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Theory of Flightengineering uk

Contents

1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES ANDSYSTEMS) .......................................................................................... 1-2

1.1 A EROPLANE AERODYNAMICS AND FLIGHT CONTROLS ................. .... 1-21.1.1 Fixed Aerofoils ......... ......... ......... .......... ......... .......... ...... 1-21.1.2 Moveable Control Surfaces .......... ......... .......... ......... ..... 1-61.1.3 High Lift Devices ......... ......... ......... .......... ......... .......... ... 1-131.1.4 Drag Inducing Devices ......... ......... .......... ......... ......... .... 1-141.1.5 Airflow Control Devices Wing Fences ......... ......... ....... 1-171.1.6 Boundary Layer Control ......... ......... ......... ......... .......... .. 1-181.1.7 Trim Tabs ......... ......... .......... ......... ......... .......... ......... ..... 1-211.1.8 Mass Balance .......... ......... ......... .......... ......... .......... ...... 1-241.1.9 Control Surface Bias ......... ......... .......... ......... ......... ....... 1-261.1.10 Aerodynamic Balance Horn Balance ......... ......... ........ 1-261.1.11 Aerodynamic Balance Inset Hinge........... ......... .......... 1-27

1.2 H IGH S PEED FLIGHT ..................................................................... 1-281.2.1 Speed of Sound .......... ......... ......... .......... ......... .......... ... 1-281.2.2 Subsonic Flight ......... .......... ......... ......... .......... ......... ..... 1-291.2.3 Transonic Flight .......... ......... ......... .......... ......... .......... ... 1-301.2.4 Supersonic Flight ......... .......... ......... ......... .......... ......... .. 1-32

1.2.5 Aerodynamic Heating ......... .......... ......... ......... .......... ..... 1-391.2.6 Area Rule ......... ......... .......... ......... ......... .......... ......... ..... 1-401.2.7 Factors Affecting Airflow in Engine Intakes of High Speed Aircraft 1-411.2.8 Effects of Sweepback on Critical Mach Number .......... .. 1-43

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1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURESAND SYSTEMS)

The principles of Aircraft Theory of Flight are covered in JAR 66 Module 8.

1.1 AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS

An aircraft is equipped with fixed and moveable surfaces, or aerofoils, whichprovide stability and control. Each item is designed for a specific function duringthe operation of the aircraft.

Typical Aircraft Flight ControlsFigure 1

1.1.1 FIXED AEROFOILS

The fixed aerofoils are the wings or mainplanes, the horizontal stabiliser ortailplane and vertical stabiliser or fin. The function of the wings is to provideenough lift to support the complete aircraft. The tail section of a conventionalaircraft, including the stabilisers, elevators and rudder, is occasionally known as

the empennage.

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1.1.1.1 Horizontal StabiliserThe horizontal stabiliser is used to provide longitudinal pitch stability and isusually attached to the aft portion of the fuselage. It may be mounted either ontop of the vertical stabiliser, at some mid-point, or below it.

Conventional horizontal stabilisers are placed aft of the wing and normally set ata slightly smaller or negative angle of incidence with respect to the wing chordline.

This configuration gives a small downward force on the tail with a valuedependent on the size of the stabiliser and its distance from the Centre of Gravity(CG).

Horizontal StabiliserFigure 2

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1.1.1.2 T-Tail ArrangementThe T-Tail Arrangement places the complete stabiliser/tailplane and elevatorassembly on top of the vertical stabiliser. This ensures that pitch control is notaffected by turbulent air from the wing. It also makes the vertical stabiliser andrudder control more effective, due to the so- called end plate effect.

However a T-Tail (and rear engine) configuration, would be dangerous if theaircraft entered what is termed a deep st all. At a very high angle of attack (i.e.:stalling angle), airflow could make pitch control non-effective (and may cause theengines to flame out). To prevent this, T- Tailed aircraft will have a stick pushsystem, in order to automatically recover them safely from excessive angles ofattack.The T-Tail has another disadvantage in that the empennage structure will beheavier than normal, due to the strengthening required to combat greater bendingloads. However since the pitch moment arm is increased, the stabiliser andelevators can be made smaller and therefore lighter than conventional designs.

Often, the complete stabiliser can be moved to provide longitudinal trim, negatingthe use of trim tabs (later in Module 11.09).

T Tail ArrangementFigure 3

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1.1.1.3 Vertical StabiliserThe vertical stabiliser for an aircraft is the aerofoils forward of the rudder and isused to provide directional stability.

A problem encountered on single-engined propeller driven aircraft is that thepropeller causes the airflow to rotate as it travels rearward. This strikes one sideof the vertical stabiliser more than the other, resulting in a yawing moment. Theseaircraft may have the leading edge of the stabiliser offset slightly, thereby causingthe airflow to pass around it in such a manner to counter the yaw.

Off-Set Vertical StabiliserFigure 4

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1.1.2 MOVEABLE CONTROL SURFACESMoveable control surfaces are normally divided into Primary and Secondarycontrols.

The primary control surfaces include the elevators, rudder, ailerons and rollspoilers. The secondary control surfaces consist of trim controls (tabs), high liftdevices (flaps and slats), speed brakes and lift dumpers (additional spoilers).

Note: Traditionally, spoilers have not been included as primary controls, but thosewhich operate in conjunction with the ailerons during roll, are considered to beprimary in the JAR 66 syllabus, so this is how these notes will define them.

The primary control surfaces are used to make the aircraft follow the correct flightpath and to execute certain manoeuvres.

The secondary controls are used to change the lift and drag characteristics of theaircraft or to provide assistance to the primary controls.

Moveable Control SurfacesFigure 5

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1.1.2.1 Roll Control - AileronsThese primary controls provide lateral (roll) control of the aircraft, that is,movement about the longitudinal axis. They are normally attached to hinges atthe trailing edge of the wing, near the wing tip. They move in opposite directions,so that the up-going aileron reduces lift on that side, causing the wing to go down,whilst the down-going surface increases the lift on the opposite side, raising thewing.

Large aircraft often use two sets of aileron surfaces on each wing, one in theconventional position near the wing tip and the other set at mid-span or outboardof the flaps. The inboard set is referred to as high speed ailerons. The outboardsurfaces, or sometimes both sets, work at low speeds to give maximum controlduring take off and landing, for example when large movements may be required.

At high cruising speed the outer ailerons are isolated and only the inboard setoperate. If the outer ailerons were permitted to operate at high speed, the stressproduced at the wing tips may twist the wing and produce aileron reversal. Thisis particularly likely with modern highly flexible thin wings, where the possibility ofstructural damage may result if the outboard surfaces were too powerful.

The ailerons are operated by a control wheel, a control column or a side-stick.Movement of any of these inputs away from neutral towards one side, will resultin the aircraft rolling to that side. Returning the control to neutral at this stage willleave the aircraft in a banked condition and a similar but opposite movement willbe required to bring the aircraft level once more.

Aileron ControlsFigure 6

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The ailerons are usually operated in conjunction with the rudder and/or elevatorduring a turn and are rarely used on their own. A co-ordinated turn is one thatoccurs without slip or skid. Too little bank will cause the aircraft to skid outwards,too much bank will cause the aircraft to slip downwards.

1.1.2.2 Roll Control - Spoilers

The use of spoilers as a primary control, will be to operate asymmetrically inconjunction with aileron movement and are normally referred to as Roll Spoilers.

Roll spoilers are mounted on the top of the wing just inboard of the outboard set

of ailerons.

Roll Spoiler ControlsFigure 7

Movement of the aileron control wheel on the flight deck will deploy each spoilerprogressively upwards with the up-going aileron, whilst on the side of the down-going aileron, the spoiler will remain flush with the upper wing camber.

.This is achieved by the control system being routed via a spoiler/aileron mixerunit. The up-going spoiler will effectively spoil the lift on the down-going wing andaugment the similar effect of the up-going aileron.

Alternatively, on some aircraft the spoilers will replace the ailerons completely toprovide the sole means of roll control.

Note: Other spoiler functions are covered later under Secondary Controls.

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1.1.2.3 Pitch Control - ElevatorsThe elevators are the control surfaces which govern the movement of the aircraftin pitch about its lateral axis. They are normally attached to the hinges on the rearspar of the horizontal stabiliser.

When the control column of the aircraft is pushed forward, the elevators movedown.. The resultant force of the airflow generated lift', acting upwards, raisesthe tail and lowers the nose of the aircraft. The reverse action takes place whenthe control is pulled back.

1.1.2.4 Pitch Control Stabilators

A special type of pitch control surface that combines the functions of the elevatorand the horizontal stabiliser is the stabilator, often referred to as a slab or all-flying tailplane . The stabilator is a complete all-moving horizontal stabiliser whichcan change its angle of attack when the control column is moved and therebyalter the total amount of lift generated by the tail.

Elevator ControlsFigure 8

Stabilator ControlsFigure 9

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1.1.2.5 Pitch Control Variable Incidence StabilisersIncorporating a conventional elevator control system, the variable incidencehorizontal stabiliser is often used for pitch trim. Normally a powerful electric motoris used to vary its angle of attack when trim switches on the flight deck areoperated.

Variable Incidence stabiliserFigure 10

1.1.2.6 Canards

Some earliest powered aircraft, such as the Wright Flyer, had horizontal surfaceslocated ahead of the wings. This configuration, with the forward surface usuallyreferred to as a canard or foreplane, has been used on occasions, up to thepresent day.

Conventional aircraft have the tailplane located at the rear of the fuselage whichprovides a small, stabilising down force. This means that the wing has to produceslightly more lift to balance this down force. As we have seen, in order for a wingto produce lift it must also generate drag.

With the tailplane located at the front of the aircraft, the stabilising force isdirected upwards. This contributes to the total lift of the aircraft, thereby reducingdrag from the lift producing wing.

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A fundamental feature of a canard design is that the angle of attack of theforeplane, (in front of the CG of the aircraft) is set at a greater angle than themain wing. This feature will ensure that the foreplane reaches the stalling anglefirst, resulting in a predictable dropping of the nose and a certain recovery.

Additionally, stall sensing systems (later), can be triggered just before theforeplane reaches its critical angle of attack, leaving the main wing safely belowthe stalling angle and still producing adequate lift.

Canard Design Beech StarshipFigure 11

1.1.2.7 Yaw Control - Rudder

The rudder is a vertical control surface that is hinged at the rear of the fin and isdesigned to apply yawing moments. The rudder rotates the aircraft about itsvertical axis and is controlled by rudder pedals that are operated by the pilotsfeet. Pushing on one pedal, the right for example, causes the rudder to move tothe right also. This causes the rudder to generate a 'lifting' force sideways to theleft which turns the nose of the aircraft to the right.

Because of the power of some rudder systems, particularly assisted systems,they may have their range reduced at high speed by means of a speed-sensitiverange limiting system.(later).

The rudder is normally a single structural unit but on large transport aircraft it may

comprise two or more operational segments, moved by different operatingsystems to provide a level of redundancy.

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Rudder controlsFigure 12

1.1.2.8 Combined-Function Controls Elevons and Ruddervators

An example of combined-function controls is found on delta-wing aircraft, wherecontrol surfaces for pitch and roll must be fitted on the trailing edge of the wing.

Controls with a dual-function (elevators and ailerons) called elevons, provideboth pitch and roll, by moving symmetrically in pitch or asymmetrically in roll via amixer unit, when the control column or control wheel are operated on the flightdeck..

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Another example are ruddervators normally used on aircraft fitted with a 'V' orButterfly tail. These surfaces serve the purposes of both rudder and elevator.

Ruddervator ControlsFigure 13

1.1.3 HIGH LIFT DEVICES

Aerodynamic lift is determined by the shape and size of the main lifting surfacesof the aircraft. In order to produce the outstanding performance achieved by alarge modern, swept wing, passenger jet such as the Boeing 777, the wing isdesigned to give optimum lift to support the aircraft whilst in cruise (typicallyMach 0.87).

This has meant, that to be able to control and land the aircraft weighing around200-tonne on runways of reasonable length, the landing speed needs to beslower than the clean stalling speed of the aircraft. In order to achieve this, morelift is required and this is obtained from so-called high lift devices.

These are divided generally into leading edge devices, namely slots, slats andKrueger flaps and trailing edge devices including plain, slotted and fowler flaps.They will increase lift and as a result, reduce the stalling speed. Consequently thelanding speed, (about 1.3 times the stalling speed), will also be reduced, sincedrag is also increased with large angles of trailing edge flap deployment.

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Flaps and SlatsFigure 14

Additionally, some aircraft incorporate ailerons, both of which are designed tomove downwards together whenever the trailing edge flaps are extended to thelanding position. These will act as additional plain flaps and provide extra drag(and lift), but will still provide roll control if required.

These surfaces are referred to as Droop Ailerons or Flaperons.

Droop AileronFigure 15

1.1.4 DRAG INDUCING DEVICES

There are several situations where the aircraft must slow down fairly quickly. Withslower, high drag, light aircraft, simply closing the throttle allows the high drag ofthe airframe and the idling propeller to slow the aircraft down, to gliding speedprior to landing approach, for example.

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As previously stated, a modern airliner is an extremely smooth, low drag design

which, if only the throttles are retarded, will continue in level flight for many milesbefore slowing down. Furthermore, if the nose were lowered more than a degreeor so, the aircraft will begin to accelerate again.

In order to overcome the problems of low drag on large aircraft with highmomentum, the designers have introduced a variety of drag inducing devices.These include spoilers, lift dumpers, speed brakes and in unusual circumstances,lowering the landing gear and operating in-flight thrust reversers.

1.1.4.1 Spoilers and Lift Dumpers.

Spoilers and Lift Dumpers are usually hinged panels located about mid-chordposition on the upper surface of the wing. Hydraulically operated, they produce alarge amount of turbulence and drag when deployed, resulting in a reduction oflift.

Lift Dump Spoilers

Figure 16Spoilers, have a variety of uses, all of which involve spoiling the lift of the wing.Some of the following facilities can be combined, so that one set of panels canhave more than one job.

Firstly, they can be the primary roll control of the aircraft as described previously.

Secondly, the spoilers can be used in a symmetrical, part-deployed position,allowing the aircraft to slow down quickly in the cruise, or descend at a muchsteeper rate without accelerating. On some aircraft, the deployment angle of thespoiler panels can be varied by changing the position of the control lever in theflight compartment.

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Lift dumpers are, as their name describes, are spoiler panels incorporated solely

to dump lift. They are normally deployed after landing, destroying the lift of thewing and producing high drag, to assist in stopping the aircraft efficiently andthereby allowing the wheel brakes to be operated more effectively.

1.1.4.2 Speed Brakes

Whilst it is true that the in-flight use of spoilers may be referred to as selecting the'speed brakes', the term more accurately describes devices which are solely forthe production of drag without any change of trim. The rear fuselage mounted'clamshell- type doors which open up on the BAe 146 and Fokker 70/100 aircraftare true speed brakes (or air brakes) and have the following major advantageover the use of spoilers for producing drag.

When the wing mounted spoilers are deployed, vibration or rumble is often felt inthe passenger cabin, which some people may find disturbing. The aft mountedspeed brakes not only produce high drag at any airspeed, but their selection isvirtually vibration free. Also, lift will be completely unaffected, thus permitting theirdeployment on approach and making a go-around much safer. (This will becovered later in powerplants).

Speed Brake InstallationFigure 17

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1.1.5 AIRFLOW CONTROL DEVICES WING FENCESThese devices are usually fitted to aircraft with swept wings. Total airflow over aswept wing, splits into two components, one moving across the wing chordparallel to the airflow and the other flowing spanwise towards the wing tip.

The fences are fitted about mid-span, on the leading edge of the wing andextending rearwards. They are designed to control the spanwise flow of theboundary layer air over the top of the wing. Also they will straighten the airflowover the ailerons, improving their effectiveness and straighten the air nearer thewing tip, resulting in less 'spillage' of air from beneath the wing to the top, therebyproducing less drag. (See Winglets later).

Wing FencesFigure 18

1.1.5.1 Airflow Control Devices Saw Tooth Leading Edges

This form of airflow control is more common on military aircraft than moderncommercial airliners. The saw tooth or notch is simply a small increase in wingchord on the outer portion of the wing. The step where the change occurs, tendsto form an invisible 'wall' of high velocity air, which flows over the wing andstraightens the spanwise flow. It functions in much the same way as the wingfence but removes the extra drag and weight penalty.

Leading Edge NotchFigure 19

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1.1.5.2 Airflow Control - WingletsThese can be seen on a variety of the later generation airliners and business jets.The outboard part of the wing are upswept to an extreme dihedral angle. Thesewinglets work best at higher speeds and, by clever aerodynamic design, will givebetter airflow control and reduce the drag produced by the wing. It does this byusing the up-flow from below the wing to produce a forward thrust from thewinglet, rather like a yacht sail. The winglets add weight to the aircraft as well asincreasing parasitic drag, but the large reduction in induced drag at the wingtip,results in a significant fuel saving.

WingletsFigure 20

1.1.6 BOUNDARY LAYER CONTROL

The boundary layer is that layer of air adjacent to the aerofoil surface (theboundary between metal and air). If measured, the air velocity in th e layer willvary from zero directly on the surface, to the relevant velocity of the free streamat the outer extremity of the boundary layer.

Normally, at the leading edge of the wing the boundary layer will be laminar, (insmooth thin sheets close to the surface), but as the air moves over the wingtowards the trailing edge, the boundary layer becomes thicker and turbulent. Theregion where the flow changes from laminar to turbulent is called the transition

point. .As airspeed increases, the transition point tends to move forward, so thedesigner tries to prevent this thus maintaining laminar flow, over the top of thewing for as far back as possible. Methods of boundary layer control are asfollows:

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1.1.6.1 Boundary Layer Control - Vortex GeneratorsOne way of stimulating the boundary layer and stopping the airflow becomingincreasingly sluggish towards the trailing edge is the use of vortex generators.

Vortex generators are small plates or wedges projecting up from the surface of anaerofoil about 25mm.(about 3 times the typical boundary layer thickness), into thefree stream air. Their purpose is to shed small but lively vortices from their tip,which act as scavengers to direct and mix the high energy free stream air into thesluggish boundary layer air and invigorate it. This action pushes the transitionpoint backwards towards the trailing edge .

In this way,the small amount of drag created by the vortices is far more thancompensated by the considerable boundary layer drag which they save. Theyalso weaken the shock wave at high speed and reduce shock drag also. (later).

Vortex GeneratorsFigure 21

1.1.6.2 Boundary Layer Control - Stall Wedges

We have seen previously that washout on a wing permits the root of the wing tostall first, allowing the pilot to retain roll control during the stall. Even with adegree of washout, the aircraft will drop a wing on occasions due to adverseboundary layer air causing the outer part of the wing to stall first. This can beovercome with the use of stall wedges, or stall strips, as they are sometimesknown.

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Stall Wedges are small, wedge-shaped strips mounted on the leading edge of thewings at about one third span. The are designed to disrupt the boundary layerairflow, at large angles of attack approaching the stall, thus ensuring the airflowbreaks away,(stalls), at the root end of the wing first.

Additionally they produce a similar effect to a wing fence at smaller angles ofattack resulting in a smoother airflow over the ailerons, thus retaining optimumroll control.

Stall WedgesFigure 22

1.1.6.3 Boundary Layer Control - Leading edge Devices

Other devices to prevent laminar separation at the low speed end of the rangeand thus control boundary layer air are leading edge droop flaps and Kreugerflaps. They can be a droop snoot or permanent droop type, or can be adjustedduring flight.

Krueger (left) and Drooped (right) Leading Edge FlapsFigure 23

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1.1.7 TRIM TABSDuring a flight an aircraft will develop a tendency to deviate from a straight andlevel hands -off attitude. This may be due to changes in fuel state, speed, loadposition or flap/landing gear selection and could be countered by applying acontinuous correcting force to the primary controls. This would be fatiguing for thecrew and difficult to maintain for long periods, so trim tabs are used for thispurpose instead.

Trim tabs move the primary control surface aerodynamically in the oppositedirection to the movement of the tab. To correct an aircraft nose down out of trimcondition, the elevator tab is moved down, resulting in the elevator moving up, thetail of the aircraft moving down , so that the nose comes up , correcting the fault.

1.1.7.1 Fixed Trim Tabs

A fixed trim tab may be a simple section of sheet metal attached to the trailingedge of a control surface. It is adjusted on the ground by simply bending it up ordown, to a position resulting in zero control forces during cruise. Alternatively, thetab is connected to the primary control by a ground-adjustable connecting rod.Finding the correct position for both types is by trial and error.

Fixed Trim TabFigure 24

1.1.7.2 Controllable Trim Tabs

A controllable trim tab is adjusted from the flight deck, with its position beingtransmitted back to a flight deck indicator showing trim units, left and right ofneutral.

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Flight deck controls are trim-wheel, lever, switch, etc., with the actuation of thetab by mechanical, electrical or hydraulic means. Trim facilities are normallyprovided on all three axes.

Controllable Trim TabFigure 25

Note: Aircraft with hydraulic fully powered controls do not have trim tabs. Sincefully powered controls are termed irreversible, trim tabs if fitted, would beaerodynamically ineffective. With these systems, trimming is achieved by movingthe primary control surface to a new neutral datum.(later).

1.1.7.3 Servo TabsSometimes referred to as the flight tabs, servo tabs are positioned on the trailingedge of the primary control surface and connected directly to the flight deckcontrol inputs. They act as a form of power booster, since pilot effort is onlyrequired to deflect the relatively small area of the servo tab into the air stream.

Movement of the flight deck control input moves the tab up or down and theaerodynamic force created on the tab, moves the primary control, until theaerodynamic load on the control surface balances that on the tab. Moving the tabdown will cause the primary control to move up and vice-versa.

Servo TabFigure 26

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1.1.7.4 Balance TabsBalance tabs assist the pilot in moving the primary control surface. The flight deckcontrols are connected to the primary control surface whereas the balance tab,hinged to the trailing edge of the primary surface, is connected to the fixedaerofoil. For example, the elevator balance tab, will be connected by anadjustable rod to the horizontal stabiliser and is so arranged, that it tends tomaintain the tab at the same relative angle to the stabiliser when the pilot movesthe elevator.

Aerodynamically, therefore, the tab is moving in the opposite direction to thecontrol surface and assists its movement. Adjusting the length of the connectingrod will alter the displacement of the effective range of the tab about the mid-pointdatum.

Some types of balance tab have more than one point of attachment and it ispossible with these so called geared balance tabs, to alter the range of tabdeflection.

The function of a balance tab can also be combined with that of a trim tab, byadjusting the length of the balance tab connecting rod from the flight deck. This isusually achieved by installing a form of linear actuator in the rod and is termed atrim/balance tab (Geared balance and trim/balance tabs will be covered later inthe notes).

Balance TabFigure 27

1.1.7.5 Anti-Balance Tabs

Anti-balance tabs operate in a similar way aerodynamically as balance tabs but

with a reverse effect. The difference is in the way it is connected to the fixedaerofoil. It is routed so that the tab moves, relative to and in the same directionas, the primary control surface. The effect is to add a loading to the pilot effort,mak ing it slightly heavier and thus providing feel, to prevent the possibility ofover-stressing the airframe structure.

Anti-Balance Tab

Figure 28

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1.1.7.6 Spring Tabs

At high speed, control surfaces operated directly from the flight deck, becomeincreasingly difficult to deflect from neutral, due to the force of the aerodynamicloads caused by the airstream around them.

The spring tab is progressive in its operation and provides increasingaerodynamic assistance in moving the control surface, with an increase in aircraftforward speed. The flight deck controls are connected to the spring tab in asimilar manner to the servo tab previously described, except the linkage is routedvia a torque rod assembly (or spring box) attached to the primary control surface.

When the aircraft is stationary or flying at low airspeed the airloads are non-

existent or very small. If the flight deck controls are deflected from neutral, therigidity of the torque tube (or spring force) causes the primary control to bedeflected together with the spring tab. The tab will remain in the same relativeposition with the primary control and consequently provides no additionalaerodynamic assistance.

As the aircraft flies faster, the increased force produced by the airflow, opposesthe movement of the primary control surface from its neutral position. Deflectionof the

flight deck controls in this case causes the torque tube to twist (or the spring tocompress), resulting in a deflection of the spring tab.

The tab deflection provides an added aerodynamic load which assists the flightdeck effort. The faster the aircraft flies, the greater the airflow force and thereforethe greater the spring tab deflection, resulting in a progressively increasingassistance in moving the primary control.

Spring TabFigure 29

1.1.8 MASS BALANCE

All aircraft structures are distorted when loads are applied. If the structure iselastic, as all good structures are, it will tend to spring back when the load isremoved, or its point of application is changed.

Since a control surface is hinged near its leading edge, the centre of gravity (C of

G) will be behind the hinge and as a consequence, there will be more weight aftof the hinge line than in front of it .

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In the case of an aileron for example, should the air load distort the wing

upwards, it is likely that the aileron will lag behind and distort downwards. Thiseffectively produces an extra upward aerodynamic force which pushes the wingup even further.

Due to its elasticity, the wing will spring back and the aileron will lag again but thistime upwards, aerodynamically forcing the wing down further than it wouldnormally go due to elastic recoil alone. Now the cycle is repeated and a highspeed oscillation will result. This unwanted phenomenon is referred to as flutter.

Flutter can be prevented if the C of G of the control surface is moved in line with,or slightly in front of, the hinge line. The normal way of achieving this is to add anumber of high density weights, either within the leading edge of the surface itselfor externally, ahead of the hinge line. The addition of these weights, normallymade from lead or depleted uranium, is closely controlled and calculated toensure that the exact balance is obtained.

This procedure of adding weights is referred to as mass balancing of the controls.

External Mass WeightsFigure 30

Integral Mass WeightsFigure 31

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1.1.9 CONTROL SURFACE BIASWhen a control surface is set so it is not in the true neutral position it is referred toas having a bias. There are many reasons for not having the controls in a truecentral position, including compensating for design features. As an example, asingle propeller aircraft may have a tendency to roll in the opposite direction tothe engines torque, to counteract this moment the ailerons could be offset withone slightly up and the other down. Once the aircraft is flying level with the biasset the trim gauge in the cabin would then be set to read zero.

1.1.10 AERODYNAMIC BALANCE HORN BALANCE

In order to overcome the high stick forces on larger aircraft at higher speeds, thesurfaces themselves are used to lighten the forces.

This is referred to as Aerodynamic Balancing and the three principal ways ofachieving it are: horn balance, inset hinge and pressure balancing.

This method, a small part of the primary control surface ahead of the hinge willproject into the airflow when the control is deflected from neutral. The airflow onthis side assists the movement of the control in the desired direction and willattempt to move the control further away from the neutral position.

Air loads on the control side, aft of the hinge, try to push the surface back towards

neutral. (This is the force that would normally make the controls heavy).If the proportion of balance area forward of the hinge and control area aft of thehinge is correct, the pilot will feel that his control loads are more manageable,making the aircraft easier to fly.

Horn BalanceFigure 32

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1.1.11 AERODYNAMIC BALANCE INSET HINGE

This method is similar to and has the same effect as the horn balance. Instead ofhaving a forward projection at one or both ends of the control surface, the hingesare set back so that the area forward of the hinge line, which projects into the airflow when the control surface is moved from neutral, is spread evenly along itswhole length.

Inset Hinge BalanceFigure 33

1.1.11.1 Aerodynamic Balance Balance Panels

A device fitted to a few aircraft is the aerodynamic balance panel. Often used inthe aileron system, the panel is fitted between the leading edge of the aileron,ahead of the hinge and the rear face of the wing. When the aileron is deflectedupwards (downwards) from neutral, the high velocity, low pressure air passingover the lower (upper) gap decreases the air pressure under (above) the balancepanel and pulls it down (up). The force on the balance panel is proportional toairspeed and control surface deflection and assists the pilot in moving thecontrols accordingly.

Aerodynamic Balance PanelFigure 34

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1.2 HIGH SPEED FLIGHT Advancement in modern aircraft and engine design has produced very largeairliners capable of cruising at 87% of the speed of sound. Typically at an altitudeof 11,000 metres (approximately 36,000feet), this will amount to an airspeed ofabout 575 miles per hour.

Earlier in the course the effects of subsonic air were considered. As airspeedincreases, the aerodynamic effects of airflow passing over an aircraft, go througha series of changes, which will now be considered.

1.2.1 SPEED OF SOUND

One of the most important measurements in high speed aerodynamics is basedon the speed of sound and so called mach number.

Mach number is named after the Austrian physicist Ernst Mach (1838-1916) andis the ratio of true airspeed of an aircraft to the local speed of sound at thataltitude. (This will be covered in more detail later).

Sound waves, like those produced by a stationary object vibrating at certainfrequencies, will cause a continuous series of pulses or pressure waves, toradiate outwards equally in all directions from the point of origin and travel inexactly the same manner as the ripples on a pond.

Pressure Waves Stationary ObjectFigure 35

The actual speed at which the waves radiate, depends on the type and density ofthe material in which they are travelling. Air and Water are both fluids but water ismore dense than air, so sound waves will travel faster (about 4 times) in waterthan in air.

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Additionally, in any one of the fluids, speed will vary with a change in

temperature. As temperature increases, the speed of sound will increase andvice-versa, so that in Air on a standard day at sea level (15 oC approx), the waveswill travel at 761mph (661.7 knots), whereas at 11,000 metres altitude, the speedwill fall to 661mph, since the temperature has dropped to -56 oC at this altitude.

Note: At altitudes above 11,000 metres and up to about 27,000 metres, thetemperature and hence the speed of sound, will remain constant.

1.2.2 SUBSONIC FLIGHT

The propagation of the pressure waves from a stationary object has beendiscussed above.

When an aircraft begins to move through the air at subsonic speeds, (a speedless than pressure wave propagation speed) the waves still travel forward and itis as if a message is sent ahead of the aircraft to warn of its approach.

On receipt of this message, the air streams begin to divide to make way for theaircraft but there is very little, if any change in the density of the air as it flowsover the aircraft. This warning message can be detected perhaps 100metres infront of the aircraft.

Consequently, anyone standing ahead of the aircraft, would hear it coming andbe able to detect the change in the nature of the pressure waves as the aircraftpassed by. It would be similar to the change in the pitch of the siren of a passingemergency road vehicle.

This is often referred to as Doppler shift or Doppler effect.

Pressure waves Subsonic FlightFigure 36

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1.2.3 TRANSONIC FLIGHT At subsonic speeds, the study of aerodynamics is simplified by the fact that airpassing over a wing experiences only very small changes in pressure anddensity. The airflow is termed incompressible as, when it passes through aventuri, the pressure changes without the density changing

At higher speeds, the change in air pressure and density becomes significant andis called the compressibility effect . When air enters a venturi at supersonicspeeds, the airflow slows down and must compress in order to pass through itsthroat. Once a fluid compresses, its pressure and density will both increase.

Subsonic AirflowFigure 37

Supersonic AirflowFigure 38

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The transonic flight range encompasses sound wave velocity and consequently isthe most difficult realm of flight since some of the air flowing over the aircraft,particularly the wings, is subsonic and some is supersonic. As the aircraftapproaches the speed of sound, the pressure waves ahead of it will be travellingat the same speed as the aircraft and are therefore relatively stationary. Theyaccumulate to form a continuous pressure wave and consequently will result inthe removal of any advance warning of the approach of the aircraft.

Transonic Flight Pressure WavesFigure 39

At these speeds other pressure waves, or shock waves form wherever the airflowreaches the speed of sound. These waves will upset the aerodynamic balance ofthe wing and this phenomenon will be covered later in the notes.

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1.2.4 SUPERSONIC FLIGHTOnce the aircraft is supersonic, all parts of it are considered to be above thespeed of sound and therefore travelling faster than the rate of propagation of thepressure waves. An infinite number of pressure waves are produced and form acone, the inclination of which will change as the aircraft speed changes.

Mach ConeFigure 40

1.2.4.1 Mach Number

As previously mentioned, Mach number is the ratio of the true airspeed of theaircraft and the local speed of sound at that altitude. An aircraft travelling atexactly the speed of sound is said to be travelling at Mach 1.

It follows therefore that an aircraft travelling at twice the speed of sound would betravelling at Mach 2 and at half the speed of sound, Mach 0.5, etc,.

The following definitions regarding airflow and mach number apply:

Subsonic Flow Mach Numbers below Mach 0.75

Transonic Flow Mach Numbers between Mach 0.75 and Mach 1.2

Supersonic Flow Mach Numbers between Mach 1.2 and 5.0

Hypersonic Flow Mach Numbers above Mach 5.0

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1.2.4.2 Critical Mach Number At any constant aircraft forward speed, the speed of the airflow will vary over thecurves and cambers on the different areas of the airframe. The behaviour of theairflow over the wing will be particularly significant, since this is the major liftprovider for the aircraft.

As air flows over the camber on the upper surface of the wing, its speed willincrease as it flows rearwards from the leading edge, reaching a maximum at thethickest part of the wing chord. This means that although the aircraft itself may betravelling at an airspeed well below Mach 1, the airflow over the thickest part ofthe wing chord, may have already reached Mach 1

As will be discussed later, many unwanted effects occur when the wingapproaches and reaches Mach 1. Therefore, the designers may eitherincorporate features that will lessen the unwanted effects, or limit the aircraft to apredetermined maximum airspeed, that will ensure the wing speed remains belowMach 1 and thus avoids the unwanted effects altogether.

For each aircraft type therefore, a unique maximum aircraft forward speed will becalculated, corresponding to a wing speed of Mach 1. This aircraft speed (alwaysbe less than Mach 1) is called the Critical Mach Number or M. crit and non-supersonic aircraft flying in the transonic flight range, will normally be limited to amaximum speed set below the Critical Mach number.

Critical Mach Number

Figure 41

A thick wing will cause the airflow to speed up over the camber and reach Mach 1more quickly than a thin wing of similar chord length. Consequently, the CriticalMach number for the thinner wing will be a higher value than the thicker wing.

This in turn will mean that the aircraft with a thin wing, will be able to fly faster inthe transonic flight range than the one with the thicker wing, before the unwantedeffects caused by the wing reaching Mach 1 ensue.

Conversely, less lift will be produced by a thin wing, than a thick wing of similar

chord length, but this can be overcome by the so called Supercritical wing chord.

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In this design, the total amount of lift lost by the shallower camber of the thin wing

is restored by making the chord longer. This is perfect for transonic cruiseconditions, but at low airspeeds, lift on a clean wing will be insufficient and soextensive use of high lift devices (slots, slats and flaps) is necessary

Supercritical WingFigure 42

1.2.4.3 Adverse Transonic Effects

Even though the onset of compressibility is gradual, it begins to have a significanteffect as the Critical Mach number is approached. Unwanted adverse effectsincluding, buffeting, shock waves, increase in drag, decrease in lift andmovement of the centre of pressure occur.

If uncontrolled, these effects could result in the aircraft becoming difficult to flyand to behave in a similar manner to a low speed high incidence stall, even

though the aircraft is at high speed and low angle of incidence.

1.2.4.4 Compressibility Buffet

Previously discussed has been the build up of the pressure wave in front of theaircraft as it approaches Mach 1, including the fact that other parts of theairframe, in particular the wing, are likely to reach Mach 1 well before thecomplete aircraft does.

When this occurs the smoothness of the airflow over the wing is severelyaffected. This region, as well as those on the flying control aerofoils, experienceviolent vibration and so-called compressibility buffeting of the airframe. If allowedto continue, control loss or possible structural damage can occur.

1.2.4.5 Shock Wave

Previously in the notes, the build up of pressure waves and the change fromincompressible to compressible flow as the aircraft or an aerofoil surfaceapproaches the speed of sound, has been discussed. Transonic flight presentsmajor design problems for the aerofoil in particular, because only a portion of theairflow passing over the wing becomes supersonic.

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When an aerofoil moves through the air at a speed below its critical Mach

number, all of the airflow is subsonic and the pressure distribution ispredictable.The first indication of a change in the nature of the flow will be abreakaway of the airflow from the aerofoil surface as described previously inboundary layer control. Any turbulence resulting from the separation will cause anincrease in drag and a corresponding reduction in the amount of lift. As speedbegins to increase, the point of separation moves forward, extending the turbulentwake.

Subsonic Flow Over all the SurfaceFigure 43

However, as flight speed reaches and exceeds the critical Mach number, theairflow over the top of the wing speeds up to supersonic velocity and a shockwave starts to form.

The First Sonic Flow is encounteredFigure 44

A Normal Shock Wave Begins to FormFigure 45

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Note: If the aerofoil is symmetrical and set at zero degrees angle of attack, the

incipient shock wave as it is called, would form equally on the upper and lowersurfaces. However, because the wing is usually set to an angle of incidence ofabout 3 degrees, even a symmetrical aerofoil section would produce the incipientwave on the top surface first.

The wave extends outwards more or less at right angles to the aerofoil surfaceand is referred to as a normal (perpendicular) shock wave This normal shockwave forms a boundary between supersonic and subsonic airflow.

As we have seen the high velocity airflow over the top of a wing creates an areaof low pressure. The shock wave causes it to decelerate to subsonic speed,resulting in a rapid rise in pressure. The separation point and turbulent wake willnow start from this point, resulting in a sudden and considerable increase in drag(about 10 times) and therefore a large loss of lift. Severe buffeting is likely, whichcould even lead to a shock stall and the centre of pressure will be altered,affecting the pitching moment.

This extra drag, so called Shock Drag, will be made up of two components,namely Wave Drag, resistance caused by the wave itself and Boundary LayerDrag, due to the increased turbulent region over the surface of the wing.Furthermore, this shock-induced separation is likely to reduce flying controleffectiveness

The velocity of the air leaving the shock wave remains supersonic, so both thestatic pressure and the density of the air increase adding to the high drag/ low liftcondition. Additionally, some of the energy in the airstream will be dissipated inthe form of heat.

As the aircraft speed continues to increase, the wave will extend outwards andbegin to move aft towards the trailing edge of the wing. A second wave begins toform on the lower surface, as the airflow here also speeds up to supersonicvelocity

Shock Induced Separation OccursFigure 46

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As the airspeed reaches the upper end of the transonic range, both shock wavesmove aft, become stronger and will eventually attach to the wing's trailing edge.

Almost all Flow is Supersonic, Some Shock Induced SeparationFigure 47

Further increases in forward speed will now result in the characteristic normalshock wave forming ahead of the aerofoil. This continuous wave, known as aBow wave, will move towards and subsequently attach itself, to the leading edgeof the wing. Once attached, all airflow over the wing will be supersonic and manyof the unwanted transonic effects are eliminated.

The Bow Wave is starting to FormFigure 48

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As can be seen in figure 49, the transonic region has a great affect on the lift and

drag. Both values rise until Mach 0.81, when shock induced separation drasticallyreduces the coefficient of lift. As speed approaches Mach 0.99, a bow wave isforming and airflow over the wing is slowed to subsonic speeds, resulting in anincrease in lift coefficient and a reduction of drag.

Lift / Drag Comparison at 2 Angle of AttackFigure 49

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1.2.5 AERODYNAMIC HEATINGOne of the biggest problems of sustained supersonic flight is aerodynamicheating of the aircraft structure. An extreme example of aerodynamic heatingmight be a shooting star, when its material overheats to the point of destruction,from the heat generated by friction heating with the earth's atmosphere.

In the commercial world, Concorde was probably the only airliner whereaerodynamic heating presents a significant problem. When the aircraft was flownat Mach 2, the friction of the air passing around the aircraft heats the skinconsiderably even at altitudes in excess of 17,000 metres. The point of maximumheating is on the nose where the rise in temperature could reach 175 0C.

As a precaution, a probe on the nose of the aircraft monitors the temperatureduring flight. When a reading of 127 0C is reached, the flight deck is directed toreduce the speed to about Mach 1.8, to bring the temperature back within limits.

Concorde used conventional aluminium alloys in its construction. If future aircraftwere required to travel within the atmosphere at even higher Mach numbers,other materials such as titanium alloy or stainless steel would need to beconsidered.

Concord Skin TemperatureFigure 50

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1.2.6 AREA RULE Area rule is an aerodynamic technique used in the design of high-speed aircraft.

If drag is to be kept to a minimum at transonic speeds, aircraft must be slim,smooth and streamlined. In general terms it means that the wings, fuselage,empennage and other appendages have to be considered together when workingout the total streamlining. This is necessary so that the cross-sectional area ofsuccessive slices of the aircraft from nose to tail, conform to those of a simplebody of streamline shape.

Area rule is defined as: For the minimum drag at the connections,(wing/fuselage), the variation of the aircrafts total cross -sectional area along itslength, should approximate that of an ideal shape having minimum wave drag.

Without area rule, the greatest frontal cross-sectional area of the fuselage wouldoccur where the wings are attached to the fuselage. Therefore, one method ofachieving area rule in this situation is to reduce the cross-sectional area of thefuselage, thereby cancelling out the increase caused by the wings.

Alternatively, the fuselage cross-section could be increased with the use ofenlarged sections behind and in front of the wings to eliminate sudden changes inthe cross-sectional area and achieve the same result.

Area RuleFigure 51

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1.2.7 FACTORS AFFECTING AIRFLOW IN ENGINE INTAKES OF HIGH SPEEDAIRCRAFT

Engine intakes on aircraft that operate in the subsonic flight range only can be ofalmost any form.

The main criteria are that the airflow reaching the compressor stage of the engineduring cruise ideally does not exceed Mach 0.5. This is normally achieved by thecareful design of the intake ducts.

Obviously, if the aircraft never exceeds Mach 0.5, a parallel intake duct could beemployed, but if the aircraft is to cruise at airspeeds in excess of this, yet belowMach 1, a divergent duct must be utilised to slow the airflow at the compressordown to Mach 0.5.

If the aircraft is designed to cruise above Mach 1, the air entering the intakes willbe supersonic and will behave in accordance with the rules of supersonic flow. Inthis case a convergent duct would be necessary to slow down the airflow to thecompressor.

However the aircraft must fly through the transonic range in order to reachsupersonic speed so both types of duct will be necessary.

One way to overcome the problem is to have moveable doors that change theintake duct shape from divergent to convergent cross-section as the aircraft

passes through Mach 1. See figure 52. This technique can be found on theintakes of Concorde.

Other methods to control airflow reaching the compressor is to make use of thefact that air passing through a shock wave slows down to a lower speed. Thistype of intake design is usually characterised by the bullet fairing, which onsome aircraft can translate in and out of the intake to reposition the shock waveduring low or high supersonic flight speeds. See Figure 53

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Intake Moveable doorsFigure 52

Bullet Fairing Intake Figure 53

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1.2.8 EFFECTS OF SWEEPBACK ON CRITICAL MACH NUMBERIn order to fly at high speed in the transonic range without encountering theproblems caused by the production of shock waves, the Critical Mach numberneeds to be as high as possible. As has already been shown, one way is to haveas thin a wing as possible. This of course is an acceptable solution in theory, butin practice there will be structural integrity problems, such as wing loading,strength and flexibility.

Another way of raising the Critical Mach number without the structural limitationsis by the use of swept wings. Sweepback not only delays the production of theshock wave, but reduces the severity of the shock stall should it occur. Thetheory behind this is that it is only the component of velocity over the wing chordthat is responsible for the pressure distribution and so for causing the shock waveto develop. The other velocity component that travels spanwise causes onlyfrictional drag and has no effect on shock wave production.

This theory is borne out by the fact that when it does appear, the shock wave liesparallel to the span of the wing. Therefore only that part of the velocityperpendicular to the shock wave, i.e. across the chord, is reduced by the shockwave to subsonic speeds.

The greater the sweepback, the smaller will be the component of velocityaffected, resulting in a higher Critical Mach number and a reduction in drag at alltransonic speeds. Additionally sweepback results in a thinner mean aerodynamicchord, which raises the Critical Mach number even more.

Effects of SweepbackFigure 54

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STRUCTURESengineering uk

CONTENTS

2 AIRFRAME STRUCTURES GENERAL CONCEPTS ................ 2-12.1 A IRWORTHINESS REQUIREMENTS FOR S TRUCTURAL S TRENGTH . .... 2-12.1.1 S TRUCTURAL CLASSIFICATION ...................................................... 2-1

2.1.2 Primary structure ......... ......... ......... .......... ......... .......... ... 2-22.1.3 Secondary Structure ......... ......... .......... ......... ......... ....... 2-42.1.4 Tertiary Structure ......... .......... ......... ......... .......... ......... .. 2-4

2.2 F AIL S AFE , S AFE LIFE AND D AMAGE TOLERANT CONCEPTS .. ......... . 2-42.2.1 Fail Safe ......... .......... ......... ......... .......... ......... ......... ....... 2-42.2.2 Safe Life ......... .......... ......... ......... .......... ......... ......... ....... 2-42.2.3 Damage Tolerance ......... ......... .......... ......... ......... ......... . 2-5

2.3 Z ONAL AND S TATION IDENTIFICATION S YSTEM ................................ 2-72.3.1 Zonal System ......... ......... .......... ......... .......... ......... ........ 2-72.3.2 Station Identification System ......... .......... ......... ......... .... 2-8

2.4 L OADS FOUND W ITHIN THE S TRUCTURE S TRESS AND S TRAIN .... .. 2-92.4.1 Compression ........ .......... ......... .......... ......... ......... ......... . 2-102.4.2 Tension ......... ......... ......... .......... ......... .......... ......... ........ 2-102.4.3 Bending ......... ......... ......... .......... ......... .......... ......... ........ 2-112.4.4 Torsion .......... ......... ......... .......... ......... .......... ......... ........ 2-122.4.5 Shear ........ .......... ......... .......... ......... ......... ......... .......... .. 2-122.4.6 Hoop Stress ......... .......... ......... .......... ......... ......... ......... . 2-132.4.7 Metal Fatigue ......... ......... .......... ......... .......... ......... ........ 2-13

2.5 D RAINAGE AND VENTILATION P ROVISIONS ..................................... 2-162.5.1 External Drains ......... .......... ......... ......... .......... ......... ..... 2-162.5.2 Internal Drains .......... ......... ......... .......... ......... .......... ...... 2-182.5.3 Ventilation ......... .......... ......... ......... .......... ......... ......... .... 2-18

2.6 L IGHTNING S TRIKE P ROVISION ...................................................... 2-192.7 C ONSTRUCTION METHODS ............................................................ 2-20

2.7.1 Stressed Skin Fuselage ......... ......... ......... ......... .......... .. 2-202.6.1 Frames and Formers ......... ......... .......... ......... ......... ....... 2-212.6.2 Bulkheads ......... .......... ......... ......... .......... ......... ......... .... 2-212.6.3 Longerons and Stringers ......... .......... ......... ......... ......... . 2-222.6.4 Doublers and Reinforcement ......... .......... ......... ......... .... 2-232.6.5 Struts and Ties ......... ......... ......... .......... ......... .......... ...... 2-23

2.6.6 Beams and Floor Structures ......... ......... .......... ......... ..... 2-242.6.7 Methods of Skinning .......... ......... .......... ......... ......... ....... 2-242.6.8 Anti-Corrosive Protection .......... ......... .......... ......... ........ 2-262.6.9 Construction Methods Wing ....................................... 2-272.6.10 Construction Methods Empennage ............ .......... ...... 2-282.6.11 Construction Methods Engine Attachments ......... ....... 2-292.6.12 Structural Assembly Techniques .......... ......... .......... ...... 2-312.6.13 Solid Shank Rivets ......... ......... .......... ......... ......... ......... . 2-312.6.14 Special and Blind Fasteners. ......... .......... ......... ......... .... 2-332.6.15 Bolts and Nuts .......... ......... ......... .......... ......... .......... ...... 2-382.6.16 Adhesive Bonded Structures ......... .......... ......... ......... .... 2-432.6.17 Methods of Surface Protection .......... ......... ......... ......... . 2-452.6.18 Exterior Finish Maintenance ......... ......... .......... ......... ..... 2-47

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2 AIRFRAME STRUCTURES GENERAL CONCEPTS

2.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURALSTRENGTH

Airworthiness requirements are necessary with respect to aircraft structures,because established standards of strength, control, maintainability, etc. willensure that all aircraft will be constructed to the safest possible standard.

Requirements for aircraft above 5700kg MTWA (maximum total weightauthorised) are listed in Joint Airworthiness Requirement 25 (EASA-25) and foraircraft below 5700kg MTWA, in EASA-23. These publications cover not only thebasic requirements, like maximum and minimum 'g' loading, but a vast range ofother requirements with respect to the structure such as:

Control Loads

Door Operation

Effect of Tabs

Factor of Safety

Fatigue

High Lift Devices Stability & Stalling

Ventilation

Weights

The list is all-embracing and provides a useful means of searching for specificstructural details.

2.1.1 STRUCTURAL CLASSIFICATION

For the purpose of assessing damage and the type of repairs to be carried out,the structure of all aircraft is divided into three significant categories:-

Primary structure

Secondary structure

Tertiary structure

Diagrams are prepared by each manufacturer to denote how the variousstructural members fall into these three categories.

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In the manuals of older aircraft the use of colour may be found to identify thethree categories. Primary Structure is shown in Red, Secondary in Yellow and

Tertiary in Green.Note: This system has been discontinued for many years, but with some aircrafthaving a life of 30 or more years and still being operated, it may still be possibleto find the old system in use.

2.1.2 PRIMARY STRUCTURE

This structure includes all portions of aircraft, the failure of which in flight or on theground, would be likely to cause:

Catastrophic structural collapse

Inability to operate a service

Injury to occupants

Loss of control

Unintentional operation of a service

Power unit failure

Examples of some types of primary structure are as follows:

Engine Mountings Fuselage Frames

Main Floor members

Main Spars

Primary Structure Engine mountingsFigure 1

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Primary Structure :Wing SparsFigure 2

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2.1.3 SECONDARY STRUCTURE

This structure includes all portions of the aircraft which would normally beregarded as primary structure, but which unavoidably have such a reserve ofstrength over design requirements that appreciable weakening may be permitted,without risk of failure. It also includes structure which, if damaged, would notimpair the safety of the aircraft as described earlier. Examples of secondarystructure include:

Ribs and parts of skin in the wings.

Skin and stringers in the fuselage

2.1.4 TERTIARY STRUCTURE

This type of structure includes all portions of the structure in which the stressesare low, but which, for various reasons, cannot be omitted from the aircraft.Typical examples include fairings, fillets and brackets which support items in thefuselage and adjacent areas.

2.2 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS

2.2.1 FAIL SAFE

A fail safe structure is one which retains, after initiation of a fracture or crack,sufficient strength for the operation of the aircraft with an acceptable standard ofsafety, until such failure is detected on a normal scheduled inspection.

This is achieved by part and full scale airframe testing and fatigue analysis byusually by the aircraft manufacturer and by subsequent in-service experience.

2.2.2 SAFE LIFE

Safe life structure and components are granted a period of time during which it isconsidered, that failure is extremely unlikely. When deciding its duration, theeffects of wear, fatigue and corrosion must be considered. For example, if testsshow that fatigue will cause a failure in 12,000 flying hours, then one sixth of thismight be quoted as the safe life.(2000 hours then scrapped) If wear or corrosionprove to be the likely cause of failure before 12,000 hours, then one of these willbe the deciding factor.

The safe life time period may be expressed in flying hours, elapsed time, numberof flights or number of applications of load, ie; pressurisation cycles.

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2.2.3 DAMAGE TOLERANCE

The fail safe method has proven to be somewhat unreliable following someaccidents that proved that the concept was not 100% guaranteed. It was also asevere limitation that the addition of extra structural members to protect theintegrity of the structure considerably increased the weight of the aircraft..

The damage tolerant concept, has eliminated much of the extra weight, bydistributing the loads on a particular structure over a larger area. This requires anevaluation of the structure, to provide multiple load paths to carry the loading. Themain advantage is that even with a crack present, the structure will retain itsintegrity and that during scheduled maintenance programmes, the crack will befound before it can become critical.

For example, a wing attachment to the fuselage, which in the past would havebeen designed with one or two large pintle bolts, will now have a larger number ofsmaller bolts in the fitting. The single or dual bolt attachment had to be heavilyreinforced to take the wing loading, adding more weight, whereas the multipleload path can be constructed in a lighter manner, whilst still maintaining itsstrength.

Single Pin AttachmentFigure 3

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Multiple Pin AttachmentFigure 4

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2.3 ZONAL AND STATION IDENTIFICATION SYSTEM

2.3.1 ZONAL SYSTEM

During many different maintenance operations including component changes,structural repairs and trouble shooting, it is necessary to indicate to the engineerwhere, within the structure, the correct location is to be found for the work to becarried out.

When attempting to establish a specific location or identifying components, somemanufacturers make use of two systems, a zonal system and a frame/station

method.The zonal system divides the airframe into a number of zones, (usually less than10), to give engineers and others a rough idea of where they need to look. Thezonal system may also be used in component labelling and work card areaidentification.

In the illustration below, an engineer might have for example a work cardnumbered 500376, indicating it was Job 376 located on the left wing (Zone 500).

Zonal IdentificationFigure 5

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2.3.2 STATION IDENTIFICATION SYSTEM

Most manufacturers use a system of station marking where, for example, theaircraft nose is designated Station 0 and other station designations are located atmeasured distances aft of this point. Component and other locations within thewings, tailplane, fin and nacelles are established from separate dedicatedstations zero.

Fuselage Locations

A particular fuselage station (or frame) would be identified, for example, asStation 5050. This means that if the metric system of measurement is employed,the frame is located at 5.05 metres (5050mm) aft of station zero.

Frame StationsFigure 6

Lateral Locations

To locate structures to the right or left of the aircraft, many manufacturers

consider the fuselage centre line as a station zero. With such a system, the wingor tailplane ribs could be identified as being a particular number of millimetres (orinches) to the right or the left of the centre line.

Vertical Locations

These are usually measured above or below a water line, which is apredetermined reference line passing along the side of the fuselage, usually,somewhere between the floor level and the window line.

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2.4 LOADS FOUND WITHIN THE STRUCTURE STRESS AND STRAIN

Aircraft structural members are designed to carry a load or to resist stress and asingle member may be subjected to a combination of stresses during flight.

When an external force acts on a body, it is opposed by a force within the body.This force is called Stress. If the body is distorted by the stress, it is said to besubject to Strain.

Stress and strain can be defined as follows:

Stress is load or force per unit area acting on a body. Stress = Load or ForceCross Sectional Area

Strain is the distortion per unit length of a body. Strain = DistortionOriginal Length

There are five major stresses and all will be found somewhere within an aircraftstructure. In the design stage, the stresses will have been assessed by thedesigner and the structure made strong enough to carry them adequately.Furthermore, a reserve of strength will also have been included for safety. Thefive types of stress are:

1. Compression

2. Tension

3. Bending (a combination of compression and tension)

4. Twisting/Torsion

5. Shear

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2.4.1 COMPRESSION

Compression is regarded as a primary stress and is the resistance to anyexternal force which tends to push the body together. Compressive stressesapplied to rivets for example, expand the shank as they are driven in, completelyfilling the hole and forming the head to hold sheet metal skins together.

CompressionFigure 7

2.4.2 TENSION

Tension is the primary stress that tends to pull an object apart. A flexible steelcable used in flying control systems is an excellent example of a componentdesigned to withstand tension loads only. It is easily bent, has little opposition tocompression, torsion or shear loads, but has an exceptional strength/weight ratiowhen subjected to a purely tension load.

TensionFigure 8

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2.4.3 BENDING

Bending, when applied to a beam, tends to try to pull one side apart while at thesame time squeezing the other side together. When a person stands on a divingboard, the top of the board is under tension while the bottom is undercompression.

Wing spars of cantilever wings are subject to bending stresses. In flight, the top ofthe spar is being compressed and the bottom is under tension while on theground, the reverse occurs, the top is in tension and the bottom is undercompression. If the wing is supported, the strut will be in tension in flight and incompression on the ground.

BendingFigure 9

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2.4.4 TORSION

A torsional stress is one that is put into a material when it is twisted. When wetwist a structural member, a tensile stress acts diagonally across the member anda compressive stress acts at right angles to the tension. A good example is acrankshaft of an aircraft piston engine which is under a torsional load when theengine is driving the propeller.

TorsionFigure 10

2.4.5 SHEAR

A shear stress is one that resists the tendency to slice a body apart. For examplea clevis bolt in a flying control system is designed to take shear loads only. It isnormally a high strength steel bolt with a thin head and a fat shank. These boltssecure the flexible steel cables to the control surfaces and allow the cable tomove with the control surface without bending. The airload on the control surfaceattempts to slice the bolt apart or shear it.

Rivet Joint in ShearFigure 11

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2.4.6 HOOP STRESS

An aircraft which has its fuselage pressurised inside to allow the carriage ofpassengers at altitude, will have other stresses acting on the fuselage skin. Thecircumferential load about the fuselage is known as hoop stress and resisted bythe fuselage frames and tension in the so called stressed skin. The longitudinal(axial) load along the fuselage is also resisted by tension in the skin and by thelongerons and stringers.

Hoop stressFigure 12

2.4.7 METAL FATIGUE

The phenomenon of metal fatigue has long been known, but has become ofgreater concern in recent years with aircraft which remain in service long aftertheir original expected fatigue life has expired.

It is relatively easy to design a structure to withstand a steady load, but aircraftare subjected to widely varying loads in flight and many components experienceload reversals, an example being the wings, where the aerodynamic forcesduring flight manoeuvres cause tension and compression loads to alternatecontinually. Unfortunately, any metal part subjected to a wide variation or reversal

of even a relatively small load is gradually and progressively weakened.The subject was vividly highlighted in 1954, with another type of load reversal,that of pressurisation cycles of the passenger cabin. which resulted in a numberof disastrous accidents with the De-Havilland Comet airliner. Small fatigue cracksin the fuselage skin accumulated around the corners of the square shapedwindows and hatches and led to a fatal explosive decompression of the cabin.

Following the incidents the most extensive research to this hitherto unwarrantedmenace was undertaken, and led to fatigue loading being included into futuredesign considerations.

Metal fatigue refers to the loss of strength, or resistance to load, experienced by acomponent or structure as the number of load cycles or load reversals increases.

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Load reversals refer to a material being continually loaded and unloaded and aslong as the elastic limit is not exceeded, the material should be unaffected and

return to its original state.In reality, however, the load application may result in minute, seeminglyinconsequential cracks, which, as the cycles continue, get larger and join up withother, newer cracks. Eventually, after many cycles, the cumulative effect will besuch that the strength of the metal will be compromised and could result incatastrophic failure.

The fatigue strength of a metal can be found by experimentation on full scalefatigue rigs, which can be subjected to a programme of load reversals, 24 hours aday, 365 days a year, to accumulate information and a fatigue life, years ahead ofthe oldest aircraft of the particular type in the fleet.

How the in-service aircraft subsequently consumes this fatigue index, depends onits operating theatre. For example, the number of times the pressurisation cyclesare applied to aircraft on long or short haul flights, steep or conventional take offand landing etc., are taken into account to calculate fatigue life consumed.

Stress amplitude can be plotted against endurance for one particular value ofmean stress, the so- called S/N Curve. Using a chart such as this, it can bedetermined at what point, in cycles, the metal has reached its minimumacceptable strength. This will be the ultimate fatigue life and is normally allotted afatigue index of 100.

Fatigue GraphFigure 13

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Even when the fatigue index of 100 is eventually reached on each individual

aircraft, the designers can extend it beyond 100, by examining, as previouslymentioned, how the fatigue was consumed and recommending specific structuralinspection and possibly strengthening or replacement of fittings and components.

Fatigue is a natural phenomenon and cannot be prevented. The ability tocorrectly predict its effects and take the necessary action is the problem faced bythe aircraft design and maintenance personnel. Different metals have differentfatigue characteristics and the way parts are designed, also affects their fatiguelife. Fastener holes, sharp changes in thickness and small seemingly insignificantcracks for example, can directly affect the fatigue life of a part.

Fatigue cracking can also accelerate the onset of corrosion, by exposing

unprotected metal to the elements. The crack growth and the consequentialincrease in corrosion, can cause serious structural problems over a relativelyshort period. With the ageing of the airliner fleet, a number of extra inspections,including non-destructive testing and structural sampling techniques have beenintroduced. The maintenance technician must carefully monitor the aircraftstructure, paying particular attention to the integrity of surface finish and generalcorrosion.

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2.5 DRAINAGE AND VENTILATION PROVISIONS

Drainage

The aircraft structure requires many different types of drain holes and paths toprevent water and other fluids such as fuel, hydraulic oil etc., from collectingwithin the structure. These could become both a corrosion and fire hazard.

The forms of drainage can be divided into two areas.

1. External drains

2. Internal drains

2.5.1 EXTERNAL DRAINS

These ports are located on exterior surfaces of the fuselage, wing andempennage to ensure fluids are dumped overboard. In small unpressurisedaircraft and unpressurised areas of larger airliners, these drains may bepermanently open. However, in pressurised aircraft, the cabin air would leakuncontrollably through the drains and so it is necessary to use drain valves toprevent loss of cabin pressure.

There are a number basic types of drain valve used for this purpose.

Two similar types rely upon pressurised air in the cabin to keep the valve closed.One valve has a rubber flapper seal and the other a spring loaded valve seal.Normally located on the keel of the fuselage, both are open when the aircraft isunpressurised on the ground, allowing the fluids to drain overboard. During flight,the increased air pressure in the cabin closes the valves, thus preventing anypressurisation losses. These valves are shown below, where it can also be seenthat a levelling compound has been used in areas which might become fluidtraps. This compound is usually a rubberised sealant which fills the cavity,bringing the level up to the lip of the drain hole.

Fuselage DrainsFigure 14

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Another similar type of drain valve also uses the cabin air pressure to close offthe drain path, this time by moving the plunger down to seal the drain. This valvewill also be open when cabin pressure is removed.

Fuselage DrainsFigure 15

Fluids from some places, such as galleys and wash basins, require more thansimple drain holes. The temperature at cruising altitude can fall to -60C andwater draining overboard could freeze and cause blockage problems.

The method used in these cases are drain masts, which are like small aerofoilsprojecting from the bottom of the aircraft skin, on the centre line, through whichthe water is discharged. The drain masts are heated to prevent icing and alsodischarge the liquids well away from the aircraft's skin.

Boeing 747 Drain MastsFigure 16

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2.5.2 INTERNAL DRAINS

To enable the external drains to function as designed, means must be providedwithin the various locations of the airframe and powerplant installation, to ensurethat all fluids are directed towards the site of the external drain points. This isachieved by using internal drain paths and drain holes.

The internal structure is provided with tubes, channels, dams and drain holes, todirect the flow of fluid towards the external drain points. All structural membersare designed so that they do not trap fluids by ensuring, for example, that alllightening holes and ribs face downwards, allowing fluids to run off them.

2.5.3 VENTILATION

It is essential that the internal cavities within the structure are properly vented toprevent the build up of flammable vapour from the drain lines and to allow anyother moisture residue to properly evaporate.

Consequently sumps, tanks and cavities will all be pr

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