thermal efficiency is defined as the ratio of the mechanicheat energy available in the fuel consumed. It increases as the turbthermal efficiency of a jet engine also increases with increasedcompressor inlet. Under static sea level conditions the thermal25%, compared to 25% - 30% for a piston engine. However, engine decreases with increasing airspeed, becoming significantlyhigher airspeeds.

Engine Developments24. As previously mentioned, jet engines may be arranged mturbojet, turbo-prop, turbo-shaft, turbo-fan. They may also be triple-spool and high or low by-pass ratio engines. These vaexplained further in the following text. In addition, the turbincentrifugal type (already encountered in piston engine supercharg

Overall Efficiency Thermal Efficiency Prop×=



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Royce Nene type. Air mass flow isanding gas from the combustionompressor. The turbine exhaust isust. Air mass flow in such an enginecompressor, limiting the propulsive

FIGUEarly Je



Single-Spool Engines25. Figure 19-10 shows an early turbo-jet engine of the Rollsprovided by a double only centrifugal compressor. Hot, expchambers powers the single stage turbine, which drives the caccelerated through the propelling nozzle to produce reaction thris limited by the low compression ratio inherent in centrifugal thrust available.

RE 19-10t Engine



Chapte

rough them is parallel to the axis oflow. Hence much greater propulsivepool turbo-jet engine with an axialuch an engine

FIGUSingle-Turbo

low about 450 mph, the propeller/ciency than the turbo-jet engine. about 350 mph a marriage of theght ratio, is attractive. Hence thehe jet engine. Figure 19-12 shows aft of the 1950’s and ‘60’s. In turbo-racted by a turbine and transmittedhis process leaves little or no useful of early turboprops a single turbineodern engines a free power turbine,ler.



26. Axial flow compressors, so called because the airflow throtation of the compressor, are capable of a much greater mass fthrust is possible from the engine. Figure 19-11 shows a single sflow compressor. The Rolls Royce Viper is a typical example of s

RE 19-11Spool Axial Jet

27. Reference to Figure 19-9 will show that, at airspeeds beengine power plant arrangement has better propulsive effiConsequently, for aircraft designed to operate at speeds up topropeller to the gas turbine, with its excellent power-to-weidevelopment of the turbo-prop engine from the earliest days of tturbo-propeller engine of the type used on many transport aircrapropeller engines the majority of the energy in the hot gas is extthrough a drive shaft and reduction gearbox to the propeller. Tenergy in the exhaust gas to provide reaction thrust. In the casedrove both the compressor and the propeller whereas in more mtotally independent of the compressor is used to drive the propel



Chapte

FIGUEarly TEngine

orsepower of a turbo-prop engine itis, in turn, depends upon the speed

(pressure rise achieved across thees increases the compression ratio.by having two or more axial flowdriven by its own turbine. This

Figure 19-13.



RE 19-12urbo Prop

Twin-Spool Engine28. In order to increase the thrust of a turbo-jet or the shaft his necessary to increase the air mass flow through the engine. Thof rotation of the compressor and the compression ratio compressor stages). Increasing the number of compression stagFor reasons to be discussed later, this is usually best achieved compressors arranged in series (one after the other), each arrangement is shown in the turbo-propeller engine illustrated at



Chapte

FIGUTwin-SFlow TEngine

rs a high-pressure turbine, which ising the HP spool. Exhaust gas from

e entry, or low-pressure compressor,peller through reduction gearing. It the LP spool. The two spools areh faster than the LP, because of thegement is known as a twin spool



RE 19-13pool Axial urbo Prop

29. Hot expanding gas from the combustion chambers powemounted on the same shaft as the high-pressure compressor, formthe HP turbine powers a second turbine, on the same shaft as thforming the LP spool. This rotating assembly also drives the prowill be noted that the HP spool is hollow and concentric withcompletely mechanically independent. The HP spool rotates mucgreater energy in the gas powering the HP turbine. This arranengine.



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FIGUTwin-SJet Eng

e mechanical arrangement is veryriven by the LP spool. Instead, therated by the propelling nozzle, givesat required for combustion, some of the combustion chambers, and theeam of the propelling nozzle.

the air is not heated and expanded)re thrust than the single-spool purens the engine makes less jet noise.unheated (cold) air, which bypassesmbustion, is known as the by-pass by-pass ratio



RE 19-14pool Turbo ine

30. Figure 19-14 shows a twin-spool turbo-jet engine. Thsimilar to that in Figure 19-13, except there is no propeller dincreased air mass flow from the duplicated compressors, acceleincreased thrust. Because the air mass flow is well in excess of ththe output of the LP compressor by-passes the HP compressor,turbines to mix with the hot gas stream in the exhaust, just upstr

31. In such an engine the lower gas velocity (because some ofis offset by the much greater mass flow, to create significantly mojet engine. Furthermore, the lower exit velocity of the gas meaEngines of this type are known as by-pass engines. The ratio of the combustion process and turbine, to (hot) air heated by coratio. In this case the ratio is less than 2:1 and is known as a low



Chapte

FIGUTwin-SShaft E

in which there is a third rotatingo an output gearbox. Such a turbine Turbo-Prop engine is an example ofugh there are three shafts, there are



RE 19-15pool Turbo ngine

32. Figure 19-15 shows a twin spool turbo-shaft engine assembly comprising a turbine driving a power shaft connected tis known as a free, or power, turbine. The Rolls Royce Advancedsuch an arrangement. It is still a twin spool engine because, althoonly two turbine/compressor spools.



Chapte

FIGUTriple-Ducte

e. The three spools are arrangedor, the LP turbine driving the LPngine the mass of air accelerated bygh the engine. With virtually no tiple, the thrust produced by the fan is the same shaft horsepower. Hence,t of the low by-pass engine and isgure 19-9).



Triple-Spool EnginesRE 19-16Shaft d Fan Engine

33. Figure 19-16 shows a triple spool ducted fan enginconcentrically with the HP turbine driving the HP compresscompressor and a final turbine driving a ducted fan. In such an ethe fan is at least five times greater than the mass passing throulosses, and acceleration of the fan air through a propelling nozzmuch greater than would be produced by a propeller driven withthe propulsive efficiency of this arrangement is better than thamaintained to higher airspeed that that of the turbo-prop (see Fi





Gas Turbine Construction Part 1 – The Cold Section

Introduction

Air Intakes

Pitot Type Intakes

Multi-Shock and Variable Area Intakes

Hazards and Factors Affecting Intake Efficiency

Compressors

Compressor Stall and Surge

Stall and Surge Prevention Devices

Pilot Actions to Prevent and Correct Compressor Stall and Surge



tion

Chapte

n Part 1 –

ich, together, make up the turbinetatic assembly of the engine - the aird those which make up the rotating

ssor, combustion chamber, turbine,

art 1 covers the cold section, whichrt 2 covers the hot section, whichelling nozzle. Part 3 considers theffects of air bleed and the auxiliary

Gas Turbine Construction Part 1 – The Cold Sec


20Gas Turbine ConstructioThe Cold Section

Introduction1. This study considers in detail the major components whengine. These may be sub-divided into those which make up the sintake, combustion chambers, exhaust and propelling nozzle anassembly of the engine - compressors and turbines.

2. Put into a sequential order these are; air intake, compreexhaust and propelling nozzle.

3. The subject is divided into three parts for convenience. Pcomprises the air intake, the compressor and the diffuser. Paincludes the combustion chambers, turbine, exhaust and propworking cycle of the engine, thrust reversal and augmentation, egearbox.



tion

Chapte

frame component. Nevertheless, it air intake are threefold. Firstly it ism to the engine. Secondly, it should

to the compressor at the maximumthis delivered pressure should be ast contribute as little as possible to

efore act as diffusers, decreasing theith supersonic aircraft the problem

d at subsonic speeds, to create anergent duct is needed for the sameetry.

he pitot type that is typically used inn stub wings on the side of the rear is reduced but this can result inat high power setting on the ground.t doors around the periphery of thehe engines demand for high airflow pressure in the intake causing the



Air Intakes4. Normally the air inlet duct is considered to be an aircontributes to efficient engine performance. The functions of therequired to admit the maximum amount of air from the free streadiffuse the airflow, with minimum loss, so as to deliver the air total pressure with minimum disturbance (turbulence). Ideally, close as possible to total free stream pressure. Thirdly it musaerodynamic drag.

5. In subsonic aircraft the inlet ducts are divergent and thervelocity and thus raising the static pressure of the incoming air. Wis complicated by the fact that a divergent passage is requireincrease in static pressure, whereas at supersonic speeds a conveffect. Inlet ducts on supersonic aircraft often have variable geom

Pitot Type Intakes6. The optimum design of air intake for subsonic aircraft is tconjunction with podded engines mounted on wing pylons or ofuselage. By minimising the diameter of such intakes, dragunacceptable restrictions to airflow to the engines when running This problem can be overcome by the use of secondary air inleintake. At high power settings when stationary on the ground tcauses airflow to accelerate into the intake. This reduces staticsecondary doors to open to increase inlet area.



tion

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a divergent duct, converting some ofis causes a reduction in air velocity

endency for shock waves to form at

akesge require a type of air intake thatpermit some of the excess air due toe compressor inlet. Control of theke, and spilling of excess air can behese secondary doors often performent the airflow to the compressorse 20-1 shows a variable area intake



7. At low to moderate (sub-sonic) speeds, the intake acts as the dynamic pressure of the incoming air into static pressure. Thand an increase in air temperature.

8. At high subsonic and transonic speeds there is a marked tthe lip of such intakes, seriously reducing their efficiency.

Multi-Shock and Variable Area Int9. Aircraft that operate in the transonic and supersonic ranreduce the shock waves and, in the case of high Mach numbers, ram effect to be spilled back to atmosphere before it reaches thshock waves can be achieved with the use of a variable area intaachieved with secondary inlet doors fitted in the intake casing. Ttwo functions, acting as a scoop in subsonic conditions to augmand as a spill valve in supersonic flight to dump excess air. Figurwith secondary doors of the type fitted to Concorde.



tion

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FIGUVariabIntake



RE 20-1le Area



tion

Chapte

he shock wave in such a way as tohilst maintaining the total pressurern military supersonic aircraft usecontrolled in all stages of flight toor variable geometry intakes.

e known as an external/internald the desired effect by producing ale strong shock wave at the lip.

FIGUExternCompIntake



10. The purpose of the variable area intake is to position tdecrease the velocity of the airflow at the compressor inlet, wwithin the duct as close as possible to ambient. Many modespecially designed intake ducts with secondary doors that are achieve shock wave and duct pressure control without the need f

11. Some early supersonic aircraft used a type of intakcompression intake, as illustrated at Figure 20-2. This achieveseries of mild shock waves in the intake throat, rather than a sing

RE 20-2al/Internal ression



tion

Chapte

d on the side of the fuselage there isnds on the ground, or if the aircraftbecomes partially obscured by thecent to, the wing root, but can alsoe disruption of the airflow can be

intake can have serious effects onive inlet area and duct shape. Thereage if significant amounts of ice arenecting vanes are invariably fitted

at this system (which normally usest a de-icing system, since use of theion. If, due to pilot inattention, de-d be de-iced one at a time with thetion.

ake Efficiencyspecially at the lip, will disrupt the velocity, resulting in loss of intake At transonic and supersonic speedsus disruption of the airflow into the



12. Airflow Separation. Where the engine intakes are mountea danger of airflow separation when operating in strong crosswiis allowed to sideslip in flight. This is because the air intake fuselage. It is especially the case with intakes buried in, or adjaoccur with rear fuselage mounted engines. In extreme cases thsufficient to cause engine flameout.

13. Inlet Icing. The formation of ice on and in the air compressor and engine performance, due to changes in the effectis also a considerable risk of flame-out and of compressor damingested. Engine inlet ducts, nose cowls and structural interconwith thermal anti-ice protection. It is important to appreciate thhot bleed air from a late compressor stage) is an anti-ice and nosystem for the latter purpose would inevitably lead to ice ingesticing of the engine intakes becomes necessary, the engines shouligniters switched ON to avoid flame-out in the event of ice inges

Hazards and Factors Affecting Int14. Inlet Damage. Distortion or damage to the air intake, eairflow into the inlet and adversely effect the duct pressure andefficiency and possibly disturbed airflow at the compressor entry.such damage will disrupt the shock wave pattern and cause seriointake duct.



tion

Chapte

ay give rise to airflow separation ate compressor inlet, possibly to thehe engine is a possibility in such is a wise precaution to select engine

icularly the compressor or fan, arects. The carefully profiled fan orts, such as stones from the groundake duct, are drawn into the enginer damage to the highly stressed fan

e cracks grow from minor nicks andrly impractical to mount protectivesure that no loose material can enter

nce of jet/propeller wash from other during ground operations. Suchwer settings on the ground.

e the following functions:



15. Heavy In-Flight Turbulence. Flight in heavy turbulence mthe intake and cause significant disruption of the airflow to thextent of causing compressor stall or surge. Flameout of tcircumstances. If flight in heavy turbulence cannot be avoided itignition on (continuous).

16. Foreign Object Ingestion. Gas turbine engines, and parthighly prone to damage through the ingestion of foreign objecompressor blades are easily damaged or broken if solid objecsurface or carelessly discarded tools or loose material in the intwhen it is rotating at high speed. More importantly, even minoand compressor blades is likely to lead to fatigue failure as fatiguscratches. From the foregoing text on intake design, it is cleamesh screens in the intake, and so great care must be taken to enthe highly vulnerable intake.

17. Careful pre-flight inspection of the intakes and the avoidataxiing aircraft will help to prevent foreign object ingestionprecautions are particularly important when operating at high po

Compressors18. Compressors in gas turbine engines are designed to achiev

(a) Increase the air mass flow.

(b) Improve combustion characteristics.



tion

Chapte

ine.

rgy of the air received from the inleter in the right quantity and at theto compress it adiabatically, and soessure.

ends upon compressor rpm, there, density and temperature and thef its outlet static pressure to its inlet

rs, and two types are currently inompressors (straight through flow).mparting kinetic energy to the air inet of divergent flow passages. Of themately 60 per cent is needed to drive pumps, etc).

nents, as shown at Figure 20-3.



(c) Increase the efficiency of the operating cycle.

(d) Increase the thrust produced by the engine.

(e) Improve fuel economy.

(f) Assist in the provision of a small and compact eng

19. The purpose of the compressor is to increase the total eneduct, compress it and discharge it into the combustion chambrequired pressure. In the compressor work is done upon the air the temperature of the air increases in direct proportion to the pr

20. The amount of air passing through the engine depatmospheric conditions at the engine inlet, such as the air pressuaircraft speed. The pressure ratio of a compressor is the ratio ostatic pressure.

21. Most gas turbines use continuous flow rotary compressouse. They are centrifugal compressors (radial flow), and axial cBoth types of compressor work on the same general principle of ia high speed rotor and converting this energy into pressure in a senergy available in the gas after the combustion process, approxithe compressor and the engine accessories (generators, hydraulic

Centrifugal Compressors22. The centrifugal compressor consists of three main compo



tion

Chapte

FIGUCentriComp

s in air and discharges it throughmounted in an outer casing, which chambers through a manifold. Thesided as shown in Figure 20-5.

wing the airflow through it, is at. Centrifugal force causes the air inn outwards to the impeller tip, orler to replace that thrown outwards.nt passages, which cause a rise in



RE 20-3fugal ressor

23. The rotating assembly is called the impeller, which drawstationary divergent passages called a diffuser. The diffuser is collects the diffuser discharge and delivers it to the combustionimpeller may be single sided, as shown in Figure 20-3 or double

24. A simplified diagram of a centrifugal compressor, shoFigure 20-4. The impeller is rotated at high speed by the turbinethe radial passages formed by the impeller vanes to be throwcircumference. Air is drawn into the centre, or ‘eye’, of the impelThe radiating vanes on the face of the impeller form divergepressure as air is thrown outwards through them.



tion

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FIGUSimplifCentriComp



RE 20-4ied fugal ressor



tion

Chapte

ent passages (diffuser ducts) in the air, due to its velocity, into pressuresor is shared approximately equallyet pressure to inlet pressure) for aompression process is an increase inty.

and casing its velocity increases dueent passages formed by the impellervelocity decreases and its pressure the combustion chambers.

entry or double entry centrifugal claimed to give the best all-rounditude. The double entry compressorss frontal area. This is an advantageer tends to be preheated as it passes as the engine designer is concerned. the other) are used, the Rolls Royce



25. On leaving the impeller the air enters stationary divergcasing of the compressor. These convert the kinetic energy of theenergy. The total pressure rise achieved by a centrifugal compresbetween impeller and diffuser. Typical compression ratio (outlcentrifugal compressor is about 4:1. The overall effect of the cstatic pressure and temperature with little or no change in veloci

26. As air flows through the radial passages between impellerto centrifugal action and its pressure increases due to the divergvanes. As the air flows through the stationary diffusers its increases due to diffusion. From the diffusers the air is ducted to

27. The engine designer has the choice of either single compressors (see Figure 20-5). The single entry compressor isefficiency with less risk of surging (to be discussed shortly) at altachieves greater mass air flow for less diameter, and therefore leto the airframe designer, but the air to the rear side of the impellthe discharge from the front side and this is a disadvantage as farIn some engines two single sided compressors in series (one afterDart is an example.



tion

Chapte

it will double the mass flow of aironly one side of the compressor it isressor does not therefore produce a

igher pressure ratios are requiredhe first is then directed through thee product of those of all the stages.chieve approximately 16:1 if used in



28. Where a double entry centrifugal compressor is employedentering the engine. Because each particle of air passes through subject to only one stage of compression. The double entry comphigher pressure ratio than a single entry compressor. If hcompressors are arranged in series so that air passing through tsecond. The overall pressure ratio of such an arrangement is thTwo compressors each of 4:1 pressure ratio for example would aseries.



tion

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FIGUDoublCentriComp



RE 20-5e Entry fugal ressor



tion

Chapte

vantages that have rendered them a very few cases. The principal

herefore low air mass flow.

e is too great the pressure losses willg results.

ages, which have resulted in theirs (APU) and helicopter turbo-shaft

damage.

ure rise).

al compressors.

ngitudinally. Consequently they are a multi-spool engine to reduce its



29. Centrifugal compressors suffer from a number of disadunsuitable for use in aircraft propulsion engines in all butdisadvantages of centrifugal compressors are:

(a) Low compression ratio (typically about 4:1) and t

(b) Larger frontal area (increased aerodynamic drag).

(c) Impeller tip clearance is critical. If the tip clearancbe excessive, if it is too small aerodynamic buffetin

30. However, centrifugal compressors have certain advantcontinued use in smaller engines, such as auxiliary power unitengines. The advantages of centrifugal compressors are:

(a) They are more robust and less prone to ingestion

(b) They are less prone to stall and surge (lower press

(c) They are much less expensive to produce than axi

(d) They are shorter and therefore occupy less space losometimes used as the final stage compressor inoverall length.



tion

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upon which aerofoil section bladese turbine, the aerofoil blades inducef an aircraft induces airflow. This



Axial Compressors31. An axial flow compressor consists of a rotating assemblyare mounted. As the compressor rotor is spun at high speed by thairflow in the same way that a propeller blade or the wing oprinciple is illustrated at Figure 20-6.



tion

Chapte

FIGUSimplifFlow C

asing so that the induced airflow isn at Figure 20-7.



RE 20-6ied Axial-ompressor

32. The rotating assembly is mounted within a stationary cducted to the combustion chambers, turbine, and so on, as show



tion

Chapte

FIGUAxial FCompInduce(One S

shown at Figure 20-8, increases the



RE 20-7low ressor d Flow tage)

33. Adding further rows, or stages, of compressor blades, as air pressure ratio.



tion

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FIGUAxial FCompInduce(Multi

the optimum angle of attack, inlet to the stationary casing (stator) of

the blades. This is illustrated at in this Chapter.



RE 20-8low ressor d Flow Stage)

34. In order to ensure that air enters the rotating blades atguide vanes are often fitted in the air intake. These are attachedthe compressor and angled to direct the incoming air intoFigure 20-9. The operation of inlet guide vanes is discussed later



tion

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FIGURotor of Atta



RE 20-9Blade Angle ck



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Chapte

FIGUAxial-FCompStator

ity) energy to the air. This is partlyadjacent blades is slightly divergent.s. The air passage between these is kinetic (velocity) energy is furthercity that occur across a multi-stage



RE 20-10low

ressor Vanes

35. The rotor blades of the compressor impart kinetic (velocconverted into pressure energy because the air passage between From the rotating blades the airflow enters the stationary vanealso slightly divergent and so a further pressure rise occurs asconverted into pressure energy. The changes in pressure and veloaxial flow compressor are illustrated at Figure 20-11.



tion

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FIGUPressuChangAxial-FComp

se the amount of diffusion has to beless, the pressure rise is cumulative

hole of a multi-stage compressor is will be up to 600°C



RE 20-11re / Velocity es through low

ressor

#

36. The pressure ratio across each stage is less than 2:1 becaulimited to avoid aerodynamic stalling of the blades. Neverthestage-by-stage and a compression ratio of up to 35:1 across the wnot unusual. At such a compression ratio the outlet temperature



tion

Chapte

blades of equal length in all stages. as velocity energy is converted tothe stages of such a machine thererease. Hence, it would be difficult toure gradient.

keep the air flowing through theonvergent. Consequently, the rotorom front to rear of an axial flowage.

ressure gradient is to flow from theustrates the changes in air pressureat the air is required to flow against of higher pressure. For this flow toamic pressure acting downstream isd the air mass flow and compressorny factor that restricts the airflow

se pressure gradient through it, orse the airflow to break down. Suchessor surge.



37. The simple axial compressor shown at Figure 20-10 hasAs already explained, a pressure rise occurs across each stage,pressure energy in the stator (diffuser) vanes. Thus, across all would be an overall velocity decrease and an overall pressure incmaintain airflow against the increasing pressure, or adverse press

38. In order to maintain uniform axial velocity, and thuscompressor from front to rear, the axial air passage is made cblades and stator vanes must be made progressively shorter frcompressor, in order to be accommodated in the convergent pass

Compressor Stall and Surge

Compressor Stall39. The natural tendency of any fluid when acted upon by a phigh pressure area to the lower pressure area. Figure 20-11 illthrough an axial flow compressor and from this it can be seen than adverse pressure gradient, from an area of low pressure to onebe possible it is necessary that the sum of static pressure and dyngreater than that acting upstream. This will be the case provideRPM are within the design specification for the compressor. Ainto the compressor, causes an excessive increase in the adverresults in an unacceptable mass flow/RPM combination will caubreak downs take one of two forms, compressor stall and compr



tion

Chapte

rofoils in that the airflow over themn their stalling angle is exceeded. Install. Stall is brought about by anygroup of blades at an angle greatera single blade or stage or a group ofotal breakdown of airflow is termed12.

e the following:

ircraft manoeuvring.

ure caused by rapid opening of the



40. The blades of axial flow compressors are like all other aevaries with angle of attack, and will detach and break down whethe case of compressors this phenomenon is termed compressor factor that causes the airflow to approach any single blade or than the stalling angle of those blades. Stall can therefore affect blades or stages. When the entire compressor stalls the resulting ta compressor surge. Compressor stall is illustrated in Figure 20-

41. Factors that are likely to promote compressor stall includ

(a) Low engine RPM during starting and idling.

(b) Strong crosswinds on the ground.

(c) Engine inlet icing.

(d) Contaminated or damaged compressor blades.

(e) Damaged air intake.

(f) High angle of attack at the air intake caused by a

(g) Excessive increases in combustion chamber pressthrottle particularly at low engine RPM.



tion

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FIGUCompAerod



RE 20-12ressor ynamic Stall



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the phenomenon in which a totalome cases a surge also results inses the result is a loss of thrust oftenoises and rapidly increasing EGT.l into one of three categories, all ofe gradient through the compressor.ssor, unacceptable RPM/mass flowtion chamber pressure.

ith RPM and air mass flow, the rate fixed and can be perfectly matchedside this combination, inefficienciesressed to match the reducing cross-uses the volume of the airflow to be

to choke. In effect the air is too bigUnder these circumstances airflowpressor to suffer aerodynamic stall.xacerbating the choking of the rear



Compressor Surge42. The term compressor surge is generally taken to refer tobreakdown of airflow through the compressor occurs. In sintermittent reversal of airflow through the compressor. In all cacoupled with vibration, loud banging, rumbling or popping nAlthough a great many scenarios might result in surge, most falwhich involve the creation of an unacceptably adverse pressurThe three common causes are aerodynamic stall of the comprecombination at low RPM, and unacceptable increases in combus

43. Whilst the efficiency of the compression process varies wof reduction of cross-sectional area of the compressor annulus isto only one RPM/mass flow combination. When operating outoccur due to the air being either over compressed or under compsection. When operating at low RPM inadequate compression careduced insufficiently, causing the rear section of the compressorto pass through the casing at the intended uniform velocity. breaks down causing the blades in the upstream stages of the comThis further decreases the efficiency of the compression process estages.



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flow there will be a pressure ratiod for all practicable combinations ite. An example of a surge envelopeh and low RPM is illustrated inhich surge will occur. The workingg in any steady state combination ofeen the working line and the surgerge.

l operating conditions and the surgeting conditions is set. In order toing fuel flow to the burners. Thisses the turbine and compressor toustion chamber pressure rise occursoves closer to the surge line beforeints A, B and C in Figure 20-13. Inl be crossed initiating a compressoreratures and loss of RPM and massine without pilot intervention. Inther than recovering as indicated byin between the working line and then starting or accelerating from lowevention devices are employed.



44. For any given combination of engine RPM and air massabove which the compressor will surge. If these points are plotteis possible to produce a map of the surge envelope of the enginmap indicating the effects of rapid throttle opening at higFigure 20-13. The surge line indicates the pressure ratio above wline indicates the point at which the engine will lie when operatinRPM, air mass flow, and compression ratio. The distance betwline at any given RPM is a measure of the engines resistance to su

45. In order to provide an adequate margin between normaline a working line defining normal engine steady state operaincrease the speed of the engine the throttle is opened increasincreases combustion chamber pressure which eventually cauaccelerate to a higher point on the working. Because the combbefore the increase in RPM and mass flow, the engine initially mreturning to the working line at a higher RPM as indicated by pothe case of excessively rapid throttle opening, the surge line wilsurge. If the surge is sufficiently severe or prolonged, rising tempflow will prevent the engine from returning to the working lextreme cases the engine will move deeper into the surge area rapoints D, E, F and G in diagram Figure 20-14. Because the margsurge line increases with RPM, surge is particularly likely wheRPM. To overcome this problem a number of stall and surge pr



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FIGUSurge Map InEffectsAcceleHigh R



RE 20-13Envelope dicating the of ration at PM



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FIGUSurge Map InEffectsAcceleLow R



RE 20-14Envelope cicating the of ration at PM



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ted by any or all of the following

on in mass flow caused by stall andhe magnitude of this effect depends

f their effect on the airflow throughduction of loud popping, knocking engine.

sruption of airflow and intermittentengine RPM.

upled with continuing combustion

in the ejection of exhaust gasses out



Indications of Stall and Surge46. The existence of a stall and/or surge condition is indicasymptoms:

(a) Unexpected loss of thrust. Because of the reductisurge the thrust output of the engine is reduced. Tupon the severity of the stall/surge.

(b) Abnormal engine noises and vibrations. Because othe engine stall and surge usually result in the proor rumbling noises and vibrations from within the

(c) Uncommanded variations in engine RPM. The direversal of airflow often cause rapid variations in

(d) The breakdown of airflow through the engine coresults in rapidly increasing EGT.

(e) In extreme cases the reversal of airflow can result of the air intake.



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s

ure that the airflow approaches therating conditions. One method ofble angle inlet guide vanes (VIGVs). By varying the angle of the bladescted onto the first stage rotor bladesded to include variable angle statorVs and the general arrangement ofand Figure 20-16.



Stall and Surge Prevention Device

Variable Inlet Guide Vanes47. In order to prevent compressor stall it is necessary to ensblades at an angle lower than the stalling angle under all opeachieving this under conditions of low RPM is to employ variasituated immediately in front of the first stage of the compressorto match prevailing mass flow and RPM conditions the air is direat an acceptable angle. In many engines this technique is extenblades in several stages of the compressor. The effect of VIGVIGVs and variable angle stators are illustrated in Figure 20-15



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FIGUEffect Angle Vanes PrevenComp



RE 20-15of Variable Inlet Guide in ting ressor Stall



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FIGUVariabStatorAxial FComp

the resulting surge are often relieved situated in the compressor casingey are open allowing excess air tor stages of the compressor. As RPMugh the engine and improve enginein Figure 20-17.



RE 20-16le Incidence Vanes in low ressor

Anti-Surge Bleed Valves48. The effects of low compressor efficiency at low RPM and or prevented by means of anti-surge bleed valves. These aretowards the mid-section of the compressor. At low RPM thexhaust to atmosphere thereby preventing the choking of the reaincreases they automatically close off to increase mass flow throefficiency. The general arrangement of these valves is illustrated



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FIGUSurge

dependently to ensure that each isure spools tend to be relatively smallkly to increases in throttle angle,revent the creation of unacceptably

iate pressure spools. In this way theduced (but not eliminated) in such



RE 20-17Bleed Valves

49. In multi-spool engines the speed of spools can vary inmatched to the prevailing mass flow conditions. Also high pressand because of their low inertia, are able to react very quicaccelerating rapidly to the required higher RPM. This tends to phigh adverse pressure gradients in the low pressure and intermedneed for variable inlet guide vanes and anti-surge valves is reengines.



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ect Compressor

l are as follows:

oved before attempting to start the

re starting to avoid intake airflow

when taxiing to avoid ingestion of

whenever possible.

rt up before attempting to accelerate

to avoid or minimise rapid throttle

for all stages of flight

event of compressor stall or surge:

ccur during engine acceleration.



Pilot Actions to Prevent and CorrStall and Surge50. The principal actions to be taken by pilots to prevent stal

(a) Ensure that all engine blanks and covers are remengine.

(b) Ensure that the aircraft is facing into wind befodisturbance and the ingestion of exhaust gas.

(c) Avoid following too close behind other aircraft exhaust gas.

(d) Avoid air turbulence and harsh flight manoeuvres

(e) Allow the engine to stabilise at idle RPM after stato higher speeds.

(f) Anticipate the need for power changes in order movements.

(g) Ensure that the appropriate power setting is used

51. The following actions should be taken by the pilot in the

(a) Reduce the rate of throttle opening if stall/surge o



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oon as other safety considerations

occur during starting.

s safe to do so if stall/surge occur

the rows of blades are mounted isft is coupled to the turbine shaft to

tator vanes are mounted, is usuallyoint. The rotor shaft bearings are,

th the blades attached to the outert Figure 20-18. The discs form an

the centrifugal loads imposed by the



(b) If stall/surge persists reduce throttle setting as spermit.

(c) Terminate the start cycle immediately if stall/surge

(d) Reduce flying control deflections as soon as it iduring flight manoeuvres.

Axial Compressor Construction52. The compressor rotating assembly, or rotor, upon whichsupported in a number of ball and roller bearings. The rotor shaform a single spool.

53. The compressor cylindrical casing in which the rows of smade in two halves with a horizontal longitudinal centreline jtypically, at either end of the stator.

54. The rotor is usually made up of a number of discs, wiperiphery of the discs. This method of construction is shown aintegral drum of low weight, but sufficient strength to withstand blades at high rotational speeds



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FIGURotor Const

uch that the stagger angle measuredm root to tip. This is intended toconstant along the blade, whilst the attack of an untwisted blade wouldto exhibit inefficiently low angles ofck at the tip. Secondly the rotorsg to deflect it outwards towards thereate a pressure gradient along the to provide this span-wise pressure axial airflow velocities across the

llustrated in Figure 20-19.



RE 20-18Disc ruction

55. The rotor blades are of aerofoil section and are twisted sbetween the blade chord and the axis of rotation increases froachieve two effects. Firstly because the axial airflow velocity is airflow velocity due to rotation increase with radius, the angle ofincrease from root to tip. Such a blade would therefore be liable attack at the root, coupled with excessively high angles of attaimpart a swirl to the airflow generating a centrifugal force tendintips of the blades. To counteract this effect it is necessary to clength of the blades. The rate of twist of the blades is designedgradient whilst ensuring acceptable angles of attack and uniformrotor disc. The varying airflow velocities and stagger angle are i



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FIGUCompTwist



RE 20-19ressor Blade



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hat the load imparted to the disc issor blades are often free to rock or concentration at the root due to clicking noise often heard when theg with the engine shut down on the



56. The rotor blades are secured to the discs in such a way tminimised. Examples are shown at Figure 20-20. The compresslide slightly in their root mounting in order to relieve stresscentrifugal force. This “looseness” gives rise to the characteristiccompressor rotates at very low rpm, such as when it is windmillinground.



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FIGURotor StaggeAttach



RE 20-20Blade r ment



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red in annular grooves in the casing.duce any tendency to vibrate. Theduces losses by monitoring leakageown at Figure 20-21.

FIGUShroudVanes



57. The stator blades are also of aerofoil section and are secuThe blades may be joined together at their inner ends, to reshrouding also forms a smooth surface along the annulus and rearound the blade tips. An example of shrouded stator vanes is sh

RE 20-21ed Stator



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ade of titanium, which has a verythe high centrifugal forces whilst

or nickel based alloys, an importanttched’ by foreign object ingestion

ines have a tendency to bend andn support (known as a ‘snubber’ orblades are mounted on their rotor.erspeed or surge these snubbers may

at is to say engines with no by-passxpanded to power the turbine and

he first, or LP compressor creates anpressor. The excess airflow passes

haust gas, to be accelerated in thesely matched to the aircraft speedr fuel consumption, than the single-ern aircraft. An example of a twin-



58. Compressor rotor blades, discs and drums are usually mhigh strength-to-density ratio and is thus able to withstand maintaining low weight. Stator blades are usually made of steel feature being the need to retain high strength even when ‘nodamage (FOD).

59. The very long blades used in high by-pass ratio fan eng‘flap’ under high aerodynamic loads. To prevent this, a mid-spa‘clapper’) is used, forming a strengthening ring when the fan Under the effects of shock, such as FOD ingestion, bird strike, ovoverlap each other - a condition known as ‘shingling’.

Multi-Spool Compressors60. Single spool compressors are used in pure jet engines, thair, in which all the air mass flow is heated by combustion, eaccelerated by the propelling nozzle.

61. Low by-pass engines are of the twin-spool type in which tair mass flow greater than that required by the second HP comaround the HP compressor and mixes with the hot turbine expropelling nozzle. The result is that the jet velocity is more clorequirements, giving much greater propulsive efficiency, and lowespool pure jet. The latter type of engine is rarely found in modspool compressor is shown at Figure 20-22.



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FIGUTwin-SComp



RE 20-22pool Axial ressor



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ge further. The single or multi-stagets output to the core of the engine,ucted to a ‘cold’ propelling nozzle tosses, typically, through two furthernd exhaust system. In a triple-spoolriple-spool compressor is shown at



62. The high by-pass turbo-fan engine takes this concept a stafan at the front of the engine delivers only about one-fifth of iknown as the gas generator. The remainder of the fan output is dproduce most of the engine thrust. The ‘core’ airflow then pacompressors before entering the combustion chambers, turbines aengine the rearmost turbine drives the fan. An example of a tFigure 20-23.



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FIGUTriple-Comp



RE 20-23Spool ressor



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wing advantages over single spool

each is able to operate at the RPMd pressure ratios thereby increasingressor stall and surge.

a small proportion of the total air. The resulting low inertia enables it accelerating to the required higherreatly reduces the probability ofcreases.

w passes through the HP spool, andound the outside, friction losses are and mechanical efficiency of the

sely matched to that of the aircraft,te airspeeds.

user casing, which forms a divergenttion chamber at a low velocity, by increasing static pressure.



63. Twin-spool and triple-spool compressors offer the follosystems.

(a) Because the spools are not physically connected which best match the prevailing air mass flow anoverall efficiency and reducing the danger of comp

(b) Because the HP spool is required to handle onlyflow, its diameter and mass can be greatly reducedto react very quickly to throttle increases rapidlyRPM, mass flow, and pressure ratio. This gcompressor stall and surge during rapid throttle in

(c) Because only a small proportion of the air mass flothe remainder passes at a much lower velocity argreatly reduced thereby improving the thermalengine.

(d) Because the velocity of the by-pass air is more clopropulsive efficiency is improved at low to modera

Diffusers64. On exit from the compressor the air passes through a diffpassage. This casing prepares the air for entry into the combusconverting some of its kinetic energy into pressure energy further




Part 2 – The


Gas Turbine ConstructionHot Section

Introduction

Combustion Chambers

Fuel Burners

Turbines

Multi-Spool Engines

Effects of Damage

Gas Temperature Monitoring

Exhaust and Jet Nozzles

Noise reduction

Thrust Reversal





Thrust Augmentation Systems

Air Bleeds and Auxillary Drives



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n Part 2 –

hat are located in what is known as the turbine, and the jet pipe, whichrbine outlet and propelling nozzle.

where fuel is burned to increase the

essure, velocity and airflow pattern

r efficient combustion.

with the remainder of the airflow to

Gas Turbine Construction Part 2 – The Hot Sect


21Gas Turbine ConstructioThe Hot Section

Introduction1. This section deals with those major engine components tthe hot section. These are the combustion chamber or chambers,comprises the propelling nozzle and the exhaust pipe between tu

Combustion Chambers2. The combustion chamber (or burner section) is a region temperature of the gas (air), adding heat energy.

3. The purposes of the combustion chamber are as follows:

(a) To provide a suitable environment in terms of prfor efficient sustained combustion.

(b) To mix fuel and air in the required proportions fo

(c) To mix the hot gasses resulting from combustion produce a uniform temperature gas flow.



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ide vanes at the optimum velocity,s.

nt combustion:

n and/or vaporisation.

front which is typically between 60

Airflow from the compressor outlet divided. Some is directed throughombustion liner, and the remainder chamber and the flame tube. This mixing with the combustion gasesipes from the flame tube serve to

onditions in each. The corrugated



(d) To deliver the hot gasses to the turbine nozzle gupressure and temperature with the minimum losse

4. The following conditions are required for sustained efficie

(a) An air:fuel ratio within the range 8:1 to 30:1.

(b) Thorough mixing of the fuel and air by atomisatio

(c) An airflow velocity not exceeding that of the flameand 80 feet per second.

5. Figure 21-1 shows an early type of combustion chamber. enters through the snout of the chamber from where the flow isswirl vanes to the combustion zone within the flame tube, or cflows along an annular passage between the outer casing of theannular airflow assists with heat insulation, whilst progressivelythrough holes in the wall of the flame tube. Interconnector pconnect adjacent combustion chambers and maintain similar cjoints in the flame tube permit expansion.



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FIGUEarly CChambRoyce



RE 21-1ombustion er (Rolls

)



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s a ratio to fuel flow, may vary fromoptimum combustion ratio of 15:1.all proportion of the airflow (up tor. Here it will mix with the fuel at arimary air. Downstream of this areaning 80% (called the secondary air)

region of 1800°C to 2000°C, thezzle guide vanes of the turbine cantion gases to between 1000°C andg determined by the materials fromBecause of steady diffusion throughpressure both fall slightly across thepical combustion chamber is shown

FIGUCombChambSplit



6. The total airflow through the hot section of the engine, a45:1 to as much as 130:1. This is obviously far too high for the The combustion chamber is therefore designed to introduce a smabout 20%) into the primary or combustion zone of the chamberatio of about 15:1. This proportion of the airflow is called the pof the combustion chamber is the dilution zone, where the remaimixes with the combustion products.

7. Since the temperature of the combustion gases is in thesecondary air serves to cool the gases to a level which the nowithstand. The temperature of the gas is raised by the combus1500°C or even higher, the limit of acceptable temperature beinwhich the turbine guide vanes and first stage blades are made. the chamber, and virtually unimpeded exit, the gas velocity and combustion chamber. The airflow and flame stabilisation in a tyat Figure 21-2.

RE 21-2ustion er Airflow



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er second, so the velocity of the air be kept low, otherwise the burning air velocity from the compressor tocombustion chamber reduces it stilltable burning and so the airflow is

f the combustion chamber, primaryw. This combines with air enteringculatory airflow called a torroidalilising and anchoring the flame, see



8. Kerosene burns at a relatively slow rate of 60 to 80 feet pentering the combustion zone of the combustion chamber mustfuel will simply be blown away. The diffuser section reduces thea lower value and further diffusion in the entry section of the further, to about 60 feet per second. This is still too high for sgiven a swirling motion to further reduce its axial velocity.

9. Immediately downstream of the snout or entry section oair passes through swirl vanes, which impart a rotational airflothrough the primary air holes to produce a low velocity recirvortex (something like a smoke ring) which has the effect of stabFigure 21-3.



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FIGUFlame

le, to ensure efficient burning of theht over a wide operating range. For

a weak limits to the fuel/air ratio, will be extinguished. Flame-out ise engines at idle power, since in thisng in a very weak mixture. A graph1-4.



RE 21-3Stabilisation

10. It is important that the combustion process remains stabfuel/air mixture, and to further ensure that the flame remains aligany particular type of combustion chamber there are rich andbeyond which combustion cannot be sustained and the flameperhaps most likely to occur during a high speed descent with thsituation there is a high mass airflow and a low fuel flow, resultiillustrating the limits of combustion stability is shown at Figure 2



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FIGUCombStabilit

opposed to combustion) of the fuelcombustion than it is to maintain it.igure 21-4.



RE 21-4ustion y

11. There are weak and rich limits beyond which ignition (aswill not be achieved, since it is always more difficult to establish These limits are within the stable region shown in the graph at F



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ng to the power plant requirementse described below.

layout, typical of centrifugal anders, like that in Figure 21-1 andted only for purposes of pressure chamber comprises an inner flamerises some 80% of the total airflow

er is its high resistance to distortionial needed for their construction and can-type, or multiple combustion



Types of Combustion Chamber12. Combustion chamber design varies considerably accordiof the aircraft. The main types of combustion chamber layout ar

Multiple or Can-Type Combustion Chambers. The earlyearly axial flow compressor fitted engines. Individual chambFigure 21-2, are disposed around the engine and interconnecequalisation and to assist flame spread during initial start. Eachtube around which there is a cooling air casing. Cooling air compsupplied to the entry section of the chamber (see Figure 21-2).

13. The principal advantage of this type of combustion chambwhen heated. The main disadvantages lie in the amount of matertheir uneconomical use of the available space. The layout ofchambers around the engine is illustrated at Figure 21-5.



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FIGUMultipCombChamb



RE 21-5le ustion ers



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to the second type of combustionn chamber, but the individual flamehis simplifies construction, easinging can be significantly reduced forterials and weight, compared to thellustrated at Figure 21-6.



Tubo-Annular, Cannular or Can-Annular are all names given chamber. The layout is similar to that of the multiple combustiotubes are mounted within a common annular air casing. Tmaintenance and inspection procedures. The diameter of the casthe same mass airflow and there is a considerable saving in mamultiple combustion chamber arrangement. A typical layout is i



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FIGUCannuCombChamb



RE 21-6lar ustion er Layout



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most modern turbo-fan engines. Itnd the engine (forming an ‘annulus’,to admit air from the compressoras to the turbine inlet nozzle guide

he front end. Since the whole of thestion chamber can be shorter for then is both lighter and simpler. Flameal ‘hot spots’ or of flameout. Total less cooling air is required, giving

burned fuel, reducing air pollution at Figure 21-7.



The Annular type of combustion chamber is the type found oncomprises a single combustion chamber annularly disposed arouor sleeve, encompassing the engine). It is open at the front discharge diffusers, and at the rear to deliver hot, expanding gvanes.

14. Fuel spray nozzles are disposed around the annulus near tarea encompassing the engine is used for combustion, the combusame overall diameter as the previous two layouts. Constructiopropagation is more efficient, so there is less likelihood of locsurface area is significantly reduced with the consequence thatgreater combustion efficiency.

15. The annular combustion chamber virtually eliminates unand increasing thermal efficiency. A typical example is illustrated



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FIGUAnnulaCombLayout



RE 21-7r ustion



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bustion chamber is often used in ase of such chambers is particularly



16. In order to reduce overall engine length the annular commodified reverse-flow form as indicated in Figure 21-8. The uprevalent in small turbo-shaft helicopter engines.



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FIGUReversCombChamb



RE 21-8e Flow ustion er



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the great temperature differencesignificant thermal loads upon there. The joint between combustionn and combustion chamber cooling

100% at sea level, reducing to 98%ducing combustion chamber inlet

pply a continuous spray of finely combustible mixture. In order torayed, must be supplied with fuel ated spray pattern. If the fuel supply

an a spray, as shown at Figure 21-9.id jet of fuel.



17. The high temperatures in the combustion chamber, andbetween it and the surrounding parts of the engine, places schamber itself and expansion loads upon the adjacent structuchamber and turbine inlet is often designed to allow for expansiois very carefully designed to minimise thermal loads.

18. The combustion efficiency of most gas turbine engines is at altitude because of reducing static air pressure and retemperature.

Fuel Burners

Burners (Atomisers)19. The purpose of the burners, or fuel injectors, is to suatomised fuel that will readily mix with air to form an easilyachieve this the burner nozzle orifice, through which the fuel is spsufficient pressure and must be designed to produce a cone-shappressure is too low the fuel will be discharged as a film, rather thIf supply pressure is too high the nozzle is liable to produce a sol



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FIGUStagesAtomi



RE 21-9 of Fuel sation



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hich are described below.

simple spray nozzle with a singlelied through tangential holes to a to the fuel assists with atomisationged

FIGUSimple

ortional to the square root of thete the range of fuel flows required0 psi at maximum rpm. Fuel pumpsnance of the required spray patternently, alternative spray nozzles werearrower fuel supply pressure range.



20. There are various types of burner, the principal ones of w

Simplex Burners. The early gas turbine engines used a discharge orifice as illustrated at Figure 21-10. Fuel is suppchamber, before being discharged. The swirling motion impartedand varying its supply pressure varies the quantity of fuel dischar

RE 21-10x Burner

21. Because the flow of a liquid through an orifice is proppressure drop across it, the range of pressures needed to crearanged from about 30 psi for minimum engine rpm to about 300of the time could not produce such high pressures, and mainteproved impossible over such a wide range of pressures. Consequdeveloped in which the fuel flow range could be achieved with a n



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s two nozzle orifices, as illustratedgs), fuel flows only to the primaryply pressure to the burner increases,it fuel to the main nozzle orifices.an be varied over a wide range and

FIGUDuplex



Duplex (Duple) Burners. The Duplex (or Duple) burner haat Figure 21-11. At low supply pressure (small throttle openinorifice. As the throttle valve is progressively opened, and fuel supthe spring-loaded pressurising valve progressively opens to admHence, for a relatively narrow supply pressure range, fuel flow cpressure remains adequate for efficient atomisation

RE 21-11 Burner



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ilar to the simplex nozzle, with theuel can be ‘spilled’ back to the fuelhamber at high pressure, ensuring a) the spill valve is closed and all theinimum engine rpm at high altitude)the swirl chamber is spilled away.g motion to ensure that the small

ised

FIGUSpill Bu

l both supply and spill pressures tohe engine.



Spill Burners 22. A spill burner nozzle is illustrated at Figure 21-12. It is simaddition of a passage from the swirl chamber through which fpump supply at a controlled rate. Fuel is supplied to the swirl cgood spray pattern. At high demand rates (maximum engine rpmsupplied fuel is sprayed from the nozzle. At low demand rates (mthe spill valve is opened so that most of the fuel supplied to However, at the high supply pressure there is sufficient swirlinquantity discharged through the nozzle orifice is efficiently atom

RE 21-12rner

23. The spill burner requires a complex fuel system to contromeet the fuel flow requirements over the full operating range of t



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r spray nozzles. Some of the primaryzle. This aeration of the fuel sprayer fuel supply pressures than other



Air Spray Nozzles. Many modern gas turbine engines use aiair for combustion mixes with the fuel supply in the spray nozachieves more efficient combustion of the fuel and requires lowburner types. An air spray nozzle is illustrated at Figure 21-13.



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FIGUAirspr



RE 21-13ay Nozzle



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energy of the hot gases into shaftrocess is, ideally, one of adiabaticngs. In turbojet engines the turbines (fuel pumps, generators, hydraulicn a turbo-prop or turbo-fan engine,tly much less energy is available toaining is negligible. In turbo-shaft

nes and output shaft.

serves to reduce the pressure andin aircraft gas turbine engines are oftor sets) varies according to engine

es mounted upon a wheel, or disc, the combustion process is directed) causes the turbine wheel to rotate.y passing it through a convergent

turbine blades. A simple turbine is



Turbines24. The function of the turbine is to transform some of thehorsepower (SHP). The expanding gas does work and this pexpansion, in which there is no loss of energy to the surroundiextracts sufficient energy to drive the compressor and accessoriepumps, etc), the remaining energy providing propulsive thrust. Ithe turbine also has to drive the propeller or fan, and consequenprovide thrust in the jet pipe. In many turbo-props the thrust remengines, virtually all of the useful energy is extracted by the turbi

25. In effect, a turbine is a compressor in reverse, since itincrease the velocity of the gas passing through it. Turbines used the axial flow type and the number of turbine stages (rotor/stapower and type.

Single Spool Turbines26. In its simplest form the turbine consists of a ring of bladattached to the rotor (rotating assembly). High velocity gas fromonto the turbine blades and the force exerted (mass x accelerationThe velocity of the combustion gas is increased (accelerated) bpassage, or nozzle, thus increasing the force it exerts upon theillustrated is Figure 21-14.



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FIGUSingle-Turbin

mbly is illustrated at Figure 21-15.



RE 21-14Spool e

27. The terminology associated with the turbine rotating asse



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FIGUSingle Turbin

ing the combustion gases, by forcingximately sonic velocity. The kineticpon impact with the turbine blades,

needed to drive the compressor.

ion gases is achieved by a ring ofgine casing between the combustionch pair of adjacent vanes forms aelerated gas onto the turbine blades



RE 21-15-Stage e Rotor

28. The energy transfer in the turbine is achieved by acceleratthem to pass through a convergent passage (a nozzle), to approenergy added to the gas during the velocity increase is absorbed ucausing the turbine to rotate at high speed, providing the power

29. In a practical gas turbine the nozzling of the combuststationary, aerofoil shaped nozzle guide vanes attached to the enchamber outlet and the turbine inlet. The passage between eaconvergent duct, and the vanes are angled so as to direct the accat the appropriate angle.



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e, reaction, and combined impulse/nes form convergent ducts through in the direction of rotation of the

the space between adjacent blades is to a change of direction whilst itsdirection equates to a change in

hat drives the turbine. The impulse



30. Three forms of turbine are possible. These are impulsreaction. In the pure impulse type of turbine the nozzle guide vawhich the direction of the gas flow is accelerated and deflectedengine. The rotor blades are of bucket shaped section such that constant. Gas passing between the blades is therefore subjectedspeed and static pressure remain constant. This change of momentum and it is the impulse of this change of momentum ttype turbine is illustrated in Figure 21-16.



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FIGUImpulsBlades



RE 21-16e Turbine



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ed. The cross sectional area of theassing through them is subject to aonstant. The rotor blades are ofergent ducts. Gas passing throughture to reduce. It is the reaction to-17.



31. In the reaction type turbine this series of events is reversspaces between the nozzle guide vanes is constant and so gas pchange of direct whilst its speed and static pressure remain caerofoil section such that the spaces between them forms convthese ducts is accelerated causing its static pressure and temperathis acceleration that drives the turbine as illustrated at Figure 21



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FIGUReactiBlades



RE 21-17on Turbine



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n systems are inefficient and so mostulse/reaction turbine (Figure 21-18)ket section, gradually changing toe vanes this change of cross section

the roots of the vanes, whilst thelose to the tips. In the rotors thetype energy extraction at the roots. Gas passing through the rotors isrted by the preceding nozzle guidebine. The general gas flow throughe changing cross section of turbine



32. When employed in isolation both the impulse and reactiopractical turbines employ a combination of the two. In the impthe blade root sections are predominantly of impulse-type bucreaction-type aerofoil section towards the tip. In the nozzle guidis reversed such that the greatest acceleration occurs close togreatest deflection in the direction of engine rotation occurs cchanging cross section of the blades results in mainly impulse-gradually changing into a mainly reaction-type process at the tipsaccelerated and deflected such that the rotational velocity impavanes is negated, leaving a more-or-less axial flow out of the turan impulse/reaction turbine is illustrated in Figure 21-18, and thblade is illustrated at Figure 21-23.



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FIGUImpulsTurbin



RE 21-18e / Reaction e



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ngine is to expand the gas, forcing itre achieved during combustion theble in the turbine. However, this iss to withstand high temperatures.

air from the compressors is passedl stresses and gas loads. An examplede vanes are usually made of nickel



Turbine Construction33. The function of the combustion process in a gas turbine einto the nozzle guide vanes. Clearly, the greater the temperatugreater the expansion and the greater the energy transfer possilimited by the ability of the nozzle guide vanes and turbine blade

34. The nozzle guide vanes are usually hollow and cooling through them to prevent overheating and damage due to thermaof this method of cooling is shown at Figure 21-19. Nozzle guialloy.



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FIGU



RE 21-19



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transfer by convection, using bladesg air is passed. The next stage ofozzle guide vanes and rotor blades.ll holes in the surface of the nozzle over the surface of the vanes andoling of the blades through complexlong the length of the blade. (See



35. Early attempts at turbine blade cooling relied upon heat with longitudinal passages through which low pressure coolindevelopment was to introduce external air film cooling of both nIn this, high-pressure air is led through internal passages to smavanes and rotor blades. The escaping air forms a cooling filmblades. This is normally supplemented by internal convection copassages that force the cooling air to make several passes aFigure 21-20).



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FIGUHigh PTurbinCoolin



RE 21-20ressure e Blade g




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FIGUImpingCoolinGuide



RE 21-21ement g of Nozzle Vanes



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tube fed with cooling air. Holes inguide vanes, the air finally escapingvane. This method is known ass have been built with nozzle guide effected by the escape of cooling airnown as transpiration cooling.

mounted are typically forged fromy airflow directed across both sides



36. In some cases the nozzle guide vanes contain an internalthe tube direct air onto the inner surfaces of the hollow nozzle into the gas flow through holes in the trailing edge of the impingement, or jet, cooling. (See Figure 21-21). A few enginevanes made from a very expensive porous material and cooling isthrough the pores from the hollow interior of the vane. This is k

37. The turbine discs, or wheels upon which the blades aresteel and bolted to the rotor shaft. The discs are usually cooled bof the disc. (See Figure 21-22).



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FIGUTurbinand SeArrang



RE 21-22e Cooling aling ements



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d so that the stagger angle increasest the gas flow from the combustiontor blade. It achieves this by causingo tip to counter the fact that turbinehe twist in the turbine rotor bladest to tip so that it enters the exhaustand Figure 21-24.



38. The turbine nozzle guide vanes and rotor blades are twistefrom root to tip. The twist in the nozzle vanes is to ensure thachamber does equal work at all points along the length of the roa gas velocity decrease, and pressure increase, from nozzle root tblade rotational velocity is higher at the tip than at the root. Trestores the gas flow to uniform velocity and pressure from roosystem without creating swirl. This is illustrated at Figure 21-23



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FIGUChangAngle to Tip



RE 21-23e of Stagger from Root



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FIGUIncomand PrUniforRoot t

ts in the engine. The temperature oftrifugal loads due to rotation often at maximum rpm. The method byof great importance. In most enginesigure 21-25 . When the turbine ischined in the wheel, so there are noe attachment.



RE 21-24ing Velocity essure m from o Tip

39. The turbine blades are the most highly stressed componenthe gas flow is sufficient to make them glow red-hot and the cenamount to several tons on each blade when the rotor is turningwhich the blades are attached to the turbine disc is consequently the "fir tree" root design is used and this is illustrated at Fstationary the blade is a loose fit in the peripheral serrations ma"in-built" stresses. During rotation centrifugal loading stiffens th



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y of the gas flow as efficiently as to gas leakage between the tips, ore to a minimum, shrouding is often

on the shrouds leave an extremely

FIGUTurbinShroud



40. Since the function of the blades is to absorb the energpossible it is important that there should be no lost energy dueends, of the blades and the turbine casing. To reduce such leakagfitted to the blade tips, as shown at Figure 21-25. The ‘ridges’small clearance between shroud and casing.

RE 21-25e Blade ing



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mum tip clearance over the wholetrol (ACC). Compressor air is usedto match the radial expansion of thehrouded turbine blade is shown at

l loads imposed upon them, turbinesion. During operation the effects of, a process known as ‘blade creep’,lade tips contact the turbine casing.the material from which the blades

hat it is made up of many crystalscrystals are uniformly aligned thend fatigue, significantly lengtheningonal solidification’.

single crystal of nickel alloy, which modern turbines use blades madef withstanding much greater speeds

by one turbine rotor, which drives may have one or more stages, orengine.



41. A more effective method of maintaining constant minioperating range of the engine is known as Active Clearance Conto cool, and actively control the expansion, of the turbine casing blades. Blade tip shrouding is then unnecessary. A typical un-sFigure 21-23.

42. In addition to the ability to withstand the high centrifugablades must also be resistant to fatigue, thermal shock and corroheat and centrifugal loading cause the blades to grow in lengthwhich has a finite limit after which ‘blade rub’ will occur as the bHence, turbine blades have a finite life, which is determined by are made. Most turbine blades are cast from nickel-based alloys.

43. Microscopic examination of the blade material reveals twhich are normally randomly aligned. By ensuring that the material becomes much more resistant to creep, thermal shock ain-service life. This manufacturing technique is known as ‘directi

44. A more recent advance is to produce each blade from a permits a substantial increase in turbine gas temperature. Somefrom reinforced ceramics, instead of metal, which are capable oand temperatures.

Multi-Spool Engines45. In a single rotor (single spool) engine, power is developedthe compressor and the engine accessories. The turbine rotorturbine wheels, depending upon the power requirements for the



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r three independent turbine rotors respective compressors. Again each

combustion chamber outlet is firstressor and the accessory gearbox. Insecond (LP) turbine. The LP turbineine also drives the propeller.

een the HP and LP, known as the some turbo-prop and turbo-shaftessors is used to drive the propeller layout is shown at Figure 21-26.



46. In a multi-spool engine power is developed by two oconnected by co-axial, mechanically independent shafts to theirturbine rotor may have one or more stages.

47. The first turbine in the sequence (the one into which thedirected) is called the HP turbine and usually drives the HP compa twin-spool unit, exhaust gas from the HP turbine is led to the drives the LP compressor. In twin-spool turbo-props, the LP turb

48. A triple-spool engine has a third turbine situated betwintermediate (IP) turbine, which drives the IP compressor. Inlayouts a free turbine rotor, independent of the HP or LP compror shaft. This is called a free/power turbine engine. A twin-spool



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FIGUTwin-STurbin



RE 21-26pool e



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ue to the variations in rpm betweene rpm of the HP section through ther is connected. Multi-spool engines

ptible to compressor surge and stallir improved power-to-weight ratio turbo-prop and turbo-fan engines

unit (propeller or fan).

sult of ingestion of foreign objects.he high gas temperatures, and the permanent extension of the turbiner which blade rub occurs as alreadyto the forces of the high velocity gas stagger angle, which reduces the

ted, no cracks are acceptable in anyrovided they are well away from theny part of the blades.

aximum acceptable turbine inlet gasth of blade material with increasedis to cause them to fracture at the



49. The LP section runs more slowly than the HP section. Dthe two turbine assemblies, the practice is to control and limit thfuel control unit (FCU). It is also to the HP rotor that the starteare capable of high compression ratios and are much less suscethan single-spool engines, especially during acceleration. Themakes for much-improved specific fuel consumption (sfc). Withstarting is easier because the HP spool does not drive the power

Effects of Damage50. Compressor damage usually occurs to the blades as a reTurbine damage, however, is almost invariably the result of tstresses arising from high rotational velocity. One of the effects isblades, or creep. This is only acceptable within certain limits, aftediscussed. Another effect is for the turbine blades to untwist due flow acting on the heated blades. The result is reduction ofefficiency of the blades.

51. Because of the stresses to which turbine blades are subjecpart of them and small nicks or indentations are only permitted proot area of the blade. Burning or distortion is unacceptable in a

52. Engine manufacturers place very stringent limits on the mtemperature, because of the marked reduction in tensile strengtemperature. The ultimate effect of overheating turbine blades shank, where the root joins the blade, due to centrifugal loading.



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the wheel radially at high velocity,. This is known as an un-containedge to adjacent parts of the aircraftned within the turbine casing, the ruin the engine.

ing of the turbine gas temperatureonitor the turbine inlet temperatureis calculable it is quite sufficient toxhaust gas temperature (EGT), and

le probes usually located at the exit often incorporated with the nozzle



53. Centrifugal force ensures that the fractured blades leaveand possibly with sufficient force to penetrate the turbine casingturbine failure and the departing blades can cause serious damastructure and systems. Even if the fractured blades are contaidamage they cause to the remaining blades is usually sufficient to

Gas Temperature Monitoring54. It will be appreciated from the foregoing that monitor(TGT) is of prime importance. It is not usually practicable to m(TIT), but since the temperature drop across the turbine stages measure the temperature of the gases leaving the turbine, the ecalibrate the display instrument accordingly.

55. Gas temperature measurement is made using thermocoupfrom the HP turbine. In the case of a multi-spool engine they areguide vanes of the successive turbine, as shown at Figure 21-27.



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FIGUThermArrang



RE 21-27ocouple ement



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located at the forward end of the jetially and connected in parallel. Thisd system redundancy in the event ofated at Figure 21-28.

FIGUJet PipThermSystem

gine power changes. In order toanged, but because of the inertia ofengine RPM does not immediately



56. In the case of single spool engines, the thermocouples are pipe. In either case a number of thermocouples are dispersed radhas the double advantage of a temperature averaging system, anfailure of individual thermocouple probes. The principle is illustr

RE 21-28e Multiple ocouple

57. EGT monitoring is of particular importance during enincrease or decrease power the fuel flow to the engine must be chthe rotating assembly (the HP turbine and its compressor) the change.



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ch clearly increases the combustion is a period where the airflow fromerature. This results in a marked’ the increasing airflow reduces the avoid excessive peak EGT.

uces the fuel flow, and combustiont takes time for the compressor to until the RPM has fallen to match

ottle operation to avoid over-rapid

parts. These are the exhaust uniticant distance apart, they are joined

f the rear turbine disc.

gas emerging from the turbine.

to reduce friction losses within the



58. During a power increase the fuel flow is increased, whitemperature. The engine RPM follows relatively slowly, so therethe compressor does not match the increased combustion temptemporary rise in EGT. Eventually, as the compressor ‘spools upEGT. Care must always be taken when advancing the throttle to

59. Similarly, when reducing power closing the throttle redtemperature, but initially the airflow does not decrease since i“spool down”. This results in a marked temporary drop in EGTthe new throttle setting. Again, care must be exercised in thrcooling of the turbine blades and vanes.

Exhaust and Jet Nozzles60. The exhaust system of a typical jet engine comprises twoand the propelling nozzle. In cases where these parts are a signifby a jet pipe. The functions of the exhaust unit are as follows:

(a) To support the turbine rear bearing.

(b) To provide a streamlined fairing for the rear face o

(c) To remove any residual swirling motion from the

(d) To diffuse the exhaust gas, reducing its velocity jetpipe.

61. The functions of the jet pipe are as follows:



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unit to the propelling nozzle with

ing to kinetic energy as much of thee.

s the cross sectional area of theorder to match prevailing mass flowing the pressure ratio capabilities of

tems where appropriate.



(a) To transfer the exhaust gas from the exhaust minimum energy losses.

(b) To house the reheat system if fitted.

(c) To house temperature and pressure sensors.

62. The functions of the propelling nozzle are as follows:

(a) To increase the velocity of the exhaust gas convertresidual pressure energy in the gas as is practicabl

(b) In the case of engines employing reheat systempropelling nozzle must be automatically varied in to maximise gas flow acceleration without exceedthe engine.

(c) To house (and comprise part of) reverse thrust sys



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FIGUBasic ENozzle

essure-sensing probes are positionedre. This pressure may be displayede pressure and displayed to the pilot



RE 21-29xhaust

63. As well as the EGT thermocouple probes, a number of prin the exhaust casing to measure the turbine discharge pressudirectly to the pilot, or alternatively compared with the air intakas engine pressure ratio (EPR).



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s of between 750 and 1200 feet per friction losses, the speed of the gas

en the exhaust cone and the wall ofl function of preventing the exhaust normal to limit the speed of the gaset per second (Mach 0.5 at exhaust

d of the gas relative to the inside of

rates the exhaust gasses before they final acceleration makes a majorximise engine thrust it is necessaryhigh as possible. Provided the gasacceleration through the nozzle ise acceleration is proportional to thef jet pipe pressure is increased the local speed of sound. At this point

creases in jet pipe pressure will notcharged to atmosphere at a pressurethe overall efficiency of the engine..



64. Gas from the turbine enters the exhaust casing at velocitiesecond. Because velocities of this magnitude would produce highflow is reduced by diffusion. The increasing passage area betwethe jet pipe achieves this. The exhaust cone serves the additionagases from flowing across the rear face of the turbine wheel. It isflow at outlet from the exhaust case to a velocity of about 950 fegas temperatures). It should be noted that this refers to the speethe exhaust unit casing.

65. The propelling nozzle is a convergent duct, which accelepass out of the engine to atmosphere. The reaction to thiscontribution to the overall thrust of the engine. In order to mathat the velocity of the gas leaving the propelling nozzle is as velocity is lower than the local speed of sound the rate of proportional to the pressure drop across it so it. That is to say thdifference between jet pipe pressure and ambient pressure. Ivelocity at the throat of the nozzle will increase until it equals thepressure waves can no longer flow upstream and so further inresult in further acceleration. Instead the exhaust gas will be disabove ambient. The gas will then expand wastefully reducing Under these conditions the propelling nozzle is said to be choked



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to increase the local speed of sound.h comprises a convergent inlet ductcreased beyond the choked nozzlepropelling nozzle, the increasing

ncrease in the local speed of sound.tion, causing the gas to continue to

to say the convergent section of thespeed of sound at the throat (the velocity continues to increase as then increase in gas velocity and hencees of gas flowing through a choked



66. In order to increase the gas velocity further it is necessary This is achieved by means of a convergent–divergent nozzle whicattached to a divergent outlet duct. As jet pipe pressure is incondition in an engine employing a convergent-divergent temperature and pressure downstream of the throat causes an iThis restores the relationship between pressure drop and acceleraaccelerate through the divergent section of the nozzle. That is nozzle remains choked with gas velocity equal to the local narrowest part of the nozzle). Downstream of the throat the gaslocal speed of sound increases in that area. The overall effect is athrust. Figure 21-30 illustrates the pressure and velocity changconvergent-divergent propelling nozzle.



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FIGUGas Fla ConvDiverg

ombustion gases from the turbineum amount of thrust. The nozzle isportantly, the velocity of the gas as

al and is specified and calibrated atve a significant effect upon engine

usted, the cold (by-pass) air and thehe two gas streams are combined ins illustrated at Figure 21-31.



RE 21-30ow through ergent - ent Nozzle

67. The propulsive nozzle is designed to expand the hot cdischarge to atmosphere in a manner that will produce the maximan orifice, the size of which determines the pressure and, more imit emerges from the engine. The area of this orifice is quite criticthe time of manufacture, and any subsequent change will haperformance.

68. In the by-pass engine there are two gas streams to be exhahot exhaust gas from the turbines. In low by-pass ratio engines ta mixer unit before passing through the propelling nozzle. This i



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FIGUInternaMixing



RE 21-31l Gas Flow



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hausted separately through hot andsome improvement in thrust if bothgrated, exhaust nozzle. These two

FIGUInternaGas Fl



69. In high by-pass ratio engines the two gas streams are excold co-axial propelling nozzles. However, it is possible to gain streams are subsequently ejected through a common, or inteconcepts are illustrated at Figure 21-32.

RE 21-32l / External

ow Mixing



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e is that caused by the exhaust jet more intense the noise created. It isthan the turbo-jet, since the bulk of

as velocity and various devices havewhen retro-fitted to the engine, are

its frequency. High frequency soundeen found that the more readily ther the frequency of the jet noise.

st gas through a number of radiallyze of the individual gas streams thed is attenuated more rapidly withnce from the aircraft is less intense

ith the corrugated perimeter nozzle.ated, to create what is, effectively, a for one single nozzle. Examples of



Noise reduction70. Much of the unacceptable noise produced by a jet enginvelocity. Broadly speaking, the higher the exhaust jet velocity thelargely for this reason that the turbo-fan engine is much quieter the exhaust is relatively low speed cold air.

71. The pure jet (no by-pass) engine has the highest exhaust gbeen introduced to reduce the noise of the jet exhaust. These, often known as ‘hush-kits’.

72. The distance that sound travels in air is dependent upon is attenuated much more rapidly than low frequency and it has bjet exhaust mixes with the surrounding atmospheric air the highe

73. One method of encouraging mixing is to pass the exhaumounted nozzles instead of one single nozzle. By reducing the sifrequency of the sound is increased. Since high frequency soundistance in air than low frequency, so the sound at a given distathan would be the case with a larger single nozzle.

74. A similar effect to the multiple-tube nozzle is achieved wThe outer perimeter of the propelling nozzle is ‘fluted’, or corrugring of nozzles with the same total outlet area as that necessarymultiple-tube and corrugated nozzles are shown at Figure 21-33.



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FIGUMultipCorruNozzle



RE 21-33le-Tube and gated Exhausts



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, the effects of air bleeding and the

eated a need for alternative andg. A simple and effective way oftion of the exhaust gas stream, thus

thrust operation are similar to take-aximum forward thrust, and is onlyaft forward speed is inadvisable duetall due to turbulent entry air.

e same with all systems. A reverseverse thrust cannot be selected until. Engine thrust cannot be increasedposition.

the event of failure of operatingrward thrust position. Operation of to show when the reversers are



75. This section covers thrust reversal, thrust augmentationauxiliary gearbox.

Thrust Reversal76. The use of high performance jet-engined aircraft crsupplementary methods of retarding the aircraft after landinreducing the landing roll was found to be by reversing the direcusing engine power to decelerate through thrust reversal.

77. Though the rpm and fuel flow conditions during reverse off power, the reverse thrust obtained is not more than 75% of mabout 50% in many cases. Use of full reverse power with no aircrto possible ingestion of debris and the possibility of compressor s

78. The method of reverse thrust selection is essentially ththrust lever is interlocked with the engine thrust lever so that rethe thrust levers have been moved to the rear of the IDLE settingto a high setting unless the reversers are in the full reverse thrust

79. The reversers are designed for fail-safe operation. In hydraulic pressure a mechanical lock holds the reversers in the fothe thrust reverser system is indicated to the flight deck crewunlocked, moving to and subsequently in the reverse position.



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rom the hot gas propelling nozzle isreaction thrust is reversed from thesic types of hot gas stream thrust

ors which, in their normal (forwardmal gas stream to flow through the to block off the normal gas stream direct the hot gas forward, creating

e designed to exert maximum forces has the effect of holding the doors loss of forward thrust. A clamshell



Hot Stream Thrust Reversal80. In pure jet and low by-pass ratio engines the gas stream fredirected to produce a forward component of velocity, so that normal, to oppose aircraft forward motion. There are two bareverser, the clamshell door and the bucket-target.

81. Clamshell Door reversers are pneumatically-operated dothrust) position close the reverse thrust ducts and allow the norexhaust nozzle. When reverse thrust is selected the doors rotateand divert the exhaust gas through ducts to cascade vanes, whicha jet thrust in opposition to aircraft motion.

82. The pneumatic rams that operate the clamshell doors arwhen the doors are in the normal, forward thrust, position. Thifirmly against the reverse duct seals, preventing gas leakage anddoor thrust reverser is illustrated at Figure 21-34.



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FIGUClamsRevers

forward thrust position the bucketnozzle, having no effect upon theactuators move the doors, by means forwards, to create a reversal of jet35.



RE 21-34hells Thrust ers

83. Bucket-Target reversers are hydraulically actuated. In theshaped doors form a concentric tube around the propelling exhausted gas stream. When reverse thrust is selected hydraulic of push rods, to deflect the jet stream from the propelling nozzlethrust. A bucket-target thrust reverser is illustrated at Figure 21-



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FIGUBucketThrust

rtion of the total thrust is producedgines for the cold stream only to bed thrust position the thrust reverser

mpletely obstruct the rearward flowarge air is diverted through cascaderward direction to the airflow as itrust.



RE 21-35-Target Reversers

Cold Stream Thrust Reversal84. In high by-pass ratio turbo-fan engines the greater propoby the cold (fan) air stream. Consequently it is usual in such enreversed when thrust reversal is selected. In the normal, forwardoors block off the gas-reversing cascade vanes.

85. When reverse thrust is selected the doors are moved to coof fan air to the cold stream propelling nozzle. All the fan dischvanes in the walls of the air duct. The cascade vanes impart a foleaves the engine, producing reversal of cold stream propulsive th



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by air motors through a system ofam thrust reverser is illustrated at

FIGUCold SThrust

l landing operation is that its use reduces the possibility of loss ofly useful when operating on a low



86. Cold stream reverser systems are usually operated eitherscrewjacks, or by hydraulic rams and push rods. A cold streFigure 21-36

RE 21-36tream Reversers

87. The main advantage of reverse thrust during a normareduces the amount of wheel braking required and thereforedirectional control during the landing roll. This is particularfriction surface such as a wet or icy runway.



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speeds the engine may ingest its own undesirable, as it tends to disrupt

At very low forward speeds the jet aircraft, obscuring or obliteratingg in crosswind conditions. Also, inment, which is clearly undesirable if

nd de-selection of reverse thrust. Inon of exhaust from a neighbouringwept wing aircraft) occurs at aboutat about that speed. An engine inbout 50 knots, so the pilot needs to

nal to aircraft speed, so it is clearlyever, in aircraft types where reverseelayed until the nose wheel is firmly

adverse effect on the power outputbe reduced if take-off is from a high high. Thrust augmentation systemsmes.



88. The disadvantage of reverse thrust is that at low forward exhaust efflux or that from an adjacent engine. This is highlyintake flow and can lead to surge and other handling problems.efflux may throw up spray, dust or loose snow ahead of theforward vision. These problems are exacerbated when operatinsome aeroplanes selection of reverse thrust creates a pitch up moit occurs whilst the nose wheel is still off the ground.

89. The pilot can reduce these problems by careful selection afour-engine aircraft with wing mounted podded engines, ingestiengine (typically the outers ingest exhaust from the inners in a s80 knots. Consequently, the inners should be throttled back reverse thrust will be liable to start ingesting its own exhaust at abegin throttling back all engines at this speed.

90. The braking effect of reverse thrust is directly proportiodesirable to select it as early as possible in the landing roll. Howthrust selection creates a pitch up moment its selection must be don the ground.

Thrust Augmentation Systems91. An increase in altitude or in ambient temperature has anof a gas turbine. Consequently the take-off power available will altitude aerodrome or one where the ambient air temperature isare used to restore lost thrust when operating from such aerodro



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o increase the take-off performancehaps with a short runway and/or a method most commonly used is toh the engine. This is achieved by

ralised water and methanol, into theeezing properties and provides an

this purpose in order to prevent theer.

ed directly into the compressor inlet

t, cooling increases air density. This thrust. The methanol burns in theature (TIT), which would otherwisey is restored but additional fuel flow

metered from a storage tank by athe compressor first stage disc. Theow between the first stage blades.



92. Alternatively, thrust augmentation systems may be used twhen operating from low altitude and cool aerodromes, but perheavy aeroplane. In commercial aircraft the thrust augmentationincrease the air mass flow by cooling the air flowing througintroducing either de-mineralised water, or a mixture of de-mineairflow. The addition of methanol to the water gives anti-fradditional source of fuel.

93. De-mineralised water (rather than tap water) is used for hot section engine component reacting with chemicals in the wat

94. The water or water-methanol mixture may either be sprayor, more usually, injected into the combustion chamber.

95. When water-methanol is sprayed into the compressor inleincreases the mass of the airflow and thus produces increasedcombustion chamber and thereby maintains turbine inlet temperfall. When water only is sprayed into the compressor inlet, densitis necessary to maintain TIT and, therefore, engine rpm.

96. With compressor inlet injection the coolant is usually control unit which distributes the coolant to radial passages in coolant is forced outwards by centrifugal action to enter the airfl



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creases the mass flow through thequent reduction of pressure and

pe pressure, resulting in additionaluel system to schedule an increasedjected the burning of the methanol minimal.

urbine driven pump discharging thensing unit ensures that coolant can

compressor discharge pressure, and

y injection is that the combustion of to re-schedule the fuel control unitg methanol with the water is that itrted in its own storage tank for use

on take-off, when maximum thrustrust restoration with use of water



97. Injection of the coolant into the combustion chamber inturbine relative to that through the compressor. The consetemperature drop across the turbine gives an increased jet pithrust. When water is injected the reduction in TIT enables the ffuel flow to increase rpm and thrust. When water-methanol is inhelps restore TIT and increased fuel flow is either unnecessary or

98. Combustion chamber injection typically employs an air-tcoolant to water jets incorporated in the fuel spray nozzles. A seonly flow to the spray jets when coolant pump pressure exceedsactuates an indicator in the cockpit.

99. The primary advantage of water-methanol over water-onlthe methanol either reduces or removes entirely the requirementfor thrust augmented take-off. A secondary advantage of mixinacts as antifreeze. Consequently water-methanol can be transpoon subsequent occasions without the risk of it freezing.

100. Note that water or water-methanol injection is only usedis essential. The graph at Figure 21-37 shows the typical thinjection on a turbo-jet engine



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FIGUWaterThrustRestor

heat, or afterburning, to increasen increase in jet pipe temperature,necessary for afterburning, to give aomical method of increasing thrust

Its main advantage is its ability toweight of the engine. A typicaligure 21-38 and Figure 21-39.



RE 21-37 Injection ation Graph

101. Military aircraft (and Concorde) frequently employ remaximum thrust. Fuel is burned in the jet pipe, resulting in apressure, velocity and hence thrust. Variable area jet nozzles are larger area during afterburner operation. This is a highly uneconand the increased jet pipe velocity makes it extremely noisy. increase thrust without significantly increasing the size or afterburning jet pipe and variable area nozzles are illustrated at F



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FIGUPrincipAfterb



RE 21-38le of urning



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FIGUAfterbPipes aPropel



RE 21-39urning Jet nd ling Nozzles



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for the engine's internal air system,ntal control system (air conditioning

e total air mass flow and thereforengine pressure ratio (EPR), which isaking air from the compressors will and therefore reduce the jet pipe.

turbines will also cause an increasell be less.

ally cause a small increase in enginesor. Depending upon the type of fueluel scheduling with a resultant rpm

he use of bleed air during maximum



Air Bleeds and Auxillary Drives

Bleed Air102. Compressed air is bled from the LP and HP compressorsfor engine and airframe anti-icing and for the aircraft environmeand pressurisation).

103. Removal of air from the engine airflow clearly reduces threduces the engine thrust. Thrust is measured and indicated as ethe ratio of jet pipe pressure to LP compressor inlet pressure. Treduce the mass of air entering the hot section of the enginepressure, reducing EPR. The resulting loss of thrust increases SFC

104. The reduced airflow through the combustion system andin turbine gas temperature (TGT), since the degree of cooling wi

105. The reduced air mass flow when air is being bled will usurpm, since the turbine has less work to do in driving the comprescontrol unit, the reduction in EPR may also lead to increased fincrease.

106. The thrust reduction and increased TGT may well limit tpower operations such as take-off and climb-out.



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or, lubricating oil pumps, hydraulice shafts from an external gearbox

rive shaft, connected through beveltypical external gearbox is shown at



Auxiliary Gearbox107. The various engine-driven accessories (fuel pumps, governpump, and electrical generator) are driven by individual drivmounted on the engine casing.

108. The external gearbox receives its drive through a radial dgearing to one of the gas turbine spools, usually the HP spool. A Figure 21-40.



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FIGUJet EngGearb



RE 21-40ine External ox



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ricating oil. To protect the gearboxts, typically generator and hydraulic. Should one of these components

els in the gearbox can occur.



109. Lubrication of the accessory drive gears is by engine lubagainst damage in the event of accessory failure, some drive shafpump drives, incorporate a deliberately weakened shear sectiontend to seize, the ‘shear-neck’ will fail before damage to gearwhe



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hrough a divergent duct?

and turbo prop engine?

ressor?

airspeed?





What does Newton's Second Law of Motion state?

QUESTION 2.

How do velocity, pressure and temperature vary when air flows t

QUESTION 3.

How is thrust generated in a turbo jet engine?

QUESTION 4.

How does air mass flow and acceleration compare in a turbo jet

QUESTION 5.

What form or type of combustion is used in a turbo jet engine?

QUESTION 6.

How and where is static pressure increased in a centrifugal comp

QUESTION 7.

What is total head pressure and how does it vary with increasing



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low compressor:

pressor?

pressor made progressively shorter

n axial flow compressor?

l compressor being too small?

ntrifugal compressor?



QUESTION 8.

Define compression ratio in a gas turbine compressor:

QUESTION 9.

State the purpose of variable angle inlet guide vanes in an axial f

QUESTION 10.

How and where is static pressure increased in an axial flow com

QUESTION 11.

Why are the blades and air passageways in an axial flow comfrom front to rear?

QUESTION 12.

Why are the blades of an axial flow compressor twisted?

QUESTION 13.

How do air velocity and static pressure change in each stage of a

QUESTION 14.

What would be the consequences of tip clearance in a centrifuga

QUESTION 15.

What is the main disadvantage of a double-entry impeller in a ce



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ikely to cause compressor surge?

ratio at which surge occurs:

position:

highest gas pressure occur:

cess in a gas turbine engine?

ccur?



QUESTION 16.

What combination of mass air flow and pressure ratio are most l

QUESTION 17.

What are the outward symptoms of compression surge?

QUESTION 18.

What term is used to describe the values of airflow and pressure

QUESTION 19.

What would be the effect of a surge bleed valve stuck in the open

QUESTION 20.

What is the purpose of surge bleed valves?

QUESTION 21.

Where in a gas turbine engine do the highest gas velocity and the

QUESTION 22.

What term is used to describe the air used in the combustion pro

QUESTION 23.

Where in a gas turbine engine does the highest gas temperature o



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combustion chamber layout?

ficiency in a jet engine?

stion?

combustion chamber outlet in a gas

ine engines?

in some gas turbines:



QUESTION 24.

How and why are individual chambers connected in the multiple

QUESTION 25.

What type of combustion engine achieves greatest combustion ef

QUESTION 26.

What air : fuel ratio is required in a jet engine for efficient combu

QUESTION 27.

What factors limit the permissible maximum temperature at the turbine engine?

QUESTION 28.

What is the function of the turbine nozzle guide vanes?

QUESTION 29.

What type of turbine blades are most commonly used in gas turb

QUESTION 30.

Describe the concept of Active Clearance Control (ACC) as used

QUESTION 31.

What is a free power turbine and where is it used?



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ine?

it passes through the turbine?

tio engine?

ore compare in a turbo-fan engine?

ersers?



QUESTION 32.

What is the main reason for shrouding of turbine blades?

QUESTION 33.

How does the gas cause the turbine to rotate in a gas turbine eng

QUESTION 34.

How do gas, velocity, temperature and static pressure change as

QUESTION 35.

Define the term choked nozzle as applied to exhaust nozzles.

QUESTION 36.

What is the purpose of the exhaust cone in gas turbine engines?

QUESTION 37.

How is the apparent noise reduced in a pure jet or low bypass ra

QUESTION 38.

How does the air flow through the fan and that through engine c

QUESTION 39.

What form of power is used to operate clamshell door thrust rev



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ange of momentum of a body is in which the force acts.

es and its pressure and temperature

ce resulting from the acceleration of

rge acceleration, but to a relatively

constant pressure (for a given thrust



QUESTION 40.

Describe the Bucket type thrust reverse system?

ANSWERS:ANSWER 1.

Newton's second law of motion states that the rate of chproportional to the applied force and takes place in the direction

ANSWER 2.

When a gas flows through a divergent duct its velocity decreasincrease.

ANSWER 3.

To every action there is an equal and opposite reaction. The fora mass of gas rearwards produces reaction thrust forwards.

ANSWER 4.

Compared to a propeller of similar thrust, a jet engine gives a lasmall mass of air.

ANSWER 5.

In the gas turbine cycle of operations combustion takes place at value).



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rise in pressure as air is throwngent passages (diffuser ducts) whichence, pressure rise occurs in both

. Dynamic pressure increases with

utlet pressure to compressor inlet

the first stage rotor blades at an

ausing a pressure increase. The aird so a further pressure rise occurs.



ANSWER 6.

The impeller vanes form a divergent passage, which causes aoutwards. On leaving the impeller the air enters stationary diverconvert the kinetic energy of the air into pressure energy. Himpeller and diffuser.

ANSWER 7.

Total head pressure is the sum of static and dynamic pressureairspeed.

ANSWER 8.

Compressor compression ratio is the ratio of compressor opressure.

ANSWER 9.

The purpose of inlet guide vanes is to direct the air flow intoacceptable angle of attack.

ANSWER 10.

The air passage between adjacent blades is slightly divergent, cpassage between the stationary vanes is also slightly divergent an



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auses its volume to decrease. If thes volume reduction would cause ageway dimensions, a uniform axial

otor blades they are "twisted" from

sure in the following stator vanes. (rotor + stator).

ffeting results.

ccur. Pressure losses would occur if



ANSWER 11.

The compression of the air as it passes through the compressor cblades and air passageway were of uniform cross section thireduction in velocity. By gradually reducing blade and air passavelocity is maintained.

ANSWER 12.

In order to maintain uniform axial velocity at all points on the rroot to tip.

ANSWER 13.

The velocity gained in the rotor blades is converted into presHence there is ideally no overall velocity increase through a stage

ANSWER 14.

Impeller tip clearance is critical, if it is too small aerodynamic bu

There would only be a danger of seizure were bearing failure to otip clearance is too large.



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ler tends to be preheated as it passes

rue of any increase in compressor for the same airflow the impeller

and a low airflow are the ideale occur at low compressor rpm.

irflow to sustain, is evidenced by ation gases forward, a marked rise in

ratio the curve separating unstable



ANSWER 15.

In a double-entry compressor the air to the rear side of the impelthe discharge from the front side.

Although increased turbine power will be necessary, this is tcapacity. An advantage of the double entry compressor is thatdiameter can be reduced, hence frontal area can be reduced.

ANSWER 16.

A high pressure ratio (excessively adverse pressure gradient)conditions for surge (reversal of airflow) to occur. Stall and surg

ANSWER 17.

Surge, which is caused by a pressure gradient too great for the acombination of a loud "bang", the expulsion of air and combusEGT, a loss of thrust, and/or vibration.

ANSWER 18.

When compressor airflow is plotted against compressor pressure(surge) and stable conditions is known as the surge line.



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mass flow to atmosphere, therefore therefore be reduced thrust and less

the early compressor stages at lowhus reducing pressure ratio.

utlet and the combustion chamber

ortion of the air flow (up to aboutthe air flow is called the primary air.

rimary zone.

ure equalisation and to assist flames purpose



ANSWER 19.

If an anti-surge bleed valve sticks open there will be a loss of airless mass flow through the "hot" section and jet pipe. There willcooling of the combustion gases.An open anti-surge bleed valve will prevent stall and surge.

ANSWER 20.

Anti-surge bleeds prevent surge by maintaining airflow throughrpm and mass flow by diverting some of the airflow overboard t

ANSWER 21.

The highest gas pressure occurs between the final compressor oinlet (the diffuser section).The highest velocity occurs in the propelling nozzle.

ANSWER 22.

The combustion chamber is designed to introduce a small prop40%) into the primary or combustion zone. This proportion of

ANSWER 23.

The highest gas temperature occurs in the combustion chamber p

ANSWER 24.

Multiple chambers are interconnected only for purposes of pressspread during initial start. Interconnection tubes are used for thi



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t fuel, reducing air pollution and

vary between 45:1 and 130:1, in thetio of about 15:1.

perature that the turbine blades and

nto the first row rotor blades at the

ngines.

such as starters and auxiliary poweractice there is always some impulse



ANSWER 25.

The annular combustion chamber virtually eliminates unburnincreasing thermal efficiency.

ANSWER 26.

Whilst the overall mixture ratio in the combustion chamber may combustion zone of the chamber air will mix with the fuel at a ra

ANSWER 27.

The combustion gas temperature is limited by the maximum temturbine wheel can safely withstand.

ANSWER 28.

Nozzle guide vanes accelerate the combustion gas and direct it ooptimum angle.

ANSWER 29.

Impulse/reaction blading is most commonly used in gas turbine e

Pure impulse turbines are generally limited to very small engines,units. Pure reaction blading is only theoretically possible. In preffect.



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e over the whole operating range ofated to ACC.

endent of the HP or LP compressorse or power turbine.

kage to a minimum.

n.

ion of the impulse force of the gas the gas being accelerated by the

ce both these values will fall. Gas and rotor blades.



ANSWER 30.

A method of maintaining constant minimum turbine tip clearancthe engine is known as Active Clearance Control usually abbrevi

ANSWER 31.

In some turbo-prop and turbo-shaft layouts a turbine rotor indepis used to drive the propeller or output shaft. This is called a fre

ANSWER 32.

The main function of turbine blade shrouding is to reduce tip lea

Compressor stator vane shrouding is also used to reduce vibratio

ANSWER 33.

The rotational force applied to the turbine rotor is a combinatbeing deflected by the blades and the reaction force due toconvergent ducts formed by the blades.

ANSWER 34.

The turbine converts heat and pressure energy into work, henvelocity is increased during its passage between both stator vanes



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gas accelerates through it until theressure ratio increases do not result

gas flowing across the rear face of

n air than low frequency.

the gas temperature will increase these. Increasing gas pressure simply

astly greater than the mass of gas



ANSWER 35.

As the pressure ratio across a convergent duct increases, the velocity at the throat equals the local speed of sound. Further pin further acceleration and the nozzle is said to be choked.

ANSWER 36.

The exhaust cone reduces gas velocity by diffusion and preventsthe turbine wheel.

ANSWER 37.

High frequency sound is attenuated more rapidly with distance i

The higher the exit gas velocity, the greater the noise. Increasing choke velocity (that's what reheat does) and increase the noichokes the nozzle, which is when it is noisiest.

ANSWER 38.

The mass of air accelerated by the fan (the cold stream) is vaccelerated by the combustion process (the hot stream).

ANSWER 39.

Clamshell Door reversers are pneumatically operated.



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the hydraulic actuators move theropelling nozzle forwards, to create



ANSWER 40.

With the target bucket system, when reverse thrust is selecteddoors, by means of pushrods, to deflect the jet stream from the pa reversal of jet thrust.





Gas Turbine Engine Systems

Introduction

Engine Starting

Ignition Systems

Fuel Systems

Gas Turbine Fuels

Gas Turbine Lubrication Systems



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ems

systems, the fuel systems and fuelhe fuels used in aircraft gas turbine

rbine will start satisfactorily. Firstly,e rotated up to a speed at which the fuel injected by the burners. fuel/air mixture in the combustion

ultaneously. It must be possible toquired for maintenance checks or tossible to use ignition alone (withouttioning of both these systems is co-



22Gas Turbine Engine Syst

Introduction1. This section deals with the engine starting and ignitioncontrol requirements of the various types aircraft gas turbine, tengines and the lubrication systems.

Starting Systems2. Two separate systems are required to ensure that a gas tuprovision must be made for the compressor and turbine to badequate air passes into the combustion system to mix withSecondly, provision must be made for ignition (light-up) of thesystem. During engine starting the two systems must operate simturn the engine with the starter motor, without ignition, when reblow out surplus fuel after a failed start. Similarly, it must be postarter) for relight, or prevention of flameout, in flight. The funcordinated during the start cycle.



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sor (and, of course, its associatedlly the one that is rotated, as this is

functions. Firstly it must rotate andfor combustion to take place andurbine until self-sustaining speed is

he engine accelerates under its own in excess of the torque required toine.

ion gearing. The electrical supply isd or, in the event of failure to start,

transmits power through reduction, from an on-board auxiliary powerype of starter unit used with modernigure 22-1.



3. Gas turbine units are started by rotating the compresturbine). In the case of dual rotor engines, the HP section is usuainvariably the hot gas section. The starter motor carries out twoaccelerate the compressor up to a particular speed suitable secondly, once the light-up has taken place, assist the engine texceeded.

4. Once self-sustaining speed is, the starter is cut out and tpower to idle speed. The torque provided by the starter must beovercome the compressor inertia and the friction loads of the eng

Engine Starting

Types of Starter Motor

Electric Starter Motor. A DC motor driving through reductautomatically cancelled when the engine has satisfactorily startewhen a time cycle is completed.

Air Starter Motor. A compressed air driven turbine rotor gearing. Air supply for the turbine may be from a ground supplyunit, or bleed air from a running engine. This is the commonest thigh by-pass engines. A typical air starter motor is illustrated at F



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FIGUJet EngStarter

ith a free turbine connected to the



RE 22-1ine Air Motor

Gas Turbine Starter. A self-contained small gas turbine wengine through reduction gearing.



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(refer to Figure 22-2). The starter is switch. As soon as the starter has through the engine, the ignition isieved by shifting the start lever fromimportant to prevent the danger of



Starting Sequence5. A typical starting sequence for a gas turbine is as followsturned to rotate the compressor by actuating the start controlaccelerated the compressor sufficiently to establish the airflowturned on and then the fuel flow. Both of these functions are achcut-off to idle. This sequence during the starting procedure is spontaneous ignition and over temperature.



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FIGUGas TuStartin

celerates. At this point the fuel flowate satisfactorily on its own. If theccelerate but far too slowly for thet energy. Hence, starter assistance isslightly above self-sustaining speed. starter motor.



RE 22-2rbine g System

6. After light up, assisted by the starter motor, the engine acrate is not sufficient to enable the engine to continue to accelerstarter is disengaged now the engine may not accelerate, may apermissible time cycle, or may even decelerate due to insufficienrequired after light up, and continues to assist the engine up to Thereafter the engine can accelerate without the assistance of the



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er and ignition are cut off, eithere now accelerates to, and stabilisesorter will be the time required until

d in between 20 and 90 seconds,under which the start is attempted.

permitted time for the start cyclen it becomes evident that either thexceed the maximum permitted peak

P cock) should be closed and theuently be motored with start lever

lready been de-energised, follow the

be cancelled by a time switch beforemay well continue to turn, but willcondition is allowed to persist, thelow will be excessive for the airflow

elerate and TGT will stay very low.and the excess fuel must be allowedllowed is, ignition off, motor enginegnition on) when the excess fuel hasmbustion chamber drain valves to

ctacular display known as torching,cted from the jet pipe as a stream of



7. After self-sustaining speed is exceeded, both the startautomatically or by releasing the start control switch. The enginat, idle rpm. The higher the rpm before the starter cuts out, the shidle rpm is reached. The entire starting cycle is accomplishedepending upon the type and make of engine and the conditions

8. It is important that the pilot is aware of the maximumevents. The pilot must be prepared to discontinue the start whestart cycle time will be exceeded (a hung start), or the EGT will estarting value (a hot start). In the latter case, the start lever (Hstarting sequence cancelled immediately. The engine can subseqshut and ignition off, to cool the hot section. If the starter has amanufacturer's recommended procedures.

9. If the engine is slow to accelerate the ignition circuit may self-sustaining speed has been exceeded. In this case the engine not accelerate beyond self-sustaining speed. If this hung-start turbine gas temperature is liable to become excessive, since fuel fat the low rpm pertaining.

10. If light up does not occur clearly the engine will not accWith fuel being scheduled the combustion chamber(s) will flood to drain off. This is known as a wet start. The procedure to be foon the starter to blow out excess fuel and only attempt light-up (igone - this may require allowing the engine to stop and the coopen. Failure to get rid of the unburned fuel can result in a spewhen the residual fuel ignites at the next start attempt and is ejeflaming fuel!



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as turbine engines are designed toeed is typically in the order of 30%

necessary to descend to a specified referred to as the re-light envelope).ot be operated and all that is neededis selected to flight start, which willer to open the HP cock, and risingbilise at idle rpm, which at altitude

e switched off and the thrust lever

a windmilling re-light with airspeedcorrect drill is to carry out a start-ring the starting sequence. A typical



11. The automatic fuel control systems fitted to modern gprevent the occurrence of hung and hot starts. Self sustaining spNz and ground idle speed approximately 60% Nz or 25% N.

Flight Start or Re-Light12. When an engine has been shut down in flight, it may bealtitude and/or reduce speed to obtain a positive re-light. (This isIf the engine is windmilling at sufficient speed, the starter need nis ignition and fuel. For a flight re-light, the start control switch start the ignition. Fuel flow is initiated by means of the start levEGT and rpm confirm a satisfactory re-light. The engine will stawill be higher than ground idle rpm. The ignition can now bgradually advanced to increase power.

13. A vital element of the re-light envelope is IAS. Attemptingtoo low will not result in a satisfactory start. At low IAS the assisted re-light to ensure adequate acceleration of the engine dure-light envelope is illustrated at Figure 22-3.



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FIGUJet EngRelightEnvelo

re a flame exists as long as fuel ist up and to stabilise the combustionint above which it will accelerate tore de-energised. The light up mustgy system is used for starting all gas



RE 22-3ine ing pe

Ignition Systems14. The gas turbine is a continuous combustion engine wheflowing. Hence, an ignition system is only required for initial lighprocess. Once the engine has reached self-sustaining speed (the ponormal running speed without further assistance) the igniters atake place under many different conditions and hence a high-enerturbine engines. A dual ignition system is frequently fitted.



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ected to its own igniter plug (not alocated in different positions in theor 115v AC supply. Normally the

the start lever (or HP cock) is set toen the engine rpm reaches a speed at

system is shown at Figure 22-4. Aich is rectified to put a high voltage stored in the condenser equals therges to an igniter plug to produce ation of the choke is to extend the



15. Each ignition system has a high energy ignition unit connsparking plug) and the two igniter plugs of a dual system are combustion section. The system operates on either a 28v DC igniters are automatically energised when, during the start cycle,deliver fuel to the rotating engine. They are then de-energised whwhich combustion becomes self-sustaining.

16. A high-energy, twelve-joule DC ignition unit electrical trembler mechanism operates an induction coil, the output of whcharge on a condenser, or reservoir capacitor. When the voltagebreakdown value of a sealed discharge gap, the condenser dischasequence of "flashovers" across the face of the plug. The funcperiod of discharge at the igniter plug.



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FIGUJet EngIgnitio



RE 22-4ine DC n Unit



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circumstances where ignition of the a flame out at high altitude. Fore joule system is adequate and willignition units are normally suppliedry in flight conditions where flame

system is shown at Figure 22-5.harge a capacitor with high voltage.e equals the breakdown value of a

of an igniter plug.



17. High-energy ignition systems are particularly necessary infuel is likely to prove difficult, such as when re-lighting aftercontinuous operation of the ignition system a low-energy, threresult in longer life of the igniter plugs. Low energy gas turbine with AC. Continuous operation of the ignition system is necessaextinction might occur, such as heavy rain, snow or icing.

18. A low-energy, three-joule AC ignition unit electrical Alternating current is passed through a step-up transformer to cAs with the previously described system, when capacitor voltagsealed discharge gap the stored energy discharges across the face



Chapte

FIGUJet EngIgnitio



RE 22-5ine AC n Unit



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s duty cycle (typically 10 minutes ofpossible malfunction of the ignitions simply a matter of remembering tor as otherwise specified.

ms that are automatically energisedndition, since compressor stall and

ch, whilst energised, produce a high- to a gap-jumping spark. A typical



19. Ignition systems normally have a duty cycle specified. Thicontinuous operation of a single system) prevents damage and systems. When operating in bad weather with the ignition on, it iswitch from one ignition system to the other every ten minutes, o

20. Many gas turbine-powered aircraft employ ignition systewhen the angle of attack indicator senses an incipient stall cosurge are likely to add to your problems in this situation.

21. Igniter plugs are usually of the surface discharge type whienergy flashover of 60-100 discharges per minute, as opposedigniter plug is shown in section at Figure 22-6.



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FIGUTypica



RE 22-6l Igniter Plug



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on system is the means by which theher thrust is required, the throttle isfuel supply to the burners increases gas temperature in the combustione to give a higher engine speed andst is produced. The fuel flow is

te to the engine burners to produce

order to maintain a selected engine

is demanded.

ng.

fuel supply.



Fuel Systems22. Regulating the quantity of fuel injected into the combustipower, or thrust, of a gas turbine engine is controlled. When a higopened by the action of the thrust lever and the pressure of the due to the greater fuel flow. This has the effect of increasing thechamber, which in turn increases the gas flow through the turbincorrespondingly greater airflow. In consequence, greater thruminimum at idle power and maximum at take-off.

23. The main functions of the fuel control system are:

(a) To deliver fuel at the correct pressure and flow rathe power required by the pilot.

(b) To adjust fuel flow as inlet conditions change in power (automatic barometric control).

(c) To provide extra fuel flow when rapid acceleration

(d) To provide adequate fuel flow during engine starti

(e) To enable the engine to be stopped by shutting off



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safety controls to prevent EGT, rpmitations. In order to accomplish the

d engine conditions, compares themo produce the desired engine thrustnd air flow/fuel flow adjustments.

means of the throttle (thrust lever),equired. Having set the throttle, fuelair density, and therefore air massircraft speed. The fuel control unitronic control system, achieves this engine is shown at Figure 22-7.

FIGUSimplifTurbinSystem



24. In addition, the fuel control system must have automatic and compressor delivery pressure from exceeding their design limcorrect fuel flow, the fuel control unit senses the atmospheric anto the throttle position set and delivers the fuel flow necessary toutput. The pilot is thus relieved of normal speed, temperature a

25. Control of power or thrust of a gas turbine engine is by with which fuel flow is manually adjusted for the power setting rflow is controlled automatically to compensate for changes in flow, due to changes in aircraft altitude, air temperature and a(FCU), which may be a hydro-mechanical device or an electautomatic fuel control. A simplified fuel system for a gas turbine

RE 22-7ied Gas e Fuel



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igh-pressure (HP) pump by a low-ly at a suitable pressure to prevent

e satisfactory engine operation. Themation of ice crystals which would

pe, delivering a variable quantity of-8.

FIGUHP Fu



26. Fuel is supplied, from the aircraft fuel tanks, to the hpressure (LP) system. The LP fuel pump ensures a constant suppvapour locking and cavitation of the HP pump supply, to ensurLP system usually incorporates a fuel heater to prevent the forblock the fuel filter.

27. The HP pump is usually of the multi-piston swash plate tyfuel at constant pressure. Such a pump is illustrated at Figure 22

RE 22-8el Pump



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mp. Fuel quantity delivered to thetrol unit. Fuel supply to the burners

put of the HP fuel pump varies thestems the HP pump output is varied all, of the following:

esigned to adjust fuel supply to the engine rpm, P1 (air intake) pressure



28. Some low-pressure fuel systems use a spur gear HP puburners is set by the throttle lever and controlled by the fuel concan be shut off altogether by the HP shut-off cock.

Fuel Control Systems29. In most gas turbine fuel control systems altering the outsupply of fuel to the burners. In hydro-mechanical fuel control syby a hydraulic servo system in response to variations of some, or

(a) Throttle position

(b) Ambient air temperature and pressure

(c) Engine rpm

(d) Rate of acceleration or deceleration

(e) Exhaust/Turbine gas temperature

(f) Compressor delivery pressure

30. Figure 22-9 shows a simplified pressure control system dburner spray nozzles in response to changes of throttle position,and exhaust gas temperature.



Chapte

FIGUSimplifEnginePressuSystem



RE 22-9ied Jet Fuel re Control



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t Figure 22-8, in which variation ofive to the pressure in a servo systemion/deceleration in such a system ist system limits the rate at which theottle lever is moved. The operationol system shown at Figure 22-9 is

e control piston. If pump dischargeswash plate angle and hence pumpn moves down, under the influencep pressure is supplied direct to the

supply through a fixed orifice.

hich control the rate at which fuel is exactly equals supply through theant. If spill exceeds supply, servose.

valve determines the pressure droptant supply pressure to the throttleross it will be. Opening the throttlet pressure), throttle inlet pressure is

ton. The pressure drop across thetrols the position of a spill valve.



31. The HP fuel pump is of the swash plate type illustrated athe swash plate angle varies pump output. A control piston sensitvaries the pump swash plate angle. Control of rate of acceleratusually incorporated in the throttle valve mechanism. A dashpothrottle valve opens or closes, regardless of how rapidly the thrand interaction of the component parts of the Pressure Contrdiscussed in the following paragraphs.

32. HP Pump pressure is maintained at a constant value by thpressure tends to rise the piston moves up, reducing the pump output. If pump discharge pressure tends to fall the control pistoof the pressure control spring, increasing pump output. HP pumthrottle valve and also supplies the servo system with a constant

33. Servo Pressure is controlled by a number of spill valves, wspilled back to the LP system. If flow through the spill valvescontrol piston fixed orifice, servo pressure will remain constpressure will fall, if spill is less than supply, servo pressure will ri

Throttle movement. The degree of opening of the throttle across the valve. Given that the control piston ensures a consvalve, the more the valve is opened the less the pressure drop acincreases the supply pressure to the burner nozzles (throttle outlemaintained constant by the servo-operated pump control pisthrottle valve is sensed by a spring-loaded diaphragm, which con



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across the diaphragm decreases andmp servo pressure increases, whichsing fuel flow to the burner nozzles.phragm, opening the spill valve toump output.

ns a spill valve. If rpm increases the reduce pump servo pressure. This

e rpm falls back to the pre-set value. force of the flyweights to close the

sule attached to a spill valve. An

reased airspeed or decreased airlow. Increased pressure will expandpump servo pressure and HP pump

gas temperature is limited, in order is sensed by thermocouples and thes a solenoid to open a spill valve asre HP pump output.

only used for turbo-jet and turbo-tems are the Flow Control system,

Pressure Ratio Control system. Theters and engine type.



34. If the throttle is opened, for more power, pressure drop the spring force moves the diaphragm to close the spill valve. Puacts upon the control piston to increase HP pump output, increaClosing the throttle increases the pressure drop across the diadecrease pump servo pressure, which results in a decreased HP p

Engine rpm is sensed by an engine-driven governor, which positiogovernor flyweights move outwards to open the spill valve andcauses the control piston to decrease HP pump output and enginIf rpm drops below this value the governor spring overcomes thespill valve, increasing servo pressure and HP pump output.

Compressor Air Intake (P1). Pressure is sensed by a cap

increase in intake pressure (due to decreased altitude, inctemperature/density) requires a corresponding increase of fuel fthe pressure-sensing capsule, closing the spill valve to increase output. A fall in air intake pressure will have the reverse effect.

Exhaust Gas Temperature. It is important that the exhaustto avoid damage to the turbine blades and inlet guide vanes. EGTelectrical output, which is proportional to temperature, activatetemperature increases, reducing pump servo pressure and therefo

35. Pressure Control fuel control systems are quite commpropeller engines. Alternative hydro-mechanical fuel control systhe Combined Acceleration and Speed Control system and the type of system chosen will depend upon engine operating parame



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t fuel pump delivery pressure is notump output (fuel flow) is controlled constant air intake conditions. The

the HP fuel pump, are contained engine speed governor, an altitudeontrol unit.

senses engine speed and produces ae (and therefore output).

trol, which senses air intake (P1)d therefore fuel flow) in conditions

pressure and adjusts a spill valve to

eeds a maximum limit a solenoid-en a spill valve and reduce servo



Flow Control System36. Flow control systems differ from pressure control in thamaintained constant, but is proportional to engine speed. Fuel pto maintain a constant pressure drop across the throttle valve atsystem is better suited to engines requiring large fuel flows.

37. The control system components, with the exception ofwithin a combined fuel control unit. These components are thesensing unit, an acceleration control unit and a gas temperature c

Engine Speed Governor. A hydro-mechanical device whichproportional hydraulic servo pressure to control fuel pump strok

Altitude Sensing Unit. This is a barometric pressure conpressure and operates a spill valve to control servo pressure (anbelow governing speed.

Acceleration Control Unit. This senses compressor deliverycontrol servo pressure to match fuel flow to airflow.

Gas Temperature Control. If engine gas temperature excoperated proportioning valve is progressively energised to oppressure, reducing fuel flow.



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ntrol System. The controlling unit is a fuel flowontrolling fuel flow. It contains twontrol governor. The speed control to the burners by means of a sleeve

due to an increased fuel flow and set value relative to engine rpm. A adjusts fuel flow to match airflow.

or automatically adjusts the throttle

tion control, which uses the ratio of, to control fuel flow. This system isn (flame out). Either condition willm will substantially reduce, or even



Combined Acceleration and Speed Co38. This is a mechanical system that does not use spill valvesregulator, which is engine driven and controls engine speed by cgovernors, a speed control governor and a pressure-drop cogovernor is set by the engine throttle lever and controls fuel flowvalve.

39. The pressure-drop governor senses any pressure drop maintains fuel flow, by adjustment of a second sleeve valve, at acapsule unit compares compressor inlet and outlet pressures andEngine gas temperature is sensed electrically and a rotary actuatmechanism if maximum temperature is reached.

Pressure Ratio Control System40. This is a mechanical system, similar to speed and acceleraHP compressor delivery pressure (P4) to air intake pressure (P1)particularly responsive to conditions of surge or flame extinctiocause the P4/P1 pressure ratio to be abnormally low and the systeshut-off, fuel flow to the burners.



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nic controls that monitor engine parameters within pre-set operatinghaust gas temperature and engine

preventing pre-set parameters fromh maintains the thrust condition set

ll authority digital engine controlumatic functions of the fuel controlntrol of the engine conditions. Thet-off cock) for safe operation in the

n FADEC, known as full authority, but does not have the same degree

e governor to match actual engineengine supervisory control systemor bleed state, engine pressure ratiod/Mach No and air temperaturesystem computes the LP spool speed angle set by the pilot and comparese the necessary N1 and trimmed to



Electronic Engine Control41. Many modern gas turbine engines incorporate electroperformance and operate the engine controls to maintain certainlimits. These parameters are, typically, engine spool speeds, expressure ratio (EPR). The system may act simply as a limiter, being exceeded, or it may be a supervisory control system, whicby the pilot, regardless of changing atmospheric conditions.

42. Currently the ultimate extension of this concept is fu(FADEC), which replaces most of the hydro-mechanical and pnesystem and virtually takes over all steady state and transient cofuel system retains only sufficient controls (throttle and HP shuevent of a major electronic failure.

43. A slightly less sophisticated electronic control system thafuel control (FAFC) uses an electronically-controlled fuel systemof transient condition control of the compressor airflow system.

44. Supervisory systems trim the fuel flow scheduled by thpower with that demanded by throttle lever position. The monitors such parameters as thrust lever angle, engine compress(turbine outlet to compressor intake) and altitude, airspeeinformation from the air data computer (ADC). The supervisory (N1) necessary to achieve the EPR demanded by the thrust levercommand EPR with actual EPR. Fuel flow is adjusted to achievachieve the command EPR.



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system continuously monitors N1,is to say, the system will not allow

es and that most commonly used isn process, between gasoline and thetly higher calorific value. Its freezingwaxy particles that are capable of and -50°C. The lower volatility of

fuel tanks or systems than gasoline.

sene type of fuel of relatively low

uels are available. These fuels are ach lower freezing point than pure

in geographic locations where lowe fuels are much more volatile thanf temperatures, making then morene to vapour locking.



45. Throughout the fuel flow/power adjustment the controlEPR and EGT and operates on a ‘lowest wins’ precept. That limiting values of any of those three parameters to be exceeded.

Gas Turbine Fuels46. Gas turbines require a less volatile fuel than piston enginkerosene. Kerosene occupies a narrow band of the fuel distillatioheavier fuel oils. Kerosene is denser than gasoline and has a slighpoint varies according to its hydrocarbon content, but solid blocking fuel filters begin to form at temperatures between -40°Ckerosene means that it is much less susceptible to vaporisation in

47. UK commercial aircraft use JET A-1, which is a kerovolatility. The equivalent in the USA is known as JET A.

48. In some countries, notably the USA, wide-cut gasoline fmix of kerosene and gasoline and have the advantage of a mukerosene, typically about -60°C. Thus they are suitable for usetemperatures prevail and at extreme altitudes. Wide-cut gasolinkerosene and produce flammable vapour over a wider range ohazardous when fuel tanks or systems are ruptured and more pro



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er specific gravity (SG) than JET A-he lower SG fuel will result in theith a change in SG of the fuel, theined acceleration and speed controling any need for adjustment upon

re at which the water content of in very cold climates this is usuallyhe principal disadvantage to the usenents are not susceptible to attack

ng with water in the fuel at lowmix below that of kerosene.

live and grow in tanks containingnvariably is to some extent. Theoloured red, brown or black. Theyan lead to serious weakening of theng the protective coating applied to

t the growth of microorganisms, aircraft whose tanks have not been



49. Wide-cut gasoline fuel is designated JET B and it has a low1. If used in an aircraft with a pressure control fuel system tgovernor maintaining a greater engine speed. Consequently, wgovernor must be reset before the engines are operated. Combfuel systems compensate for changes in SG of the fuel, precludchange of fuel.

Fuel Additives50. Additives have been devised that reduce the temperatukerosene will freeze. For aircraft operating at extreme altitude ora preferable alternative to the use of volatile wide-cut gasoline. Tof icing inhibitors is the need to ensure that fuel system compofrom them. In general, an icing inhibitor works by combinitemperatures, to reduce the freezing point of the water/inhibitor

51. Microorganisms, in the form of fungi or bacteria can kerosene, provided water is present in the fuel, which it imicroorganisms appear as a slimy deposit in the fuel, typically cproduce chemicals that are highly corrosive to metal and which cwing structure in aircraft with integral fuel tanks, even penetratithe inner surface of the tanks.

52. Fuel additives are commercially available that prevenparticularly fungi, and these are essential in gas turbine-poweredcoated with fungus-resistant material.



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nd icing inhibitor. When it contains

of ice crystal formation at low-existent. Unfortunately, fuel has anthough every precaution is taken tond conveniently sinks to the bottomis particularly difficult to detect and

lution is largely dependent uponater the saturation level. As fuel

l separates as small water particles,

ides lubrication and cooling for allth other moving parts (for examplebo-prop engine the oil system alsooil for the propeller pitch control



53. JET A1 is available with or without fungus suppressant athese additives it is known as JET A1 (FSII).

Water in Fuel54. If there were no water content in fuel the problemstemperatures and microorganism growth would be virtually nonaffinity for water and it is impossible to remove it altogether, alexclude it as far as possible. Whilst bulk water is clearly visible aof a fuel container, it is dissolved water (water in solution) that remove.

55. The amount of water that kerosene can hold in sotemperature, the higher the temperature of the fuel the gretemperature falls, that water content above the saturation levewhich combine with the fuel to form a frozen gel.

Gas Turbine Lubrication Systems56. The lubricating oil system of a gas turbine engine provparts of the engine where moving parts are in contact either wigears) or with stationary parts (for example bearings). In a turprovides lubrication for the propeller reduction gearing and mechanism.



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sity must be low enough to permitand and absorb mechanical loads.ngine, gas turbine lubricating oil isasier and normal starts, without the°C. Mineral oils are generally not ofre ranges of a gas turbine engine, so

y lubrication systems that are self-und the engine by oil pumps, withosses in flight. Some short durationpended overboard after circulationtem, the pressure relief valve systemfe operation are oil temperature and

ntrolled by maintaining a constant valve on the outlet side of the oiln side of the pump. The valve is set

speed. The faster the engine (and,intain constant supply pressure. Aigure 22-10.



57. The oil used must protect against corrosion and its viscoflow at low starting temperatures, but high enough to withstBecause these loads are generally much lower than in a piston eusually of lower viscosity. This makes low temperature starting eneed to pre-heat the oil are possible at temperatures down to -40sufficiently constant viscosity over the wide operating temperatusynthetic low viscosity oils have been developed.

58. Commercial aircraft gas turbine engines use recirculatorcontained. Oil is contained in an oil tank and distributed arosufficient reserve of oil in the tank to cope with normal, minor lengines employ a total loss lubrication system in which oil is exthrough the engine. There are two basic types of recirculatory sysand the full flow system. In both systems the factors crucial to saoil pressure and these are invariably indicated on the flight deck.

Pressure Relief Valve System59. In this system the oil flow to the bearing chambers is copressure in the oil supply line. A spring-loaded pressure reliefpressure pump opens to return oil to the oil tank or to the suctioto begin opening at a pressure corresponding to engine idlingtherefore, pump) speed the more the relief valve opens to maschematic diagram of a pressure relief valve system is shown at F



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FIGUJet EngPressuSystem



RE 22-10ine re-Relief Oil



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ncy of bearing chamber pressure tore, which reduces the oil flow to the relief valve system is unsuitable forecause of high bearing loads. Theessure relief valve spring force withre is necessarily high an alternative

ead feeds oil pressure pump outputhigher the engine (and pump) speedall the oil pumped is supplied to theve, pump sizes can be significantly

w system, pressure-limiting by-passs and filters, which could otherwise Full Flow System are dry sump

rporates a sight glass or dipstick forr gravity filling. The tank has a de-rning from the bearing chambers. A



60. An inherent problem with this type of system is the tendeincrease with increased engine speed. This creates a back-pressubearing chambers as engine speed increases. Hence, the pressureengines where a high bearing chamber pressure is necessary bproblem can be alleviated to some extent by augmenting the prbearing chamber pressure, but in engines where chamber pressulubricating system, the full flow system, is used.

Full Flow System61. This system does not use a pressure relief valve, but instdirect to the oil supply line to the bearing chambers. Thus, the the greater the quantity of oil supplied to the bearings. Because bearings, with none being spilled back by a pressure relief valsmaller with this type of system.

62. Because of the higher pressures associated with the full flovalves are fitted in conjunction with components such as coolerbe damaged. Both the Pressure Relief Valve System and thelubrication systems.

Lubricating Oil System Components63. The oil tank is mounted externally on the engine and incochecking contents. Oil tank replenishment may be by pressure oaeration device in the return line to remove air from the oil retutypical gas turbine engine oil tank is shown at Figure 22-11.



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FIGUTypicaTurbin

f the spur gear type, although vane The scavenge pumps and pressure the accessories gearbox.

s of a spray from a jet orifice.



RE 22-11l Gas e Oil Tank

64. The scavenge pumps and pressure oil pump are usually otype or gerotor (Roots type) pumps are used on some engines.pump are contained in a common casing with a single drive from

65. Oil distribution to gears, bearings, etc is usually by mean



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of tubes through which the coolingd heat is transferred by conduction some engines both fuel-cooled and use at high powers when the fuel-

shown at Figure 22-12.

FIGUFuel-CCoole



66. The oil cooler is a heat exchanger comprising a matrix medium flows. Oil is directed over the outside of the tubes anfrom oil to coolant. The cooling fluid is either fuel or air and inair-cooled oil coolers are used, the latter only being brought intocooled cooler cannot cope. An example of a fuel-cooled cooler is

RE 22-12ooled Oil r



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in event of blockage of the cooler. Aher value than fuel pressure, is also

e, oil leaks into the fuel rather than

bers and gearboxes are collected by. Small permanent magnets collect

ng of potential component failure.ification of precisely which gear orconnected to a flight deck warningre. An example of a magnetic chip

FIGUMagneDetec



67. A by-pass valve both protects the cooler, and the engine, pressure-maintaining valve, which maintains oil pressure at a higa common feature. This ensures that, in the event of tube failurthe reverse, which could be damaging to engine bearings.

68. Small particles of metallic debris from the bearing chammagnetic chip detectors fitted in the scavenge side of the systemferritic metal particles that, upon examination, can give warniIdentification of the collected material can often provide identbearing is failing. Magnetic chip detectors (MCD's) are often system to give in-flight indication of impending component failudetector is illustrated at Figure 22-13.

RE 22-13tic Chip tor



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In order to reduce the transmission squeeze film is often used. Oil from the outer race (cage) of the bearing absorbs the radial shock loads of



69. Gas turbine engine bearings are of the ball or roller type.of vibration from the rotating assembly to the bearing housing athe pressure supply system is fed to a narrow clearance betweenand the bearing housing. The film of oil filling the clearancevibration. A squeeze film bearing is illustrated at Figure 22-14.



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FIGUSqueezBearin

gs, is separated and exhausted toreather as shown at Figure 22-15.



RE 22-14e Film g

70. Air, which tends to accumulate in the oil at the bearinatmosphere, before the oil returns to the tank, by a centrifugal b



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FIGUCentriBreath



RE 22-15fugal er



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ystem to prevent the circulation ofcavenger pump inlet. Fine, pressuref the oil feed jets. In addition, smalln example of a pressure filter, with

Figure 22-16.



71. The lubricating oil is filtered at various points in the sdebris. Coarse strainers are fitted at the oil tank outlet and/or sfilters are fitted at the pressure pump outlet to prevent blockage othread-type filters are often fitted just upstream of the feed jets. Aa number of wire-wound elements providing edge filtration, is at



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FIGUJet EngFilter



RE 22-16ine Pressure





Gas Turbine Engine Performance and Operation

Introduction

Factors Affecting Power Output

Gas Turbine Operation and Monitoring



n

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ormance and

upon the mass of air that passesby the engine. In the turbo-jet and at the propelling nozzle or nozzles,).

e the thrust developed, is dependente. The forward speed of the aircraftm effect at the air intake when the

uced by the engine when the aircraft thrust is the thrust produced by they an engine is largely the product off the exhaust gas at the propelling the speed of the aircraft.

Gas Turbine Engine Performance and Operatio


23Gas Turbine Engine PerfOperation

Introduction1. The performance of a gas turbine engine is dependentthrough the engine and the acceleration imparted to that mass turbo-fan engine the performance is measured in terms of thrustin the turbo-prop engine it is measured as shaft horsepower (SHP

2. The mass of air passing through the engine, and thereforupon the air density, which varies with temperature and pressuralso affects air mass flow through the engine, since there is a raaircraft is moving.

3. Static thrust, or gross thrust, is the amount of thrust prodis stationary on the ground or the engine is on a test bench. Netengine when the aircraft is in flight. The static thrust produced bthe mass of air passing through the engine and the velocity onozzle. To calculate net thrust it is necessary to take into account



n

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of air (wa) divided by accelerationnd the velocity of the jet efflux (vj).

forward speed of the aircraft in thedifference between the exhaust gaslocity is, of course, the same as the substituted with V2 – V1 and the

zzles in (section 3.37), a turbo-jetfrom the pressure drop across there difference across the nozzle and



The Thrust Formula4. The mass of air passing through the engine is the weightdue to gravity (g). The thrust produced is the product of this aThe static thrust force is given by the formula:

5. To calculate the net thrust it is necessary to include the equation, because the thrust developed will depend upon the velocity (V2) and the initial gas velocity (V1). The initial gas veaircraft velocity. Hence, vj in the static thrust formula can beformula for calculating static or net thrust becomes:

6. When calculating static thrust, V1 will be zero.

7. As described in the section dealing with propelling nooperating under choked conditions derives additional thrust propelling nozzle. This added thrust is the product of the pressuthe area of the nozzle throat and is given by:

FWa Vj×

g---------------------=

FWag

-------- V2 V1–( )=

F P P0–( )= a×



n

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ting with the propelling nozzle in a

reams to be taken into account, the fan. Given that the hot gas exhausting the total thrust may be given as:

shaft horsepower (SHP) or thrusthe residual jet thrust of a turboprop

a

1) Fan( )



Where: p = static pressure at the nozzle throat

p0 = ambient atmospheric pressure

a = area of the propelling nozzle throat

8. Consequently, the static thrust of a turbo-jet engine operachoked condition is given by the formula:

9. When considering a turbo-fan engine there are two gas sthot gas stream from the jet pipe and the cold gas stream from theis normally un-choked in such an engine, the formula for calculat

Turboprops10. The output of a turboprop is measured normally as equivalent shaft horsepower (TESHP). TESHP takes account of texhaust as shaft horsepower.

FWag

-------- V2 V1–( )= P P0–( )+ ×

FWag

-------- V2 V1–( ) Jet( )=Wag

-------- V2 V–(+



n

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e engine must decrease as airspeedreasing. The greater the value of V1,h at Figure 23-1.

ease with increasing forward speed,take converts the increased velocity increasing the engine thrust. Theh at Figure 23-1.

the reduced differential between V2wa due to increased density, becauseat importance in increasing engineve C of the graph at Figure 23-1.



Factors Affecting Power Output

Effect of Airspeed11. From the foregoing it appears that the net thrust of thincreases since, at any given altitude, V2 is constant and V1 is incthe lower the net thrust. This is indicated by curve A on the grap

12. However, the ram effects at the compressor air intake incrforcing air into the engine at greater velocity. The design of the inof the intake air into pressure, increasing the density, therebyincreased thrust due to ram effect is shown in curve B of the grap

13. In terms of the thrust formula, the loss of thrust due to and V1, as airspeed increases, is largely offset by the increase in of ram effect. At supersonic airspeeds ram effect takes on grethrust. The overall effect of airspeed upon thrust is shown in cur



n

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FIGURam E

ched by an increased fuel flow, asonic airspeeds, net thrust decreasess shown in the upper two graphs atn (sfc) of a turbo-jet engine must is defined as pounds of fuel burneds is shown in the lower graph at



RE 23-1ffect

14. The increased airflow, due to ram effect, must be matdiscussed in the section covering fuel control systems. At subsslightly with increased airspeed and fuel consumption increases aFigure 23-2. From this it follows that specific fuel consumptioincrease with increasing airspeed, since specific fuel consumptionper hour per pound of thrust developed (lb/hr/lb.thrust). ThiFigure 23-2.



n

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FIGUEffect - Turbo



RE 23-2of Airspeed jet



n

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. Because SHP (shaft horsepower)o-prop aircraft up to the maximumat Figure 23-3.



15. In turbo-prop aircraft sfc is defined as lb.fuel/hr/SHPincreases with increased airspeed, sfc shows an increase in turbpropeller efficiency airspeed (about 350 mph). This is illustrated



n

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FIGUEffect - Turbo



RE 23-3of Airspeed prop



n

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re and temperature decrease. Thee tropopause this effect is partlydensity decrease with altitude. Any mass entering the compressor andtly thrust of a gas turbine engine

e (36,089 feet) where temperature. With no compensating effect from much more rapid rate above theis reason that the optimum cruising

fect upon engine performance for amance of gas turbine engines.

ed density will increase the mass of is greater. However, the compressorenser air so the engine will requiree speed will decrease.



Effect of Altitude16. As altitude increases the ambient atmospheric pressureduction in pressure decreases the air density but, up to thcompensated by the reduction in temperature, which limits the reduction in air density must, however, result in a reduced airtherefore a reduced mass flow through the engine. Consequendecreases with increased altitude.

17. The effect becomes most marked above the tropopausremains constant up to 65,617 feet but pressure continues to falltemperature, air density due to pressure reduction falls at atropopause, with a consequent rapid decrease in thrust. It is for thaltitude for long-range cruising is 36,000 feet.

Effect of Temperature18. At lower levels, outside air temperature has a marked efgiven altitude. This is of particular importance to take-off perfor

19. When air temperature is lower than standard the increasair entering the compressor so, for a given engine rpm the thrustwill require greater power to maintain the same speed in the dmore fuel. Alternatively, if the fuel supply is not increased, engin



n

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d density will result in a decrease inh beyond 100% rpm so, if the massestored. This means that some formmethanol injection, as discussed in

ompressors is dependent upon their Further, the change of thrust (massigh engine speeds. For this reason,lso for this reason that it is usuallynd, especially as gas turbine engines

bine engine is directly related to therottle (thrust) lever is set for a givenf engine thrust. For example, if theecrease with altitude because of the

ic engine control systems, this iseed to maintain thrust at a valuelayed as engine pressure ratio (EPR)-4.



20. When air temperature is higher than standard the reducethrust for a given rpm. The engine cannot be safely operated mucflow for take-off thrust is to be achieved, the air density must be rof thrust augmentation must be used, such as water or water-section 3.38.4.

Effect of Engine RPM21. The air mass flow produced by the gas turbine engine cspeed of rotation, the higher the rpm the greater the mass flow.flow) for a given change in engine rpm is most marked at hcruising rpm is usually 85% to 90% of maximum rpm. It is adesirable to maintain high engine rpm during the approach to laare inherently slow to ‘spool up’ (increase rpm)

Flat Rated Thrust22. It is clear from the foregoing that the thrust of a gas turdensity of the air in which it is operating. Consequently, if the thfuel flow then any change in air density will result in a change oaircraft climbs with the throttle lever set, the engine thrust will ddecreasing air density.

23. In many modern engines, with full authority electronovercome by automatic compensation that adjusts engine spdemanded by the pilot’s throttle lever angle. Engine thrust is dispon an EPR gauge, an example of which is illustrated at Figure 23



n

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FIGUEPR G

the throttle lever angle until the bug set by the thrust management

ystem will automatically adjust fuel varying altitude, airspeed and inletnt (flat) value regardless of changes

engines equipped with it are known



RE 23-4auge

24. To select the required thrust (EPR), the pilot adjustscommand EPR needle on the gauge is aligned with the referencecomputer in the electronic engine control system. The control sflow, and thus engine speed, to maintain the required thrust withair temperature. In other words, thrust is maintained at a constain air density.

25. Such a system is known as a flat rated thrust system and as flat rated engines.



n

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oringingle lever, known as the throttle or, thrust. Engine rpm is normally

number of respects from the pistonth on the ground and in the air.

s that there is a considerable influx. On the ground great care must be, since even quite small solid objectsll nicks in these blades are sufficient

of sufficiently high velocity to berable distance behind the engine.

s forward of and behind the engines. a running aircraft turbine engine.

articularly during engine start, sinceighly susceptible to damage from



Gas Turbine Operation and Monit26. Control of a gas turbine engine is usually by means of a smore correctly, power lever. This is used to select the desiredgoverned within certain limits.

27. The characteristics of the gas turbine engine differ in a engine and these affect the way in which it must be operated, bo

28. The much greater air mass flow through the engine meanof air at the intake, even when the engine is operating at idle rpmtaken to ensure that foreign objects are not drawn into the intakecan cause significant damage to fan and compressor blades. Smato create damaging out-of-balance forces at high rpm.

29. The jet efflux from a turbo-jet or turbo-fan engine is hazardous to personnel and ground equipment to a considePersonnel and equipment should be kept clear of the hazard areaIt is not unknown for ground personnel to have been ingested by

30. It is important to monitor the turbine gas temperature, pthe turbine nozzle guide vanes and first stage blades are hoverheating.



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um efficiency at high rpm, typicallyn the engines are being operated atore slowly than a piston engine.erably longer to be met. The fuelrate at which fuel flow increases, to

susceptible to the ingestion of waterchambers when flying in heavy rained.

be damaging to the airframe. Gasage from bird ingestion and many

e exercised when flying at low level,

e listed below.

tio (EPR). Usually the pressure ratioough on large fan engines it is oftenpressor inlet pressure.

measured by a torquemeter. Powering and a reference value.



31. The gas turbine compressor only operates close to maximabove 85% of maximum rpm. This must be borne in mind whelow rpm in flight, as a turbine engine accelerates much mConsequently demands for an increase in power take considscheduling system will almost certainly be ‘slugged’ to limit the avoid compressor stall and surge.

32. The high air mass flow of the engine also means that it is in sufficient quantity to extinguish the flame in the combustion or snow. In these conditions continuous ignition should be select

33. In any aircraft, encountering a bird when in flight canturbine compressor and fan blades are particularly prone to damengines have been wrecked in such encounters. Great care must bas during approach and take-off.

Instrumentation34. The engine operating parameters indicated to the pilot ar

Thrust. Generally indicated in the form of engine pressure rabetween jet pipe pressure and LP compressor inlet pressure, althan integration of fan discharge/turbine outlet pressure to LP com

Torque. The power output of a turboprop engine is usuallyassessment is made by comparison between the torquemeter read



n

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f maximum rpm. On multi-spooln speed (N1) is usually indicated as

pe temperature (JPT), exhaust gasET) is the most critical of engine

rotational speeds of gas turbines,eration.

upply is provided.

engine, since it provides a valuableicator is included in the engine

w levels of vibration in normallure. Typical causes of vibration are lead to catastrophic engine failure.vibration within critical frequency

tems, and their operation, are dealtents, at the end of this book.



RPM. Engine speed is measured usually as a percentage oengines HP spool rpm (N2) is always measured, LP spool or fawell.

TGT. Turbine gas temperature, also referred to as jet pitemperature (EGT), turbine inlet/entry temperature (TIT or Ttemperatures. It is measured by a system of thermocouples.

Lubricating Oil Temperature and Pressure. At the high monitoring of oil temperature and pressure is essential to safe op

Fuel Temperature and Pressure. Indication of the LP fuel s

Fuel Flow. An indication of the fuel flow is given for each indication of unit performance. Frequently a fuel-used indinstrumentation.

Vibration Monitoring. Gas turbine engines have very looperation so vibration is an indication of incipient potential faidamaged fan, compressor or turbine blades - each of which couldVibration monitors transmit a signal of relative amplitude of ranges appropriate to the engine and its components.

35. The various powerplant-monitoring instruments and syswith in detail under Powerplant and System Monitoring Instrum






Location

Fuel System

Starting and Ignition

Fire Protection and Cooling Systems

Shutting Down

Ram Air Turbine



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ity of transport aircraft to supply ground with its main engines shut. Normally the APU is operated onlyaircraft can utilise the APU in flightn in flight is usually only possible

nstant speed gas turbine coupled toto the aircraft's main generators iss (fuel pump, lubricating oil pump,ion, speed governing and overspeed

at Figure 24-1. A load compressorg the same air supply as the power

stem. The gearbox containing thetating assembly



24Auxiliary Power Units

1. Auxiliary power units (APU’s) are fitted in the majorelectrical power and compressed air when the aircraft is on thedown, in order to reduce the need for ground support equipmentwhen the aircraft is on the ground, but many modern transport as an alternative source of electrical power, if needed. Operatiowithin specified maximum airspeed and altitude limitations.

2. The APU is typically a self-contained unit comprising a coa gearbox, from which a generator of similar type and rating driven. The APU gearbox also drives the gas turbine accessorietachometer and a centrifugal switch controlling the starter, ignitprotection circuits).

3. A typical Auxiliary Power Unit is illustrated in section mounted on the same shaft as the APU power section and sharinsection compressor, supplies air to the aircraft pneumatic sygenerator and ancillary drives is shown at the end of the APU ro



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FIGUA TypiPower



RE 24-1cal Auxillary Unit



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t, usually in the rear fuselage andess to the compartment is external.

FIGULocatioAuxillaUnit



Location4. The APU is located in an un-pressurised compartmenseparated from the remainder of the aircraft by a firewall. AccFigure 24-2 shows the location of the APU in the Boeing 757.

RE 24-2n of ry Power



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ducted intakes. These normally takef the start sequence and close whenndicated at the APU control panel in

um chamber, from which air forPU gas turbine. In some units, a

amber for the aircraft's cabin air

fuel tanks via a remotely operatedly. The FCU controls accelerationload conditions, when the unit is

s through the unit gearbox and isparate APU starter battery is fitted.quence is basically the same for all



5. Air supply for the gas turbine compressor is admitted via the form of doors, which open automatically at the beginning othe APU is shut down. The APU intake door positions are often ithe flight compartment.

6. The APU gas turbine compressor discharges to a plencombustion is supplied to the combustion chambers of the Aregulated supply of bleed air is ducted from the plenum chconditioning and main engine air-starting systems.

Fuel System7. The APU fuel is supplied from one of the aircraft main(solenoid) valve. A fuel control unit (FCU) regulates fuel suppduring starting and maintains constant speed, under varying running.

Starting and Ignition8. The APU utilises an electric starter motor, which driveusually powered from the aircraft batteries. In some cases a seThe ignition unit is of the high-energy type and the starting setypes of APU.



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t (via one of the landing gear squat start principle. The in-flight starthigher airspeed being required thane APU inlet duct door arrangementffect to windmill the engine up to anergising of the igniter circuits.

two-engined aircraft) the available not depleted by what may be annative source of generated electrical in sequencing the APU inlet duct

ged period of flight at high altitude,

the aircraft flight deck. The startingle or push type). This opens the airfuel supply and ignition controls areeved. A typical APU control panel is



9. In some systems the APU starter motor is isolated in flighswitches), and in-flight starts are achieved using the windmillenvelope for the APU will be narrower (a lower altitude and a with a motor-assisted air-start) with this system. Furthermore, thneeds to be more complex in order to provide increased ram air eadequate rotational speed for the introduction of fuel and the en

10. However, in the event of a double generator failure (on aemergency source of electrical power (the aircraft batteries) isunsuccessful attempt to start the APU in order to provide an alterpower. The only electrical power required being that involveddoors.

11. APU’s are notoriously difficult to start following a prolonwhich has resulted in well and truly cold-soaked lubricating oil.

12. The main APU control and indication panel is located in sequence is initiated by closing the master control switch (toggintake doors and the starter motor then motors up the APU. The activated as and when the appropriate rotational speeds are achishown at Figure 24-3.



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FIGUTypicaPowerContro



RE 24-3l Auxillary Unit l Panel



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with the continued aid of the starterthis point the starter motor circuit iser its own turbine power. At aboutcentrifugal switch and combustion

a minute or so, to allow all parts tond pneumatic system air loads. AnPU if governed speed is exceeded,

at light up and governed speed ares not occur. The number of re-start maintenance manuals.

mss-wire fire detection system and itsting visual and aural warnings, thely activate the APU shut down. The on the APU control panel. In some

detection circuit is activated.

ooling and ventilation of the APUil cooler.



13. After light up has taken place the APU engine accelerates,motor, typically to between 35% and 50% of governed rpm. At de-energised and the engine accelerates to governed speed und95% governed speed the ignition circuit is cancelled by the becomes self-sustaining.

14. It is good practice to allow the unit to run at no load for reach normal working temperatures, before selecting electrical aoverspeed sensing system will automatically shut down the Atypically when speed reaches 110% of governed speed.

15. During starting the APU must be monitored to ensure thachieved within specified time limits, and that overheating doeattempts permissible is also specified in the aircraft operating and

Fire Protection and Cooling Syste16. The APU compartment is fitted with its own continuouown single-shot fire extinguishing system. In addition to activadetection circuit is normally arranged so that it will automaticalAPU fire extinguisher is activated by manually operated switchesAPU’s the fire extinguisher is discharged automatically when the

17. A fan, driven from the accessories gearbox, provides ccompartment and cooling air for the generator and lubricating o



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(or STOP) position on the masterr about two minutes to assist with

result of any or all of the following

.

exceeded.

switch should be selected OFF (or



Shutting Down18. The APU is normally shut down by selecting the OFF control switch, after first allowing the unit to run on no-load foeven cooling of the engine.

19. Automatic shut down of the APU will typically occur as aconditions:

(a) Overspeed (typically 110% of governed rpm).

(b) Excess exhaust gas temperature (EGT).

(c) Loss of EGT signal to the control system.

(d) Low lubricating oil pressure.

(e) High lubricating oil temperature.

(f) APU compartment fire detection system operation

(g) Excess APU bleed air outlet temperature.

(h) When specified airspeed or altitude limitations are

20. Following an automatic shut down the master control STOP).



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g and automatic shut-down systemided externally under a panel in thermit APU shut-down to be initiatedill be an APU fire warning klaxon.n the flight deck is unoccupied.

r requirement. Some APU’s have a governed by associated equipmentimit may be lower than the aircraft’sudes its output may be restricted ie.9,000 ft.

eumatic/electric or both) exceed theto protect the APU from exceedingld be more critical. In conditions ofPU EGT will reduce the bleed air

air-driven turbine that drives anine-driven alternators.

ct drive to the alternator, which isrs and capable of meeting essentialtment closed by a hinged door.



21. Depending upon the complexity of the APU monitorinprovided, a second APU control panel may or may not be provaircraft skin (normally adjacent to the APU itself), which will pefrom outside the aircraft. Associated with this external panel wAppreciate that it is common for the APU to be left running whe

22. Operating envelopes for APU’s vary according to type orestriction on the maximum altitude for starting, which may belimitations i.e. battery or by the starting envelope. The altitude lmaximum ceiling. Whilst an APU can be operated at higher altita generator may go from 100% output at 25,000 ft to 60% at 3

23. Should the demand on systems provided by the APU (pndesigned APU load, automatic control devices will be activated engine operating temperature limits (EGT). Higher altitudes couhigh electrical load and air bleed the limiting function of the Asupply but permit the APU to maintain generator output.

Ram Air Turbine24. Many commercial transport aircraft are fitted with anemergency alternator, for use in the event of failure of all the eng

25. A typical unit comprises a single stage turbine with direusually of lower output capacity to the engine driven alternatoelectrical requirements. The unit is normally stowed in a compar



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deployed into the air stream, whereVariable incidence inlet guide vanesd is maintained by variation of ther.

used as the driving unit and certainump.

wing root fairing and is deployedom the flight deck.



26. When activated, the door opens and the ram air turbine isit is rotated by the airflow passing through the turbine blades. control the airflow into the turbine wheel. Constant rotary speeinlet guide vanes under the influence of a flyweight type governo

27. In some instances a variable pitch two-bladed propeller isaircraft have a ram air turbine driving an emergency hydraulic p

28. Typically, the unit is located in the underside of the mechanically by spring action when a release catch is activated fr




Aeroplane

hing


Emergency Equipment-

Doors and Emergency Exits

Passenger Emergency Exits

Crew Emergency Exits

Accessibility

Normal And Emergency Operation

Door Markings

Cut-In Areas

Smoke Detection

Fire Protection Systems

Fire Detection Systems

Automatic Toilet Fire Extinguishers

Cargo Compartment Fire Extinguis

Oxygen Systems




itter (ELT)


Continuous Flow Oxygen Systems

Diluter Demand Oxygen Systems

Chemical Oxygen Generators

Masks

Indications

Cylinder Charge Pressure

Diluter Demand Regulator

Normal Operation

Emergency Equipment

Portable Fire Extinguishers

Smoke Hoods

Portable Oxygen Sets

Emergency Locator Beacon/Transm

Life Jackets

Life Rafts



entory


Emergency Torches

Emergency Lighting

Megaphone

Crash Axe

Fireproof Gloves

Emergency Equipment – Typical Inv



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eroplane

g aircraft must be equipped withuations.

ntry doors are situated at the frontr capacity will determine the totalare designed to be capable for an

exit in the cabin which is usually aindicate operation for normal and

crew would use the passenger cabint the flight deck crew from reachingm the flight deck. This can take the the fight deck. Because of the longn inertia reel or rope system which,olled rate.

Emergency Equipment-Aeroplane


25Emergency Equipment-A

Doors and Emergency Exits1. Statutory regulations require that all passenger-carryinadequate crew and passenger exits for normal and emergency sit

Passenger Emergency Exits2. In most large passenger transport aircraft normal exit/eand rear of the passenger cabin. Cabin length and passengenumber of doors located either side of the cabin. All doors emergency evacuation – some aircraft may have an over-wing removable window. All doors/window are clearly marked to emergency use.

Crew Emergency Exits3. In the event of an emergency evacuation, the flight deck exits as their primary means of escape. If circumstances preventhe cabin they are provided with an alternative means of exit froform of either an opening side window or a hatch in the roof ofdrop from this position each crew member will have available awhen grasped, will allow the crew member to descend at a contr



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ency situation it is paramount thatd. The manufacturer will design thet slow down the flow of passengers

sure that any items brought on byired pathways. This would meanred in the spaces provided once onall exits is a serious and mandatory

nther inward or outward from thef the plug type, in which the cabinof this type of door is only achievedanually or electrically from outside

leasing the mechanical locks, beforey a motor supplied from the aircraftvent of normal system failure.

d outside of the door.

m to power the door open when



Accessibility4. To ensure swift evacuation from the aircraft in an emergthe route to all emergency exits is completely clear and unimpedeinterior fixtures and fittings to ensure a clear path that does noproceeding to the exit(s).

5. It is then left to the operator to maintain this and to enpassengers are not placed where they would impede the requcontrolling the items carried on and ensuring that they are stoboard the aircraft. The provision of an unimpeded pathway to requirement and must not be ignored.

Normal And Emergency Operatio6. There are a variety of aircraft door types, opening eipassenger cabin. In modern passenger aircraft the doors are odifferential pressure keeps the door closed and sealed. Opening when differential press is zero. The doors can be opened either mor inside. Manual operation is achieved by moving a handle, repushing/pulling the door open. Electrical operation is achieved belectrical system, backed up by the standby electrical system in e

7. Method of operation is clearly displayed on the inside an

8. Some large, heavy doors have a pneumatic assist systeoperated in the auto emergency situation.



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cing the closed door into the ‘arm’in the cabin floor. When the doorflation mechanism of the emergencyy combine the use of the slide and occupants to vacate the aircraft. Ifal means of operation is provided.sets off the bottle firing mechanism.

ystem will be disarmed if the door isbe used a lever must be operated tothe cabin). Exit is then through thermed at commencement of taxi-out

he inside and outside to indicateon have an exit sign adjacent to thetive quality. Operating instructions

ompanied with pictorial guides i.e.

e visually displayed by a protuding

ngers to evacuate the aircraft in 90



9. Emergency operation of the door is achieved by first plaposition. This action operates a bar which engages with lugs handle is operated to open, the locked bar will trigger off the inescape chute, or slide. Some slides have a dual role where thefunction as a life raft. The slide will inflate and deploy to allowthe slide inflation bottle fails to operate automatically a manuThis is usually a handle or lanyard which is pulled by hand and If the bottle fails to function the slide cannot be used.

10. Irrespective of the position of the door arming lever, the sopened from the outside of the cabin. If the overwing exit is to unlock a removable window section (usually pulled inwards to aperture and onto the wing surface. Door slides are normally aand disarmed just prior to reaching parking area.

Door Markings11. All doors and exit positions are clearly marked on toperation. All doors and windows used for emergency evacuatifacility. The sign may be illuminated or it may have a light reflecfor normal and emergency use are printed boldly and are accarrows indicating direction of handles/switches.

12. Indication that a door is in the auto armed position can bpin or the position of the arming lever/switch.

13. Evacuation routes must be capable of enabling all passeseconds.



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ncy lighting in the B757.

FIGUDoorsEmergEquipm



14. Figure 25-1 shows the location of exits, slides and emerge

RE 25-1, Exits & ency ent



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follows;

ssible external door

m opening in flight (by person orh door must be capable of beingif persons are crowded against the can be either inward or outwardsimple and obvious. The markedght.

jamming as a result of fuselage

be so located that persons using

is not inward must have provisionechanism to determine if its fully

must be provided to warn the flighted. For doors with initial openingld be totally free of any erroneous

locked it should be such that anyossible. It must also be shown that.



JAR-OPS Requirements - Doors and Emergency Exits

The JAR-OPS requirement for door and emergency exits are as

(a) Each cabin must have at least one easily acce

(b) Each external door must be safeguarded frofailure of a single structural element). Eacopened from both inside and outside, even door on the inside of the aeroplane. Doorsopening. The means of opening must be instructions must be easily read by day or ni

(c) Each external door must be free from deformation in a minor crash.

(d) On propeller driven aeroplanes, exits mustthem are not endangered by the propellers.

(e) Any doors, whose initial opening movementfor direct visual inspection of its locking mclosed and locked. A visual warning system crew of any door not fully closed and lockmovement, not inward, the indication shouclosed and locked indication.

(f) If an external door is not fully closed or attempt of initiation of pressurisation is impinadvertant opening is extremely improbable



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fuselage must be either a Type A,it. Each emergency exit must beselage deformation) with either theone or more of the landing gearsation to fully open.

ith passenger emergency exit sill

ft) above the ground with the ding gear extended; or

tres (6 ft) above the ground after of, one or more legs of the landing e was first applied for on or after 1

sub-paragraphs (i) or (ii) apply, toergency.

rovided at overwing exits if theture at which the escape routem the landing gear extended, and, whichever flap position is higher

ergency exit for the flight crew and:



(g) Each passenger entry door in the side of theType 1 or Type 2 passenger emergency excapable of being opened (when there is no fuaeroplane in the normal ground mode or collapsed, within 10 seconds from door actu

Means for Emergency Evacuation

(a) An operator shall not operate an aeroplane wheights:

(i) Which are more than 1.83 metres (6aeroplane on the ground and the lan

(ii) Which would be more than 1.83 methe collapse of, or failure to extend gear and for which a Type CertificatApril 2000.

unless it has equipment or devices available at each exit, whereenable passengers and crew to reach the ground safely in an em

(b) Such equipment or devices need not be pdesignated place on the aeroplane structerminates is less than 1.83 metres (6 ft) frothe flaps in the take off or landing positionfrom the ground.

(c) In aeroplanes required to have a separate em




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e fatally delayed by distorted and areas are marked on the outside ofHere in Emergency’ in capital letterse cut through with relative ease, todesign of cut-in panels is controlled

e fuselage suitable for break-in byeas shall be marked as shown below.ry they shall be outlined in white toe than 2 metres apart, intermediate2 metres between adjacent marks.

emergency exit is more than 1.83the landing gear extended; or

irst applied for on or after 1 Aprilres (6 ft) above the ground after thene or more legs of the landing gear,

in descending to reach the ground



Cut-In Areas15. Access to a crashed aircraft by rescue services could bjammed doors. To allow for prompt access, rectangular break-inthe fuselage by right-angled corner markings and the words ‘Cut across the centre. These indicate areas of structure which can bprovide ‘door-sized’ holes in the fuselage. The size, colour and by law and is shown in Figure 25-2.

16. An operator shall ensure that, if designated areas of threscue crews in emergency are available on an aeroplane, such arThe colour of the markings shall be red or yellow, and if necessacontrast with the background. If the corner markings are morlines 9 cm x 3 cm shall be inserted so that there is no more than

(i) For which the lowest point of the metres (6 ft) above the ground with

(ii) For which a Type Certificate was f2000, would be more than 1.83 metcollapse of, or failure to extend of, o

there must be a device to assist all members of the flight crewsafely in an emergency.




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FIGUCut-In

6 or more.

e detector system that provides ar audible warning in the cabin that

o compartment fire detection the

ation to the flight crew within one



RE 25-2 Areas

Smoke DetectionJAR-OPS Requirement Related to Smoke Detection

(a) Lavatory. For aeroplanes with a passenger capacity of 2

(i) Each lavatory must be equipped with a smokwarning on the flight deck or provides a visual owould be readily detected by the cabin crew.

(b) Cargo compartment. For aircraft certified with cargfollowing must be met for each compartment.

(i) The detection system must provide a visual indicminute after the start of a fire.



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ys are fitted with smoke detectorsf certain parameters are exceeded.

ost passenger aircraft. The detector automatic and operates from thereen light to indicate that power isction with the aural warning when

ning, but the alarm will sound again

a beam of light which occurs whenttering of the light increases the to operate a warning circuit.

chambers which measure alpharbed by smoke, which reduces them.

t a temperature significantly belowplane is substantially decreased.

the functioning of each fire

shown for all approved operating



17. Freight holds, baggage compartments and equipment bawhich sample the air in the compartment and activate an alarm i

18. Smoke detectors are fitted in the toilet compartments of mprovides an aural warning to alert the cabin crew, it is fullyaircraft’s 28v DC power supply. The detector unit displays a gbeing supplied to it and a large red display illuminates in conjunsmoke is detected. A reset switch enables cancellation of the warif smoke is still present. Smoke detectors are of four main types:

(a) Photoelectric cells. These detect the diffusion of the beam is interrupted by smoke. The scaconductance of the cell and its output is amplified

(b) Alpha particle detectors. These are ionisationradiation from radium. Alpha particles are absoionisation current of the device, to operate an alar

(ii) The system must be capable of detecting a fire athat at which the structural integrity of the aero

(iii) There must be a means to allow the crew to testdetection circuit.

(iv) The effectiveness of the detection system must beconfigurations and conditions.

JAR-OPS Requirement Related to Smoke Detection



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tted as alarm verification devices.

aircraft of American manufacture,vate a warning system.

deck by illuminating a master firee case of multiple areas for detection

e installations will incorporate annal test facility will be provided ton systems are sometimes subject to

tamination other than smoke in thetegories of freight, especially in all

gent, the rate of discharge and the in the designated zone. For mostncentration.

o prevent bursting of the container.er to indicate that the container hasecessary for operation. If any toxicnted from entering any personnelately or by defect.



(c) Visual smoke indicators. These are usually only fi

(d) Carbon monoxide detectors. Found mainly in these devices detect concentrations of CO and acti

19. Activation of a smoke detector is indicated on the flightwarning light and a red warning light for the affected area. In tha panel with a specific light for each area will be provided. Somaural warning in addition to the visual indications. A functioenable the crew to check the system at any time. Smoke detectiofalse warnings. This can occur if the detector is exposed to consampled air e.g. dust, sand or impurities given off by certain cacargo aircraft. (Live animals).

Fire Protection SystemsJAR-OPS for Fire Extinguishing Equipment

A fire extinguishing system, the quantity of the extinguishing adischarge distribution must be adequate to extinguish the firefirezones, two discharges must be provided of adequate agent co

Each extinguishing agent container must have a pressure relief tThere must be a means for each fire extinguishing agent containdischarged or that the charging pressure is below the minimum nagent is used, harmful concentration of liquids must be prevecompartments as a result of the extinguisher discharging deliber



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ystems incorporate a fire protectionion system is provided on the engineve warning on the flight deck. The located in strategic positions, or a

tail at the end of this section.

a steady RED warning light and an the fire and the alarm bell can bengine is covered by its own warningten doubled up, either by the use offire bottle.

nel by means of discharge switches.e aircraft fire extinguishing system.switch is pulled, certain automatic

valves



20. All gas turbine engines and their associated installation ssystem for the detection and rapid extinguishing of fire. A detectto sense an overheat condition or the occurrence of a fire and gidetector system can consist of either a number of detector unitscontinuous sensing element. The latter is discussed in greater de

21. The occurrence of a fire is indicated on the flight deck byalarm bell. The red light will usually indicate the location ofsilenced by a cut-out switch, leaving the light remaining. Each esystem and is provided with an extinguishing system, which is ofmore than one fire bottle or the optional use of another engine's

22. The system is controlled from the pilot's fire control paFigure 25-3 shows a typical schematic layout for a twin-enginDepending on aircraft type/engine type, when the engine fire actions are triggered, i.e:

(a) Closure of the hydraulic shut-off valve

(b) Closure of the pneumatic system supply bleed air

(c) The fire bottle squibs are armed

(d) The generator field control relay is tripped

(e) The fuel system tank shut-off valve is closed.



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FIGUTwin EExtingSystem

the fire bottle and its discharge line.ge which is electrically detonated to



RE 25-3ngine Fire uishing

23. A discharge device, often called a squib, is fitted between The squib consists of a breakable disc and a small explosive charbreak the disc and discharge the contents of the bottle.



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a fire is detected in the number oneb in each fire bottle. Rotating thed fires its left squib, discharging thekwise will fire the number two fire

o extinguisher is not used it remainsitch will arm the right engine bottlearge the number two bottle into the

Pulling the APU fire switch arms the is normal to wait 30 seconds beforeead system it must be borne in mindtection for the other engine.

stallations consisting of pressurisedrols. The types of extinguishant areochlorodiflouromethane (BCF) and

bottle and a discharge head through by remote operation of electricallyuit is controlled by switches or teeutomatically initiated in the event ofnit which, when fired, ruptures aorce a plug to the end of a tube,



24. Two fire extinguisher bottles are installed. Let us supposeengine. Pulling the LEFT fire switch arms the left engine squiswitch anti-clockwise selects the number one fire extinguisher anbottle into the left engine. Subsequently rotating the switch clocextinguisher into the left engine. Alternatively, if the number twavailable for use on the right engine. Pulling the RIGHT fire swsquibs and turning it clockwise will fire its right squib and dischright engine.

25. A single fire extinguisher bottle is provided for the APU. squib, which is fired by rotating the switch in either direction. Itcontemplating the use of a second shot. In the case of the dual hthat the use of the second charge will deprive the pilot of fire pro

26. Power plants and APU's use fixed fire extinguishing inextinguishant containers, distribution piping and operating contusually toxic or semi-toxic Freon compounds such as brombromotriflouro-methane (BTM).

27. Fixed extinguishers normally consist of a steel or copper which the extinguishant is discharged to the distribution systeminitiated cartridge units in the discharge head. The firing circhandles on the flight deck. In some aircraft types firing may be aa crash landing. The discharge head contains a cartridge udiaphragm and allows the pressure of the extinguishant to fconnecting the bottle to the distribution system.



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ly flush with the tube end cap. Thep once the cartridge has been fired,discharged. In some aircraft thetate of charge. Figure 25-4 shows a



28. At the end of the discharge tube is a pin which is normalmovement of the plug causes the pin to protrude from the end caproviding a visual indication that the extinguisher has been extinguisher bottle incorporates a pressure gauge to indicate the sfixed fire extinguisher.



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FIGUFixed FExting



RE 25-4ire

uisher



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le, due for example to overheating,es under high temperature or burstsand discharges overboard through apically on the outside of the enginedicators are usually in the form of a. Alternatively, it may be arranged

ue to excess pressure. With thistle.

ht deck to show when extinguisher present in the distribution system.

fire or overheat detectors in eachailpipe sections of turbine enginensuring prompt detection of fire in

installed so that:

loads during operation.

ed of any severance of a detectionerative.

of any short circuit of a detection erative.



29. In the event of a build-up of excess pressure in the bottprotection against bursting is provided by a disc which either fusunder excess pressure. The disc is situated in the discharge head pressure relief line. An indicator located in a visible position, tynacelle, will be exposed by the pressure in the relief line. The incoloured disc, red in most cases, but occasionally green or yellowfor the disc to be blown out when the bottle discharges darrangement, the absence of the disc warns of a depleted fire bot

30. Some aircraft incorporate electrical indicators on the fligcartridge units have been fired or that extinguishant pressure isThese may be magnetic indicators, or warning lights.

Fire Detection SystemsJAR-OPS Requirements for a Fire Detector System

JAR-OPS states that there must be approved, quick acting designated fire zone, and in the combustion, turbine, and tinstallations. They must be in sufficient numbers and locations ethose zones. Each fire detector system must be constructed and

(a) It will withstand vibration, inertia and other

(b) There is provision for the crew to be warnsystem where it would render the system inop

(c) There is provision for the crew to be warned system where it would render the system inop



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sually divided into fire zones. Thoseg and extinguisher systems. Others,ning systems.

lly protected by smoke detectioncess temperature detectors.

and extinguishing equipment to the

er fluid or fumes present in their

the system at any time.

the fire zone must be fire resistant.

nt may pass through another fire

lse warnings resulting from fires in

ly protected by the same detector

e approved alarm actuation time.



31. On civil transport aircraft the engine compartments are uin which the likelihood of fire is greatest are protected by warninsuch as jet pipe surrounds, may only be fitted with overheat war

32. Equipment bays and baggage compartments are usuaequipment and areas adjacent to hot air ducts usually contain ex

33. Auxiliary Power Units (APUs) have similar fire detectionmain engines, but usually incorporating automatic shut-down.

(d) A detector is not affected by oil, water or othlocation

(e) The crew can carry out a functional check of

(f) Wiring components of any detector system in

(g) No fire or overheat detector system componezone unless;

(i) It is protected against possibilty of fazones through which it passes or

(ii) Each zone involved is simultaneousand extinguishing system

(h) Each fire detection system must not exceed th

JAR-OPS Requirements for a Fire Detector System



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tions on the flight deck and often steady red indication. All detectionsophisticated type which monitor

or each engine, but the warning bell

and continuous types.

or switches which are operated byitor specific points where excessivend a potential fire zone to provide

hich the central conducting core isve material. These detectors may be

in resistance of the insulation withow from core to sheath and activateort-circuit between core and sheath

fire warning.

citance which occurs with increasedcharge, with increased temperature activate the warning circuit. If acapacitor, but does not produce a



34. Fire detector signals activate warning lamps and/or capaudible warnings also. Fire warning lamps conventionally give asystems include functional test circuits and many are of a temperature trends in engine bays. There is one warning lamp fwill be activated by any detection circuit.

35. Detection equipment falls into two main categories, unit

36. Unit type detectors usually employ either thermocouplesdifferential expansion of metals. Unit detectors are used to montemperatures might occur, continuous detectors are routed aroumaximum coverage.

37. Continuous wire detectors consist of a co-axial cable in winsulated from the outer, earthed, sheath by a temperature sensitiof either the capacitive or resistive variety.

38. Resistive continuous detectors make use of the decreaseincreasing temperature, which will eventually allow current to fla warning circuit. The disadvantage of these detectors is that a shdue to crushing or chafing will cause them to initiate a spurious

39. Capacitive continuous detectors use the increase of capatemperature. The increase of stored charge, and therefore discreates a back emf and current which eventually is sufficient tocapacitive detector is short-circuited it may cease to act as a spurious fire warning.



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rsautomatic fire extinguishers locatedesigned to discharge when a pre-set freon gas through one, or both, oftowel disposal container, the others changes to aluminium when thecated on the inside of an access doorurn black when exposed to highp has changed colour it should bed placard renewed.

ishing bottles are provided for fire controls detected in either compartment theing arming switch arms the selected number one fire extinguisher bottlement with the overheat warning, byne cargo bottle discharged warning indicate low bottle pressure. The whereupon the discharge light andswitch is provided for checking the



Automatic Toilet Fire Extinguishe40. Most modern passenger transport aircraft are fitted with under the wash basin in the toilet compartments. The system is dtemperature is exceeded. The extinguisher discharges non-toxictwo heat-activated nozzles. One nozzle discharges toward the directly under the wash basin. The colour of the nozzle tipextinguisher is discharged. A temperature indicator placard is lobelow each wash basin. White dots on the placard will ttemperature. If an indicator has turned black or a nozzle tiassumed that the extinguisher has discharged and extinguisher an

Cargo Compartment Fire Extingu41. In the system shown at Figure 25-5, two fire extinguisherof the forward and aft cargo compartments. When an overheat iassociated fire warning light illuminates. Pressing the correspondsquib (FWD or AFT) in each of the two extinguisher bottles. Theis the larger of the two and is discharged first, into the compartpressing the number one bottle discharge switch. The number olight illuminates and a CARGO BTL 1 caption is displayed tonumber two bottle may be manually discharged at a later stage,advisory message appears for the second bottle. A squib test electrical continuity of the squib firing circuits



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FIGUCargoExtingSystem



RE 25-5 Fire uishing



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method of operation and its effect in direct contact with oxygen maymust be protected agains incorrectved hand should be used. Wheret be complete and sealed beforeto both crew and passengers the

on duty and a separate supply for or

ans to separately reserve the ht crew on duty.

of continuous flow, diluter demand to meet the crew or passenger



Oxygen SystemsJAR-OPS Requirement for Aeroplane Oxygen Equipment

The oxygen system must be free from hazards in itself, in its upon other components e.g. no material should be used which,give off noxious or toxic gases. Any couplings or connectors assembly and if required for connection, disconnection a gloelectrical connections are combined, the oxygen circuit muselectrical connection is attempted. When oxygen is supplied distribution system must be designed for either

(i) a source of supply for the flight crewpassengers and other crew members

(ii) a common source of supply with meminimum supply required by the flig

(iii) Portable walk around oxygen units or straight demand may be usedrequirements



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h demand equipment and a quickld be able to be placed on the facen with one hand within 5 secondsy duties. Normal communicationent must be immediately availableew to determine whether oxygen is). Pressure limiting devices (relief protect the system from excessive

it for each occupant for whomst be designed to cover the nose andans to retain the unit in position onygen must have provisions for the



Each crew member on flight deck duty must be provided witdonning mask connected to an oxygen supply. The mask shoufrom its stowed position, secured, sealed and supplying oxyge(without disturbing eyeglasses or causing delays in emergencduties must be able with mask donned. Portable oxygen equipmfor each cabin attendant. There must be a means to allow the crbeing delivered to the dispensing equipment (flow indicatorvalves) and protective devices (rupture disc) should be fitted topressure. Venting should be overboard.

JAR-OPS Equipment Standards for Oxygen Dispensing Units

If oxygen dispensing units are installed, the following apply:

(a) There must be an individual dispensing unsupplemental oxygen is to be supplied. Units mumouth and must be equipped with a suitable methe face. Flight crew masks for supplemental oxuse of communication equipment.

JAR-OPS Requirement for Aeroplane Oxygen Equipment



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25,000 ft is requested, an oxygenh the unit of oxygen dispensingwhich ensures that the oxygen isch of each crew member. For anysing equipment must be located toapplicable National Operational

ulations, if certification for st be oxygen dispensing equipment

unit connected to oxygen supplych occupant, wherever seated. If00 ft is requested, the dispensingw must be automatically presentedsure altitude exceeds 15,000 ft andual means to make the dispensing of failure of the automatic system.nd outlets must exceed the numberts must be as uniformly distributed



(b) If certification for operation up to and includingsupply terminal, either a supply terminal witequipment already connected or a connection immediately available, must be within easy reaother occupants the supply terminals and dispenallow use of oxygen as required by the Regulations.

(c) Except as specified in National Operational Regoperation above 25,000 ft is requested, there mumeeting the following requirements:

(i) There must be an oxygen dispensing terminals immediately available to eacertification for operation above 30,0units providing the required oxygen floto the occupants before the cabin presthe crew must be provided with a manunits immediately available in the eventThe total number of dispensing units aof seats by at least 10%. The extra unithroughout the cabin as practicable.

JAR-OPS Equipment Standards for Oxygen Dispensing Units



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conditions inside the cabin equal toaft altitude above this figure. Underand crew, but oxygen equipment isfailure.

ment may be installed for the use ofove 10,000 feet. Where no oxygen

ten also provided in large transport during pressurisation emergencies.

ow or diluter demand type, or at types. Gaseous oxygen is stored in

nders are provided with an excessenting the cylinder contents to the in the cylinder. In most cases ane to excess cylinder pressure.

rs are used for passenger oxygenies of pressurised oxygen.

. Gaseous oxygen is supplied to the reduced pressure to the distribution or the drop out type.



42. Most civil transport aircraft are pressurised to maintain an altitude of approximately 8000 feet, regardless of actual aircrthese conditions oxygen is not normally needed for passengers installed for emergency use in the event of pressurisation system

43. In passenger aircraft without pressurisation, oxygen equippassengers and crew when it is necessary for the aircraft to fly absystem is fitted, portable oxygen sets are provided, these are ofaircraft for therapeutic purposes and for use by cabin attendants

44. Gaseous oxygen systems may be of the continuous flcombination of the two, especially on the larger transport aircraf

cylinders at approximately 1800 lb/in2. Oxygen storage cylipressure rupture disc, fitted to the shut-off valve body and voutside of the aircraft in the event of a dangerous pressure riseindicator is fitted which will show that discharge has occurred du

45. In many transport aircraft chemical oxygen generatosupplies, since these obviate the need for storage of large quantit

Continuous Flow Oxygen Systems46. A continuous flow oxygen system is shown at Figure 25-6pressure reducing valve from the storage cylinder and thence at asystem and mask connection points. These may be of the plug in



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FIGUContinOxyge

k supply tubes are connected to thee embody a selector lever for thiseach mask connection, one givingrporate self-closing shut-off valves

by plugging in the mask connecting



RE 25-6uous Flow n System

47. With plug in systems the points at which the oxygen massystem usually provide for ‘normal’ or ‘high’ flow rates. Sompurpose, whilst others have two alternative socket points for normal flow and the other high flow. The socket points incowhich are spring-loaded to the closed position and are opened tube.



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d from the stowed position to theism, activated when cabin altitudeitch on the flight deck. When the

erated valve opens to supply oxygen

n used in light aircraft which only the functions of shut-off valve and plug-in variety.

5-7. Oxygen is diluted with air andle, that is to say only when the user and are additional to the passengermember.

n control lever, which controls the

mber's mask when the user inhales.sed as cabin altitude increases until,s supplied.

oxygen is supplied when the user



48. With drop out systems the mask is automatically ejecte"half hang" position by an aneroid controlled release mechanexceeds about 14,000 feet, and over-ridden, if necessary, by a swmask is further pulled towards the passenger's face a lanyard-opto the mask.

49. A simpler version of the continuous flow system is oftecarry a pilot and perhaps five passengers. In this type of systempressure regulator are combined and mask connections are of the

Diluter Demand Oxygen Systems50. A diluter demand oxygen system is illustrated at Figure 2the mixture is supplied as demanded by the user's respiration cycinhales. Diluter demand systems are provided only for crew useoxygen system. There is a mask connection point for each crew

51. The diluter demand regulator incorporates a four-positiooxygen flow to the crew masks as follows:

Normal oxygen. Diluted oxygen is supplied to the crew meThe proportion of oxygen in the mixture is automatically increaat a cabin altitude of 32,000 feet, approximately 100% oxygen i

100% oxygen. The regulator air valve is closed and 100%inhales, regardless of cabin altitude.



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other noxious fumes in the cabined at a pressure greater than cabin the flow is continuous (as opposed

at when the emergency position ise mask and equipment for leakage.

FIGUDiluteOxyge



Emergency. Used to provide protection against smoke oratmosphere. When selected, a flow of 100% oxygen is supplipressure, to prevent any leakage into the mask. In some systemsto demand) when the emergency selection is made.

Test mask. Oxygen is supplied at a higher pressure than thselected. This is used to test the mask for facial fit and to test th

RE 25-7r Demand n System



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FIGUChemiGener

tor.

barometrically ejected when cabinn a mask is pulled towards the user sodium chlorate and iron powderical reaction creates a flow of lowormally be maintained for about 15e oxygen itself is at a comfortableed mechanically. Chemical oxygen



Chemical Oxygen GeneratorsRE 25-8cal Oxygen ator

52. Figure 25-8 shows a diagram of a chemical oxygen genera

53. The unit consists of (typically) three drop out masks altitude exceeds a preset value (usually about 14,000 feet). Whea lanyard trips the electrical firing mechanism, which ignites acharge block. As the temperature of the block rises, a chempressure oxygen through a filter to the mask. Oxygen flow will nminutes and, despite the very high temperatures generated, thtemperature. In some sets the electrical firing circuit is initiatgenerators have a shelf/service life of ten years.



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m leakage around the mask, andhe aircraft communication system.ich deploy automatically (drop-outor may need to be physically held inployment the masks may be plastic

sually on a cockpit overhead panel.e pressure warning lights illuminatesplayed instead of cylinder pressure.

in2 at 21°C. At higher or lowerarging table is provided, giving therature.

nel system diluter demand oxygen



Masks54. Crew masks fit snugly to the user's face, with minimuincorporate a microphone and jack-plug for connection to tPassenger masks are usually cup-shaped rubber mouldings whmasks). They may be provided with a simple elastic head strap, place. In passenger systems which do not embody automatic debags with a head strap.

Indications55. Storage cylinder and supply line pressures are indicated, uCylinder pressures are usually displayed on pressure gauges, linwhen a system is in use. In some systems cylinder contents are di

Cylinder Charge Pressure56. Oxygen cylinders are charged to a nominal 1800 lb/temperatures the cylinder pressure will be higher or lower. A chpressure to which the cylinder must be charged at ambient tempe

Diluter Demand Regulator57. Figure 25-9 shows a schematic layout of a narrow paregulator.



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FIGUDiluteRegula



RE 25-9r Demand tor



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ygen Selection Lever to NORMAL,es a pressure differential across the, and admit oxygen from the openeroid-controlled air metering valve increased the reducing barometricr supply and increasing the oxygent to open when the user inhales and

% the air inlet port is closed and the oxygen to the outlet whenever the

ion, the demand valve is held opener responds to differential pressure to the outlet and oxygen supply

anel oxygen regulator.



Normal OperationFor normal operation the Emergency Lever is set to OFF, the Oxand the Supply Lever to ON. When the user inhales this creatdemand diaphragm, causing the demand valve to open partiallysupply valve to mix with air from the air inlet valve. The andetermines the oxygen/air ratio of the mixture. As altitude ispressure causes the aneroid capsule to contract, reducing the aisupply to the outlet. The air inlet valve is a non-return valve, secreates a pressure differential across it.

100% Oxygen. With the Oxygen Selection Lever set to 100supplementary oxygen valve is held open, connecting undiluteduser inhales.

Emergency. With the Emergency Lever set to the ON positand the demand diaphragm is mechanically loaded and no longwhen the user inhales. Undiluted oxygen flows continuouslypressure holds the air inlet valve closed.

58. Figure 25-10 shows a typical control panel for a narrow p



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FIGUOxygePanel

the oxygen flow and visible throughof view.

d containing, usually, 120 litres ofh depressurisation and paramedicalth straps so that it may be worn by by a flexible hose.



RE 25-10n Regulator

59. The flow indicator is in the form of a float suspended in a sight window. When no oxygen is flowing the float drops out

Emergency Equipment

Portable Oxygen Sets60. These consist of a cylinder, charged to 1800 ib/in2 anoxygen. They are provided for use by cabin crew in dealing witincidents. The cylinder is contained in a fabric carrying bag, withe bearer, and a mask connected to the cylinder regulating valve



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e cases) ‘emergency’ flow rates frome endurance of a 120 litre portableminutes respectively.

or BCF as the extinguishant. Thein an extinguishant suitable for the should be easily accessible and are

cal fires. The released CO2, gas namely heat and oxygen (the thirdamage if used on an engine fire.

re therefore widely used in aircraft.t it is tackling, it gives off noxiousis therefore necessary to employ aire, either with a water extinguishery only luke warm).

materials (furnishings, paper, woodires, since the water jet can conductr is also unsuitable for use on metalwater to break down into hydrogenfire where water extinguishers arend result in the spread of the fire.



61. It is usually possible to select ‘normal’, ‘high’ and (in somthese sets, corresponding to 2, 4 and 10 litres per minute. Thoxygen set at these flow rates would therefore be 60, 30 and 12

Portable Fire Extinguishers62. Portable extinguishers in aircraft use either CO2 WATERtype of extinguisher chosen for a particular location will contatype of fire to be expected in that compartment. Extinguisherstherefore retained in brackets by quick-release fittings.

63. CO2 extinguishers are particularly suitable for electriexcludes two of the three components required for combustion,being fuel). Because of its rapid cooling effect, CO2 can cause d

64. BCF extinguishers are suitable for all types of fires and aThe disadvantages of BCF are that, once heated by the fire thagases. Also, BCF does not cool the fire-affected area, and it follow-up action to cool the area and prevent re-ignition of the for a convenient coffeepot (as we know, airline coffee is invariabl

65. WATER extinguishers are suitable only for use on dry and so on). They should never be used on electrical equipment felectricity and may lead to electrocution of the fire-fighter. Watefires, since the high temperature of burning metal may cause the (a fuel) and oxygen. Liquid fuel fires are another type of unsuitable, since the burning fuel could well float on the water a



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tinguishers are provided for use ind galleys in accordance with the

t must be suitable for the kinds ofe the extinguisher is intended to be

t minimise the hazards of toxic gas

her, containing Halon 1211r equivalent as the extinguishingight deck for use by the flight crew;

cated in, or readily accessible for assenger deck;

guisher must be available for use in mpartment and in each Class E members in flight; and



JAR Requirements

JAR–OPS 1.790 Hand fire Extinguishers

An operator shall not operate an aeroplane unless hand fire excrew, passenger and, as applicable, cargo compartments anfollowing:

(a) The type and quantity of extinguishing agenfires likely to occur in the compartment wherused and, for personnel compartments, musconcentration;

(b) At least one hand fire extinguis(bromochlorodifluoromethane, CBrCIF2), oagent, must be conveniently located on the fl

(c) At least one hand fire extinguisher must be louse in, each galley not located on the main p

(d) At least one readily accessible hand fire extineach Class A or Class B cargo or baggage cocargo compartment that is accessible to crew



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extinguishers must be conveniently

, they must be evenly distributed in

uishers located in the passengerimum approved passenger seatinghan 60, and at least two of the firepartment of an aeroplane with auration of 61 or more must contain, CBrCIF2), or equivalent as the



(e) At least the following number of hand fire located in the passenger compartment(s):

(f) When two or more extinguishers are requiredthe passenger compartment.

(g) At least one of the required fire extingcompartment of an aeroplane with a maxconfiguration of at least 31, and not more textinguishers located in the passenger commaximum approved passenger seating configHalon 1211 (bromochlorodifluoromethaneextinguishing agent.

JAR–OPS 1.790 Hand fire Extinguishers



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shers should be such as to provideen of the number and size of the

imise the hazard of toxic gass etc.

eing greater than the minimum

itable for both flammable fluid andeck. Additional extinguishers mayrtments accessible to the crew inot be used on the flight deck, or inrom the flight deck, because of thef non-conductive interference with

d in the passenger compartments itstation, where provided.

s are required in the passengerwise dictated by consideration oflocated near each end of the cabinbin as evenly as is practicable.

on should be indicated by a placardpplement such a placard or sign.



AMC Requirements

AMC OPS 1.790 Hand Fire Extinguishers

(a) The number and location of hand fire extinguiadequate availability for use, account being takpassenger compartments, the need to minconcentrations and the location of toilets, galley

(b) These considerations may result in the number bprescribed.

(c) There should be at least one fire extinguisher suelectrical equipment fires installed on the flight dbe required for the protection of other compaflight. Dry chemical fire extinguishers should nany compartment not separated by a partition fadverse effect on vision during discharge and, ielectrical contacts by the chemical residues.

(d) Where only one hand fire extinguisher is requireshould be located near the cabin crew member’s

(e) Where two or more hand fire extinguishercompartments and their location is not otherparagraph 1 above, an extinguisher should be with the remainder distributed throughout the ca

(f) Unless an extinguisher is clearly visible, its locatior sign. Appropriate symbols may be used to su



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smoke hood as provided in many current usage and is a light-weight respiratory and eye protection in

mprises a single size hood whichchest contains a life support pack ineroxide. This chemical reacts withroduce pure oxygen. The chemical

his takes a few seconds to start andtly a quick-start toggle is located ates oxygen for the wearer until the



Smoke Hoods66. The description which follows is of the Drager Oxycrewaircraft. It is reasonably representative of most smoke hoods inportable breathing device, designed to provide the wearer withoxygen deficient, smoke-filled or toxic atmospheres.

67. A Drager mask is illustrated at Figure 25-11 and cocompletely covers the head. An apron extending down over the the form of a chemical oxygen canister filled with potassium supwater vapour and carbon dioxide in the wearer’s exhaled air, to paction can be initiated by simply breathing into the mask, but tthese may not be available in an emergency situation. Consequenthe bottom of the apron. When pulled this immediately providpotassium superoxide reaction commences.



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FIGUSmoke



RE 25-11 Hood



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er 1 April 2000, an unpressurisedng 5700 kg or having a maximum

mouth of each flight crew membern for a period of not less than 15uipment (PBE) may be provided byS 1.770(b)(1) or JAR-OPS 1.775more than one and a cabin crewrried to protect the eyes, nose and

provide breathing gas for a period

es, nose and mouth of all requiredas for a period of not less than 15

n the flight deck and be easily er at their assigned duty station.

each required cabin crew member



JAR Requirements

JAR-OPS 1.780 Crew Protective Breathing Equipment

An operator shall not operate a pressurised aeroplane or, aftaeroplane with a maximum certificated take-off mass exceediapproved seating configuration of more than 19 seats unless:

(a) It has equipment to protect the eyes, nose and while on flight deck duty and to provide oxygeminutes. The supply for Protective Breathing Eqthe supplemental oxygen required by JAR-OP(b)(1). In addition, when the flight crew is member is not carried, portable PBE must be camouth of one member of the flight crew and toof not less than 15 minutes; and

(b) It has sufficient portable PBE to protect the eycabin crew members and to provide breathing gminutes.

PBE intended for flight crew use must be conveniently located oaccessible for immediate use by each required flight crew memb

PBE intended for cabin crew use must be installed adjacent toduty station.



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mitter (ELT)

and located at or adjacent to thed (d) except that, where the fireust be stowed outside but adjacent

ired by JAR-OPS 1.685, JAR-OPS

ped with an automatic Emergencyanner that, in the event of a crash,

aximised and the possibility of the

on the distress frequencies

g temperature of -20ºC.



Emergency Locator Beacon/Trans

JAR Requirements

An additional, easily accessible portable PBE must be providedhand fire extinguishers required by JAR-OPS 1.790 (c) anextinguisher is located inside a cargo compartment, the PBE mto the entrance to that compartment.

PBE while in use must not prevent communication where requ1.690, JAR-OPS 1.810 and JAR-OPS 1.850.

JAR-OPS 1.820 Automatic Emergency Locator Transmitter

An operator shall not operate an aeroplane unless it is equipLocator Transmitter (ELT) attached to the aeroplane in such a mthe probability of the ELT transmitting a detectable signal is mELT transmitting at any other time is minimised.

An operator must ensure that the ELT is capable of transmitting

121. 5 MHz and 243 MHz for a period of 48 hrs at an operatin

JAR-OPS 1.780 Crew Protective Breathing Equipment



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as follows:

LT is intended to be permanentlyh and is designed to aid SAR teams

T is intended to be rigidly attachedovable from the aeroplane after a

sh sequence. If the ELT does notted antenna may be disconnectedse) attached to the ELT. The ELT

type of ELT is intended to aide SAR

of ELT is intended to be rigidlyautomatically ejected and deployedsh has occurred. This type of ELT teams in locating the crash site.

impact, the Automatic Emergencyucture as far aft as practicable withbability of the signal being radiated



IEM Requirements

IEM OPS 1.820 Automatic Emergency Locator Transmitter

Types of automatic Emergency Locator Transmitters are defined

(a) Automatic Fixed (ELT (AF)). This type of Eattached to the aeroplane before and after a crasin locating a crash site;

(b) Automatic Portable (ELT (AP)). This type of ELto the aeroplane before a crash, but readily remcrash. It functions as an ELT during the craemploy an integral antenna, the aircraft-mounand an auxiliary antenna (stored on the ELT cacan be tethered to a survivor or a lift-raft. This teams in locating the crash site or survivor(s);

(c) Automatic Deployable (ELT (AD)). This typeattached to the aeroplane before the crash and after the crash sensor has determined that a crashould float in water and is intended to aid SAR

To minimise the possibility of damage in the event of crash Locator Transmitter should be rigidly fixed to the aeroplane strits antenna and connections so arranged as to maximise the proafter a crash.



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senger seats and crew lift jackets inift jacket in general use, they are allder pressure in a small cylinder-cord. A stand-by mouth-operatedain inflation over extended periods. be deflated. Subsequent re-inflationoperations, life jackets are equipped battery is activated by contact witht dye marker and shark repellent Passenger life jackets are generally

aeroplane: When flying over watere shore; or

e take-off or approach path is so would be likelihood of a ditching,rvivor locator light, for each personn easily accessible from the seat orckets for infants may be substitutedvivor locator light.



Life Jackets68. Passenger lift jackets are normally stored beneath the paseasily accessible locations. Although there are several types of lbasically similar. Inflation is by means of CO2, stored unincorporated in the life jacket and activated manually by a pullinflation valve is provided for emergency inflation and to maintIf the CO2 inflation is inadvertently activated the life jacket cancan only be effected through mouth-operation. To aid in rescue with a battery-operated light and a mouth-operated whistle. Thewater. Some types of life jacket also incorporate fluorescencompounds, which stain the surrounding water when immersed.coloured yellow and crew life jackets are coloured dayglow.

JAR Requirements

JAR-OPS 1.825 Life Jackets

Land aeroplanes. An operator shall not operate a land and at a distance of more than 50 nautical miles from th

When taking off or landing at an aerodrome where thdisposed over water that in the event of a mishap thereunless it is equipped with life jackets equipped with a suon board. Each life jacket must be stowed in a positioberth of the person for whose use it is provided. Life jaby other approved flotation devices equipped with a sur



Chapte

ate a seaplane or an amphibian onth a survivor locator light, for each position easily accessible from thed. Life jackets for infants may bed with a survivor locator light.

ered to be flotation devices.

ater Flights

e at a distance away from land, that corresponding to:

l miles, whichever is the lesser, forto an aerodrome with the criticalpoint along the route or planned



IEM Requirements

Life Rafts

Jar Requirements

Seaplanes and amphibians. An operator shall not operwater unless it is equipped with life jackets equipped wiperson on board. Each life jacket must be stowed in aseat or berth of the person for whose use it is providesubstituted by other approved flotation devices equippe

IEM OPS 1.825 Life Jackets

For the purpose of JAR-OPS 1.825, seat cushions are not consid

JAR-OPS 1.830 Life-rafts and Survival ELTs for Extended Overw

On overwater flights, an operator shall not operate an aeroplanwhich is suitable for making an emergency landing, greater than

(a) 120 minutes at cruising speed or 400 nauticaaeroplanes capable of continuing the flight power unit(s) becoming inoperative at any diversions; or

JAR-OPS 1.825 Life Jackets



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miles, whichever is the lesser, for all

phs below is carried.

cess rafts of enough capacity arerated capacity of the rafts mustloss of one raft of the largest rated

staining life as appropriate to the (b) (2); and

ransmitters (ELT). (See IEM OPS

ater Flights

another;

;

ater Flights



AMC Requirements

(b) 30 minutes at cruising speed or 100 nautical other aeroplanes,

(c) unless the equipment specified in sub-paragra

Sufficient life-rafts to carry all persons on board. Unless exprovided, the buoyancy and seating capacity beyond the accommodate all occupants of the aeroplane in the event of a capacity. The life-rafts shall be equipped with:

(a) A survivor locator light; and

(b) Life saving equipment including means of suflight to be undertaken (see AMC OPS 1.830

(c) At least two survival Emergency Locator T1.820.)

AMC OPS 1.830 (b) (2) Life-rafts and ELT for Extended Overw

The following should be included in each life-raft:

(a) Means for maintaining buoyancy;

(b) A sea anchor;

(c) Life-lines, and means of attaching one life-raft to

(d) Paddles for life-rafts with a capacity of 6 or less

JAR-OPS 1.830 Life-rafts and Survival ELTs for Extended Overw



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tically switched ON when removedf the torch battery is indicated by a

long as the battery has an adequatettery power diminishes and when itcement. The red light also monitorsment fails. Once removed from itsprocedure, since the wall bracketn to remove the torch. A typical

ents;

distress signals described in ICAO

ife-raft is designed to carry:

ained in a pack.

ater Flights



Emergency Torches69. The emergency torches are self-powered, they are automafrom their stowage and OFF when refitted. The state of charge ored light on the torch body which flashes every 3 to 4 seconds asreserve of power. The rate of flashing becomes slower as the bareaches one flash every 10 seconds the battery is in need of replathe bulb filament, the red light will cease to flash when the filastowage, refitment of an emergency torch is a maintenance retaining band which holds the torch in place must be brokeemergency torch stowage is illustrated at Figure 25-12.

(e) Means of protecting the occupants from the elem

(f) A water resistant torch;

(g) Signalling equipment to make the pyrotechnicalAnnex 2;

(h) For each 4, or fraction of 4, persons which the l

(i) 100 g glucose tablets;

(j) First-aid equipment

(k) The above three items, inclusive, should be cont

AMC OPS 1.830 (b) (2) Life-rafts and ELT for Extended Overw



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FIGUEmerg

of the main electrical circuits on thecape path lighting will illuminatesential/vital/standby bus or from itstion for a minimum of 10 minutes.



RE 25-12ency Torch

Emergency Lighting70. Emergency lighting and exit signs must operate when all aircraft are rendered inoperative. Cabin emergency and esautomatically and will be supplied from the battery bus or an esown dedicated battery. The lighting must be capable of illumina



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hich has a maximum approved ided with an emergency lighting cuation of the aeroplane. The

ting configuration of more than 19:

eas; and

g signs.

te or equivalent was filed in a JAAlying by night, exterior emergency

eans are required.

te or equivalent was filed in a JAAflying by night, exterior emergency

a JAA Member State or elsewherecape path marking system in the

ting configuration of 19 or less and



JAR Requirements

JAR-OPS 1.815 Emergency Lighting

An operator shall not operate a passenger carrying aeroplane wpassenger seating configuration of more than 9 unless it is provsystem having an independent power supply to facilitate the evaemergency lighting system must include:

For aeroplanes which have a maximum approved passenger sea

(a) Sources of general cabin illumination;

(b) Internal lighting in floor level emergency exit ar

(c) Illuminated emergency exit marking and locatin

For aeroplanes for which the application for the type certificaMember State or elsewhere before 1 May 1972, and when flighting at all overwing exits, and at exits where descent assist m

For aeroplanes for which the application for the type certificaMember State or elsewhere on or after 1 May 1972, and when lighting at all passenger emergency exits.

For aeroplanes for which the type certificate was first issued inon or after 1 January 1958, floor proximity emergency espassenger compartment(s).

For aeroplanes which have a maximum approved passenger seaare certificated to JAR –23 or JAR-25:



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g signs.

ting configuration of 19 or less andin illumination.

assenger carrying aeroplane which or less unless it is provided with a of the aeroplane. The system mayd on the aeroplane and which are been switched off.

mum approved passenger seatingsengers unless it is equipped withuse by crew members during an



Megaphone

JAR Requirements

(a) Sources of general cabin illumination;

(b) Internal lighting in emergency exit areas; and

(c) Illuminated emergency exit marking and locatin

For aeroplanes which have a maximum approved passenger seaare not certificated to JAR-23 or JAR-25, sources of general cab

After 1 April 1998 an operator shall not, by night, operate a phas a maximum approved passenger seating configuration of 9source of general cabin illumination to facilitate the evacuationuse dome lights or other sources of illumination already fittecapable of remaining operative after the aeroplane’s battery has

JAR-OPS 1.810 Megaphones

An operator shall not operate an aeroplane with a maxiconfiguration of more than 60 and carrying one or more pasportable battery-powered megaphones readily accessible for emergency evacuation, to the following scales:

JAR-OPS 1.815 Emergency Lighting



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s when the total passenger seating.

sible from a cabin crew member’sey should be suitably distributed iners assigned to direct emergencye positioned such that they can beer’s seat.



AMC Requirements

(a) For each passenger deck:

For aeroplanes with more than one passenger deck, in all caseconfiguration is more than 60, at least 1 megaphone is required

AMC OPS 1.810 Megaphones

Where one megaphone is required, it should be readily accesassigned seat. Where two or more megaphones are required, ththe passenger cabin(s) and readily accessible to crew membevacuations. This does not necessarily require megaphones to breached by a crew member when strapped in a cabin crew memb

JAR-OPS 1.810 Megaphones



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f gloves. Current designs are madele fire extinguishers.

nventoryency equipment typically fitted in a 25 persons and the life rafts aboutcent to cabin attendants’ door seats.towed in the cockpit), torches and

rtificated take-off mass exceeding iguration of more than 9 seats ted on the flight deck. If the n 200 an additional crash axe or rd galley area.

must not be visible to passengers.



Crash Axe

JAR Requirements

Fireproof Gloves71. Most commercial aircraft carry at least one set of fireprooin Kevlar. Such gloves are normally located alongside the portab

Emergency Equipment – Typical I72. The table at Figure 25-13 shown the quantity of emergBoeing 757 aircraft. The slide rafts will typically support about40 persons. Emergency locator beacons are usually stowed adjaAdditional cabin emergency equipment consists of jemmies (smegaphones for use by cabin staff.

JAR-OPS 1.795 Crash Axes and Crowbars

An operator shall not operate an aeroplane with a maximum ce5700 kg or having a maximum approved passenger seating confunless it is equipped with at least one crash axe or crowbar locamaximum approved passenger seating configuration is more thacrowbar must be carried and located in or near the most rearwa

Crash axes and crowbars located in the passenger compartment



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FIGUEmergEquipm



RE 25-13ency ent B757

Cabin Emergency Equipment

8 Doors

6 Inflatable Slides

2 Inflatable Slides - Dual Lane

8 BCF Fire Extinguishers

8 Kelvar Gloves

14 Portable Oxygen Sets

8 Smoke Hoods

228 Passenger Life Jackets

8 Crew Life Jackets

23 Infant Life Jackets

23 Child Seat - Belts

8 Emergency Torches

3 First Aid Kits

3 Megaphones

4 Infant Flotation Cots

4 Life Rafts (EROPS Aircraft)

2 Sabre Emergency Location Beacons (EROPS)



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Flight Deck Emergency Equipment

First-Aid Kits

4 Slide Rafts

4 Survival Packs

2 BCF Fire Extinguishers

4 Smoke Goggles

1 Crew Oxygen Set with Full Face Mask

1 Jemmy

2 Windows with Escape Straps

4 Crew Life Jackets

4 Emergency Torches

1 Kelvar Gloves

The following should be included in the First-Aid Kits:

Bandages (unspecified)

Burns dressings (unspecified)

Wound dressings, large and small



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). This should include information

first-aid kit should, where possible,



Adhesive tape, safety pins and scissors

Small adhesive dressings

Antiseptic wound cleaner

Adhesive wound closures

Adhesive tape

Disposable resuscitation aid

Simple analgesic e.g. paracetamol

Antiemetic e.g. cinnarizine

Nasal decongestant

First-Aid handbook

Splints, suitable for upper and lower limbs

Gastrointestinal Antacid +

Anti-diarrhoeal medication e.g. Loperamide +

Ground/Air visual signal code for use by survivors

Disposable Gloves

A list of contents in at least 2 languages (English and one otheron the effects and side effects of drugs carried.

NOTE: An eye irrigator whilst not required to be carried in the be available for use on the ground.

+ For aeroplanes with more than 9 passenger seats installed.



Chapte

arried in the aeroplane:



Emergency Medical-Kit

The following should be included in the emergency medical kit c

Sphygmomanometer - non mercury

Stethoscope

Syringes and needles

Oropharyngeal airways (2 sizes)

Tourniquet

Coronary vasodilator e.g. nitro-glycerine

Anti-smasmodic e.g. hyascene

Epinephrine 1 : 1000

Adrenocortical steroid e.g. hydrocortisone

Major analgesic e.g. nalbuphine

Diuretic e.g. fursemide

Antihistamine e.g. diphenhydramine hydrochloride

Sedative/anticonvulsant e.g. diazepam

Medication for Hypoglycaemia e.g. hypertonic glucose

Antiemetic e.g. metoclopramide

Atropine

Digoxin



Chapte

equipped with first-aid kits, readily

xtent possible, that contents are intended use; and

ce with instructions contained on



Uterine contractant e.g. Ergometrine/Oxytocin

Disposable Gloves

Bronchial Dilator - including an injectable form

First-Aid Kits

(a) An operator shall not operate an aeroplane unless it is accessible for use, to the following scale:

(b) An operator shall ensure that first-aid kits are:

(i) Inspected periodically to confirm, to the emaintained in the condition necessary for their

(ii) Replenished at regular intervals, in accordantheir labels, or as circumstances warrant.



Chapte

imum approved passenger seatingd with an emergency medical kit iftes flying time (at normal cruisingassistance could be expected to be



Emergency Medical-Kit

(a) An operator shall not operate an aeroplane with a maxconfiguration of more than 30 seats unless it is equippeany point on the planned route is more than 60 minuspeed) from an aerodrome at which qualified medical available.



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