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Contents1 Basic mechanics 12 Working cycle and airflow 113 Compressors 194 Combustion chambers 355 Turbines 456 Exhaust system 597 Accessory drives 658 Lubrication 739 Internal air system 85

10 Fuel system 9511 Starting and ignition 12112 Controls and instrumentation 13313 Ice protection 14714 Fire protection 15315 Thrust reversal 15916 Afterburning 16917 Water injection 18118 Vertical/short take-off and landing 18719 Noise suppression 19920 Thrust distribution 20721 Performance 21522 Manufacture 22923 Power plant installation 24324 Maintenance 25125 Overhaul 263

Appendix 1; Conversion factors 277

Rolls-Royce Trent 800

Developed from the RB211, the Trent covers a thrust range of 71,000 lb to 92,000 lb thrust, with the capabilityto grow beyond 100,000 lb. The Trent 800 features a 110 inch diameter wide-chord fan, high flow compressorsand Full Authority Digital Engine Control (FADEC).

Detailed engineering design began in 1988 to meet the propulsion requirements of the Airbus A330 (Trent 700)and Boeing 777 (Trent 800). The Trent first ran in August 1990, and in January 1994 a Trent 800 demonstrateda world record thrust of 106,087 lb.

The engine entered service in March 1995 in the Airbus A330.

IntroductionThis book has been written to provide a simple andself-contained description of the working andunderlying principles of the aero gas turbine engine.The use of complex formulae and the language ofthe specialist have been avoided to allow for a clearand concise presentation of the essential facts. Onlysuch description and formulae, therefore, as arenecessary to the understanding of the function andthe theory of the engine are included.It will be noted that the emphasis in this book is onthe turbo-jet engine and that no special part dealswith the propeller-turbine engine. This is because theworking principles of both engine types areessentially the same. However where differences infunction or application do exist, these are described.The aero gas turbine is being continually developedto provide improved performance for each newgeneration of aircraft; the fourth edition of this bookhas been revised and expanded to include the latestaero gas engine technology.

Rolls-Royce RB183 Mk 555

On 1 April, 1943, Rolls-Royce assumedresponsibility for the Power Jets W2B which, amonth earlier, had made its first flight in theGloster E28/39 at 1200lb thrust. Later knownas the B23 Welland it was, during April, putthrough a 100 hr test at the design rating of1600 Ib thrust. In June, 1943, it flew in aGloster Meteor at 1400lb thrust. ProductionWelland-Meteors were in action against V-1flying bombs in August 1944.

Rolls-Royce B23 Welland

1: Basic mechanics

Contents Page

Introduction 1Principles of jet propulsion 2Methods of jet propulsion 3


1. The development of the gas turbine engine as anaircraft power plant has been so rapid that it isdifficult to appreciate that prior to the 1950s very fewpeople had heard of this method of aircraftpropulsion. The possibility of using a reaction jet hadinterested aircraft designers for a long time, butinitially the low speeds of early aircraft and theunsuitably of a piston engine for producing the largehigh velocity airflow necessary for the �jet� presentedmany obstacles.

2. A French engineer, René Lorin, patented a jetpropulsion engine (fig. 1-1) in 1913, but this was anathodyd (para. 11) and was at that period impossibleto manufacture or use, since suitable heat resistingmaterials had not then been developed and, in thesecond place, jet propulsion would have beenextremely inefficient at the low speeds of the aircraftof those days. However, today the modern ram jet isvery similar to Lorin's conception.

3. In 1930 Frank Whittle was granted his first patentfor using a gas turbine to produce a propulsive jet,

but it was eleven years before his engine completedits first flight. The Whittle engine formed the basis ofthe modern gas turbine engine, and from it wasdeveloped the Rolls-Royce Welland, Derwent, Neneand Dart engines. The Derwent and Nene turbo-jetengines had world-wide military applications; theDart turbo-propeller engine became world famous asthe power plant for the Vickers Viscount aircraft.Although other aircraft may be fitted; with laterengines termed twin-spool, triple-spool, by-pass,ducted fan, unducted fan and propfan, these areinevitable developments of Whittle's early engine.


Fig. 1-1 Lorin's jet engine.

4. The jet engine (fig. 1-2), although appearing sodifferent from the piston engine-propellercombination, applies the same basic principles toeffect propulsion. As shown in fig. 1-3, both propeltheir aircraft solely by thrusting a large weight of airbackwards.

5. Although today jet propulsion is popularly linkedwith the gas turbine engine, there are other types ofjet propelled engines, such as the ram jet, the pulsejet, the rocket, the turbo/ram jet, and the turbo-rocket.


6. Jet propulsion is a practical application of SirIsaac Newton's third law of motion which states that,'for every force acting on a body there is an oppositeand equal reaction'. For aircraft propulsion, the 'body'is atmospheric air that is caused to accelerate as itpasses through the engine. The force required togive this acceleration has an equal effect in theopposite direction acting on the apparatus producingthe acceleration. A jet engine produces thrust in a

similar way to the engine/propeller combination. Bothpropel the aircraft by thrusting a large weight of airbackwards (fig. 1-3), one in the form of a large airslipstream at comparatively low speed and the otherin the form of a jet of gas at very high speed.

7. This same principle of reaction occurs in all formsof movement and has been usefully applied in manyways. The earliest known example of jet reaction isthat of Hero's engine (fig. 1-4) produced as a toy in120 B.C. This toy showed how the momentum ofsteam issuing from a number of jets could impart anequal and opposite reaction to the jets themselves,thus causing the engine to revolve.

8. The familiar whirling garden sprinkler (fig. 1-5) isa more practical example of this principle, for themechanism rotates by virtue of the reaction to thewater jets. The high pressure jets of modern fire-fighting equipment are an example of 'jet reaction',for often, due to the reaction of the water jet, the hosecannot be held or controlled by one fireman. Perhapsthe simplest illustration of this principle is afforded bythe carnival balloon which, when the air or gas isreleased, rushes rapidly away in the directionopposite to the jet.

9. Jet reaction is definitely an internal phenomenonand does not, as is frequently assumed, result fromthe pressure of the jet on the atmosphere. In fact, the

jet propulsion engine, whether rocket, athodyd, orturbo-jet, is a piece of apparatus designed toaccelerate a stream of air or gas and to expel it athigh velocity. There are, of course, a number of ways

Basic mechanics


Fig. 1-2 A Whittle-type turbo-jet engine.

Fig. 1-3 Propeller and jet propulsion.

of doing this, as described in Part 2, but in allinstances the resultant reaction or thrust exerted onthe engine is proportional to the mass or weight of airexpelled by the engine and to the velocity changeimparted to it. In other words, the same thrust can beprovided either by giving a large mass of air a littleextra velocity or a small mass of air a large extra

velocity. In practice the former is preferred, since bylowering the jet velocity relative to the atmosphere ahigher propulsive efficiency is obtained.


10. The types of jet engine, whether ram jet, pulsejet, rocket, gas turbine, turbo/ram jet or turbo-rocket,differ only in the way in which the 'thrust provider', orengine, supplies and converts the energy into powerfor flight.

11. The ram jet engine (fig. 1-6) is an athodyd, or'aero-thermodynamic-duct to give it its full name. Ithas no major rotating parts and consists of a ductwith a divergent entry and a convergent or

convergent-divergent exit. When forward motion isimparted to it from an external source, air is forcedinto the air intake where it loses velocity or kineticenergy and increases its pressure energy as itpasses through the diverging duct. The total energyis then increased by the combustion of fuel, and theexpanding gases accelerate to atmosphere throughthe outlet duct. A ram jet is often the power plant formissiles and .target vehicles; but is unsuitable as anaircraft power plant "because it requires forwardmotion imparting to it before any thrust is produced.

12. The pulse jet engine (fig. 1-7) uses the principleof intermittent combustion and unlike the ram jet itcan be run at a static condition. The engine is formedby an aerodynamic duct similar to the ram jet but,due to the higher pressures involved, it is of morerobust construction. The duct inlet has a series ofinlet 'valves' that are spring-loaded into the openposition. Air drawn through the open valves passesinto the combustion chamber and is heated by theburning of fuel injected into the chamber. Theresulting expansion causes a rise in pressure, forcing

Basic mechanics


Fig. 1-4 Hero�s engine - probably the earliestform of jet reaction.

Fig. 1-5 A garden sprinkler rotated by thereaction of the water jets.

Fig. 1-6 A ram Jet engine.

the valves to close, and the expanding gases arethen ejected rearwards. A depression created by theexhausting gases allows the valves to open andrepeat the cycle. Pulse jets have been designed forhelicopter rotor propulsion and some dispense withinlet valves by careful design of the ducting to controlthe changing pressures of the resonating cycle. Thepulse jet is unsuitable as an aircraft power plantbecause it has a high fuel consumption and is unableto equal the performance of the modern gas turbineengine.

13. Although a rocket engine (fig. 1-8) is a jetengine, it has one major difference in that it does notuse atmospheric air as the propulsive fluid stream.Instead, it produces its own propelling fluid by thecombustion of liquid or chemically decomposed fuelwith oxygen, which it carries, thus enabling it tooperate outside the earth's atmosphere. It is,therefore, only suitable for operation over shortperiods.

14. The application of the gas turbine to jetpropulsion has avoided the inherent weakness of therocket and the athodyd, for by the introduction of aturbine-driven compressor a means of producingthrust at low speeds is provided. The turbo-jet engineoperates on the 'working cycle' as described in Part2. It draws air from the atmosphere and aftercompressing and heating it, a process that occurs inall heat engines, the energy and momentum given tothe air forces It out of the propelling nozzle at avelocity of up to 2,000 feet per second or about 1,400miles per hour. On its way through the engine, the airgives up some of its energy and momentum to drivethe turbine that powers the compressor.

15. The mechanical arrangement of the gas turbineengine is simple, for it consists of only two mainrotating parts, a compressor (Part 3) and a turbine(Part 5), and one or a number of combustionchambers (Part 4). The mechanical arrangement ofvarious gas turbine engines is shown in fig. 1 -9. Thissimplicity, however, does not apply to all aspects ofthe engine, for as described in subsequent Parts thethermo and aerodynamic problems are somewhatcomplex. They result from the high operating tem-peratures of the combustion chamber and turbine,the effects of varying flows across the compressor

Basic mechanics


Fig. 1-7 A pulse jet engine.

Fig. 1-8 A rocket engine.

Basic mechanics


Fig. 1-9-1 Mechanical arrangement of gas turbine engines.

Basic mechanics


Fig. 1-9-2 Mechanical arrangement of gas turbine engines.

and turbine blades, and the design of the exhaustsystem through which the gases are ejected to formthe propulsive jet.

16. At aircraft speeds below approximately 450miles per hour, the pure jet engine is less efficientthan a propeller-type engine, since its propulsiveefficiency depends largely on its forward speed; thepure turbo-jet engine is, therefore, most suitable forhigh forward speeds. The propeller efficiency does,however, decrease rapidly above 350 miles per hourdue to the disturbance of the airflow caused by thehigh blade-tip speeds of the propeller. These charac-

teristics have led to some departure from the use ofpure turbo-jet propulsion where aircraft operate atmedium speeds by the introduction of a combinationof propeller and gas turbine engine.

17. The advantages of the propeller/turbinecombination have to some extent been offset by theintroduction of the by-pass, ducted fan and propfanengines. These engines deal with larger comparativeairflows and lower jet velocities than the pure jetengine, thus giving a propulsive efficiency (Part 21)which is comparable to that of the turbo-prop andexceeds that of the pure jet engine (fig. 1-10).

Basic mechanics


Fig. 1-10 Comparative propulsive efficiencies.

18. The turbo/ram jet engine (fig. 1-11) combinesthe turbo-jet engine (which is used for speeds up toMach 3) with the ram jet engine, which has goodperformance at high Mach numbers.

19. The engine is surrounded by a duct that has avariable intake at the front and an afterburning jetpipe with a variable nozzle at the rear. During take-off and acceleration, the engine functions as a con-

ventional turbo-jet with the afterburner lit; at otherflight conditions up to Mach 3, the afterburner isinoperative. As the aircraft accelerates through Mach3, the turbo-jet is shut down and the intake air isdiverted from the compressor, by guide vanes, andducted straight into the afterburning jet pipe, whichbecomes a ram jet combustion chamber. This engineis suitable for an aircraft requiring high speed and

Basic mechanics


Fig. 1-12 A turbo-rocket engine.

Fig. 1-11 A turbo/ram jet engine.

sustained high Mach number cruise conditionswhere the engine operates in the ram jet mode.

20. The turbo-rocket engine (fig. 1-12) could beconsidered as an alternative engine to the turbo/ramjet; however, it has one major difference in that itcarries its own oxygen to provide combustion,

21. The engine has a low pressure compressordriven by a multi-stage turbine; the power to drive theturbine is derived from combustion of kerosine andliquid oxygen in a rocket-type combustion chamber.Since the gas temperature will be in the order of3,500 deg. C, additional fuel is sprayed into the

combustion chamber for cooling purposes before thegas enters the turbine. This fuel-rich mixture (gas) isthen diluted with air from the compressor and thesurplus fuel burnt in a conventional afterburningsystem.

22. Although the engine is smaller and lighter thanthe turbo/ram jet, it has a higher fuel consumption.This tends to make it more suitable for an interceptoror space-launcher type of aircraft that requires highspeed, high altitude performance and normally has aflight plan that is entirely accelerative and of shortduration.

Basic mechanics


Rolls-Royce/Snecma Olympus

Rolls-Royce RB37 Derwent 1

A straight-through version of the reverse-flowPower Jets W2B, known as the W2B/26, wasdeveloped by the Rover Company from 1941to 1943. Taken over by Rolls-Royce in April1943 and renamed the Derwent, it passed a100hr. test at 2000 lb thrust in November 1943and was flown at that rating in April 1944. Theengine powered the Gloster Meteor III whichentered service in 1945.

2: Working cycle and airflow

Contents Page

Introduction 11 Working cycle 11The relations between pressure,volume and temperature 13 Changes in velocity and pressure 14Airflow 17


1. The gas turbine engine is essentially a heatengine using air as a working fluid to provide thrust.To achieve this, the air passing through the enginehas to be accelerated; this means that the velocity orkinetic energy of the air is increased. To obtain thisincrease, the pressure energy is first of all increased,followed by the addition of heat energy, before finalconversion back to kinetic Energy in the form of ahigh velocity jet efflux.


2. The working cycle of the gas turbine engine issimilar to that of the four-stroke piston engine.However, in the gas turbine engine, combustionoccurs at a constant pressure, whereas in the pistonengine it occurs at a constant volume. Both enginecycles (fig. 2-1) show that in each instance there isinduction, compression, combustion and exhaust.These processes are intermittent in the case of the

piston engine whilst they occur continuously in thegas turbine. In the piston engine only one stroke isutilized in the production of power, the others beinginvolved in the charging, compressing andexhausting of the working fluid. In contrast, theturbine engine eliminates the three 'idle' strokes, thusenabling more fuel to be burnt in a shorter time;hence it produces a greater power output for a givensize of engine.

3. Due to the continuous action of the turbineengine and the fact that the combustion chamber isnot an enclosed space, the pressure of the air doesnot rise, like that of the piston engine, duringcombustion but its volume does increase. Thisprocess is known as heating at constant pressure.Under these conditions there are no peak orfluctuating pressures to be withstood, as is the casewith the piston engine with its peak pressures inexcess of 1,000 lb. per sq. in. It is these peakpressures which make it necessary for the pistonengine to employ cylinders of heavy construction and


to use high octane fuels, in contrast to the low octanefuels and the light fabricated combustion chambersused on the turbine engine.

4. The working cycle upon which the gas turbineengine functions is, in its simplest form, representedby the cycle shown on the pressure volume diagramin fig. 2-2. Point A represents air at atmosphericpressure that is compressed along the line AB. FromB to C heat is added to the air by introducing andburning fuel at constant pressure, thereby consider-ably increasing the volume of air. Pressure losses inthe combustion chambers (Part 4) are indicated bythe drop between B and C. From C to D the gasesresulting from combustion expand through theturbine and jet pipe back to atmosphere. During thispart of the cycle, some of the energy in theexpanding gases is turned into mechanical power by

Working cycle and airflow


Fig. 2-1 A comparison between the working cycle of a turbo-jet engine and a piston engine.

Fig. 2-2 The working cycle on a pressure-volume diagram.

the turbine; the remainder, on its discharge toatmosphere, provides a propulsive jet.

5. Because the turbo-jet engine is a heat engine,the higher the temperature of combustion the greateris the expansion of the gases. The combustiontemperature, however, must not exceed a value thatgives a turbine gas entry temperature suitable for thedesign and materials of the turbine assembly.

6. The use of air-cooled blades in the turbineassembly permits a higher gas temperature and aconsequently higher thermal efficiency.


7. During the working cycle of the turbine engine,the airflow or 'working fluid' receives and gives upheat, so producing changes in its pressure, volumeand temperature. These changes as they occur areclosely related, for they follow a common principlethat is embodied in a combination of the laws ofBoyle and Charles. Briefly, this means that theproduct of the pressure and the volume of the air atthe various stages in the working cycle is proportion-al to the absolute temperature of the air at those

Working cycle and airflow


Fig. 2-3 An airflow through divergent and convergent ducts.

stages. This relationship applies for whatever meansare used to change the state of the air. For example,whether energy is added by combustion or bycompression, or is extracted by the turbine, the heatchange is directly proportional to the work added ortaken from the gas.

8. There are three main conditions in the engineworking cycle during which these changes occur.During compression, when work is done to increasethe pressure and decrease the volume of the air,there is a corresponding rise in the temperature.During combustion, when fuel is added to the air andburnt to increase the temperature, there is a corre-sponding increase in volume whilst the pressureremains almost constant. During expansion, whenwork is taken from the gas stream by the turbineassembly, there is a decrease in temperature andpressure with a corresponding increase in volume.

9. Changes in the temperature and pressure of theair can be traced through an engine by using theairflow diagram in fig. 2-5. With the airflow beingcontinuous, volume changes are shown up aschanges in velocity.

10. The efficiency with which these changes aremade will determine to what extent the desiredrelations between the pressure, volume andtemperature are attained. For the more efficient thecompressor, the higher the pressure generated for agiven work input; that is, for a given temperature riseof the air. Conversely, the more efficiently the turbineuses the expanding gas, the greater the output ofwork for a given pressure drop in the gas.

11. When the air is compressed or expanded at 100per cent efficiency, the process is said to beadiabatic. Since such a change means there is noenergy losses in the process, either by friction,conduction or turbulence, it is obviously impossibleto achieve in practice; 90 per cent is a good adiabaticefficiency for the compressor and turbine.


12. During the passage of the air through theengine, aerodynamic and energy requirementsdemand changes in its velocity and pressure. Forinstance: during compression, a rise in the pressureof the air is required and not an increase in itsvelocity. After the air has been heated and its internalenergy increased by combustion, an increase in thevelocity of the gases is necessary to force the turbineto rotate. At the propelling nozzle a high exit velocityis required, for it is the change in the momentum of

the air that provides the thrust on the aircraft. Localdecelerations of airflow are also required, as forinstance, in the combustion chambers to provide alow velocity zone for the flame to burn.

13. These various changes are effected by meansof the size and shape of the ducts through which theair passes on its way through the engine. Where aconversion from velocity (kinetic) energy to pressureis required, the passages are divergent in shape.Conversely, where it is required to convert the energystored in the combustion gases to velocity energy, aconvergent passage or nozzle (fig. 2-3) is used.These shapes apply to the gas turbine engine wherethe airflow velocity is subsonic or sonic, i.e. at thelocal speed of sound. Where supersonic speeds areencountered, such as in the propelling nozzle of therocket, athodyd and some jet engines (Part 6), aconvergent-divergent nozzle or venturi (fig. 2-4) isused to obtain the maximum conversion of theenergy in the combustion gases to kinetic energy.

14. The design of the passages and nozzles is ofgreat importance, for upon their good design willdepend the efficiency with which the energy changesare effected. Any interference with the smooth airflowcreates a loss in efficiency and could result incomponent failure due to vibration caused by eddiesor turbulence of the airflow.

Working cycle and airflow


Fig. 2-4 Supersonic airflow through aconvergent-divergent nozzle orventuri.

Working cycle and airflow


Fig. 2-5-1 Airflow systems.

Working cycle and airflow


Fig, 2-5-2 Airflow systems.


15. The path of the air through a gas turbine enginevaries according to the design of the engine. Astraight-through flow system (fig. 2-5) is the basicdesign, as it provides for an engine with a relativelysmall frontal area and is also suitable for use of theby-pass principle. In contrast, the reverse flowsystem gives an engine with greater frontal area, butwith a reduced overall length. The operation,however, of all engines is similar. The variations dueto the different designs are described in thesubsequent paragraphs.

16. The major difference of a turbo-propeller engineis the conversion of gas energy into mechanicalpower to drive the propeller. Only a small amount of'jet thrust' is available from the exhaust system. Themajority of the energy in the gas stream is absorbedby additional turbine stages, which drive the propellerthrough internal shafts (Part 5).

17. As can be seen in fig. 2-5, the by-pass principleinvolves a division of the airflow. Conventionally, allthe air taken in is given an initial low compressionand a percentage is then ducted to by-pass, theremainder being delivered to the combustion systemin the usual manner. As described in Part 21, this

principle is conducive to improved propulsiveefficiency and specific fuel consumption.

18. An important design feature of the by-passengine is the by-pass ratio; that is, the ratio of cool airby-passed through the duct to the flow of air passedthrough the high pressure system. With low by-passratios, i.e. in the order of 1:1, the two streams areusually mixed before being exhausted from theengine. The fan engine may be regarded as anextension of the by-pass principle, and therequirement for high by-pass ratios of up to 5:1 islargely met by using the front fan in a twin or triple-spool configuration (on which the fan is, in fact, thelow pressure compressor) both with and withoutmixing of the airflows. Very high by-pass ratios, in theorder of 15:1, are achieved using propfans. Theseare a variation on the turbo-propeller theme but withadvanced technology propellers capable of operatingwith high efficiency at high aircraft speeds.

19. On some front fan engines, the by-passairstream is ducted overboard either directly behindthe fan through short ducts or at the rear of theengine through longer ducts; hence the term 'ductedfan'. Another, though seldom used, variation is that ofthe aft fan.

Working cycle and airflow


De Havilland H1 Goblin

Development of the de Havilland Goblinbegan in 1941 with the Halford H1 with adesign thrust of 3000 lb. The engine passed a25 hr special category test in September 1942and was cleared for flight at 2000 lb thrust.This took place in a Gloster Meteor on 5March 1943 and was also the first flight of thataircraft type. In September 1943 the first flightof a de Havilland DH100 Vampire was madewith a Goblin of 2300 lb thrust.

Rolls-Royce RB211-22B

3: Compressors


1. In the gas turbine engine, compression of the airbefore expansion through the turbine is effected byone of two basic types of compressor, one givingcentrifugal flow and the other axial flow. Both typesare driven by the engine turbine and are usuallycoupled direct to the turbine shaft.

2. The centrifugal flow compressor (fig. 3-1) is asingle or two stage unit employing an impeller to

accelerate the air and a diffuser to produce therequired pressure rise. The axial flow compressor(fig. 3-7 and fig. 3-8) is a multi-stage unit employingalternate .rows of rotating (rotor) blades andstationary (stator) vanes, to accelerate and diffusethe air until the required pressure rise is obtained. Insome cases, particularly on small engines, an axialcompressor is used to boost the inlet pressure to thecentrifugal.

3. With regard to the advantages and disadvan-tages of the two types, the centrifugal compressor isusually more robust than the axial compressor and isalso easier to develop and manufacture. The axialcompressor however consumes far more air than a

Contents Page

Introduction 19The centrifugal flow compressor 21

Principles of operation Construction Impellers Diffusers

The axial flow compressor 22Principles of operation Construction RotorsRotor blades Stator vanes

Operating conditions 28Airflow control 29Materials 29Balancing 33


centrifugal compressor of the same frontal area andcan be designed to attain much higher pressureratios. Since the air flow is an important factor indetermining the amount of thrust, this means theaxial compressor engine will also give more thrust forthe same frontal area. This, plus the ability toincrease the pressure ratio by addition of extrastages, has led to the adoption of axial compressorsin most engine designs. However, the centrifugalcompressor is still favoured for smaller engineswhere its simplicity and ruggedness outweigh anyother disadvantages.

4. The trend to high pressure ratios which hasfavoured the adoption of axial compressors isbecause of the improved efficiency that results,

which in turn leads to improved specific fuelconsumption for a given thrust, ref. fig. 3-2.



Fig. 3-1 A typical centrifugal flow compressor.

Fig. 3-2 Specific fuel consumption andpressure ratio.


5. Centrifugal flow compressors have a single ordouble-sided impeller and occasionally a two-stage,single sided impeller is used, as on the Rolls-RoyceDart. The impeller is supported in a casing that alsocontains a ring of diffuser vanes. If a double-entryimpeller is used, the airflow to the _rear side isreversed in direction and a plenum chamber isrequired.

Principles of operation6. The impeller is rotated at high speed by theturbine and air is continuously induced into thecentre of the impeller. Centrifugal action causes it toflow radially outwards along the vanes to the impellertip, thus accelerating the air and also causing a risein pressure to occur. The engine intake duct maycontain vanes that provide an initial swirl to the airentering the compressor.

7. The air, on leaving the impeller, passes into thediffuser section where the passages form divergentnozzles that convert most of the kinetic energy intopressure energy, as illustrated in fig. 3-3. In practice,it is usual to design the compressor so that about halfof the pressure rise occurs in the impeller and half inthe diffuser.

8. To maximize the airflow and pressure risethrough the compressor requires the impeller to berotated at high speed, therefore impellers aredesigned to operate at tip speeds of up to 1,600 ft.

per sec. By operating at such high tip speeds the airvelocity from the impeller is increased so that greaterenergy is available for conversion to pressure.

9. To maintain the efficiency of the compressor, it isnecessary to prevent excessive air leakage betweenthe impeller and the casing; this is achieved bykeeping their clearances as small as possible (fig. 3-4).

Construction10. The construction of the compressor centresaround the impeller, diffuser and air intake system.The impeller shaft rotates in ball and roller bearingsand is either common to the turbine shaft or split inthe centre and connected by a coupling, which isusually designed for ease of detachment.

Impellers11. The impeller consists of a .forged, disc withintegral, radially disposed vanes on one or both sides(fig. 3-5) forming convergent passages in conjunctionwith the compressor casing. The vanes may beswept back, but for ease of manufacture straight



Fig. 3-3 Pressure and velocity changesthrough a centrifugal compressor.

Fig. 3-4 Impeller working clearance andair leakage.

radial vanes are usually employed. To ease the airfrom axial flow in the entry duct on to the rotatingimpeller, the vanes in the centre of the impeller arecurved in the direction of rotation. The curvedsections may be integral with the radial vanes orformed separately for easier and more accuratemanufacture.

Diffusers12. The diffuser assembly may be an integral part ofthe compressor casing or a separately attachedassembly. In each instance it consists of a number ofvanes formed tangential to the impeller. The vanepassages are divergent to convert the kinetic energyinto pressure energy and the inner edges of the

vanes are in line with the direction of the resultantairflow from the impeller (fig. 3-6). The clearancebetween the impeller and the diffuser is an importantfactor, as too small a clearance will set upaerodynamic buffeting impulses that could betransferred to the impeller and create an unsteadyairflow and vibration.


13. An axial flow compressor (fig. 3-7 and fig. 3-8)consists of one or more rotor assemblies that carryblades of airfoil section. These assemblies aremounted between bearings in the casings whichincorporate the stator vanes. The compressor is amulti-stage unit as the amount of pressure increaseby each stage is small; a stage consists of a row ofrotating blades followed by a row of stator vanes.Where several stages of compression operate inseries on one shaft it becomes necessary to vary thestator vane angle to enable the compressor tooperate effectively at speeds below the designcondition. As the pressure ratio is increased theincorporation of variable stator vanes ensures thatthe airflow is directed onto the succeeding stage ofrotor blades at an acceptable angle, ref. para. 30,Airflow Control.

14. From the front to the rear of the compressor, i.e.from the low to the high pressure end, there is agradual reduction of the air annulus area between



Fig. 3-5 Typical impellers for centrifugalcompressors.

Fig. 3-6 Airflow at entry to diffuser.



Fig. 3-7 Typical axial flow compressors.

the rotor shaft and the stator casing. This isnecessary to maintain a near constant air axialvelocity as the density increases through the lengthof the compressor. The convergence of the airannulus is achieved by the tapering of the casing orrotor. A combination of both is also possible, with thearrangement being influenced by manufacturingproblems and other mechanical design factors.

15. A single-spool compressor (fig. 3-7) consists ofone rotor assembly and stators with as many stagesas necessary to achieve the desired pressure ratioand all the airflow from the intake passes through thecompressor.

16. The multi-spool compressor consists of two ormore rotor assemblies, each driven by their own

turbine at an optimum speed to achieve higherpressure ratios and to give greater operatingflexibility.

17. Although a twin-spool compressor (fig. 3-7) canbe used for a pure jet engine, it is most suitable forthe by-pass type of engine where the front or lowpressure compressor is designed to handle a largerairflow than the high pressure compressor. Only apercentage of the air from the low pressurecompressor passes into the high pressurecompressor; the remainder of the air, the by-passflow, is ducted around the high pressure compressor.Both flows mix in the exhaust system before passingto the propelling nozzle (Part 6). This arrangementmatches the velocity of the jet nearer to the optimumrequirements of the aircraft and results in higher



Fig. 3-8 Typical triple spool compressor.

propulsive efficiency, hence lower fuel consumption.For this reason the pure jet engine where all theairflow passes through the full compression cycle isnow obsolete for all but the highest speed aircraft.

18. With the high by-pass ratio turbo-fan this trendis taken a stage further. The intake air undergoesonly one stage of compression in the fan beforebeing split between the core or gas generator systemand the by-pass duct in the ratio of approximatelyone to five (fig. 3-8). This results in the optimumarrangement for passenger and/or transport aircraftflying at just below the speed of sound. The fan maybe coupled to the front of a number of corecompression stages (two shaft engine) or a separateshaft driven by its own turbine (three shaft engine).

Principles of operation19. During operation the rotor is turned at highspeed by the turbine so that air is continuouslyinduced into the compressor, which is thenaccelerated by the rotating blades and sweptrearwards onto the adjacent row of stator vanes. Thepressure rise results from the energy imparted to theair in the rotor which increases the air velocity. Theair is then decelerated (diffused) in the following

stator passage and the kinetic energy translated intopressure. Stator vanes also serve to correct thedeflection given to the air by the rotor blades and topresent the air at the correct angle to the next stageof rotor blades. The last row of stator vanes usuallyact as air straighteners to remove swirl from the airprior to entry into the combustion system at areasonably uniform axial velocity. Changes inpressure and velocity that occur in the airflowthrough the compressor are shown diagrammaticallyin fig. 3-9. The changes are accompanied by aprogressive increase in air temperature as thepressure increases.

20. Across each stage the ratio of total pressures ofoutgoing air and inlet air is quite small, beingbetween 1:1 and 1:2. The reason for the smallpressure increase through each stage is that the rateof diffusion and the deflection angle of the .bladesmust be limited if losses due to air breakaway at theblades and subsequent blade stall are to be avoided.Although the pressure ratio of each stage is small,every stage increases the exit pressure of the stagethat precedes it. So whilst this first stage of acompressor may only increase the pressure by 3 to4 lb. per sq. in., at the rear of a thirty to onecompression system the stage pressure rise can beup to 80 lb, per sq. in, The ability to design multi-stage axial compressors with controlled air velocitiesand straight through flow, minimizes losses andresults in a high efficiency and hence low fuelconsumption. This gives it a further advantage overthe centrifugal compressor where these conditionsare fundamentally not so easily achieved.

21. The more the pressure ratio of a compressor isincreased the more difficult it becomes to ensure thatit will operate efficiently over the full speed range.This is because the requirement for the ratio of inletarea to exit area, at the high speed case, results inan inlet area that becomes progressively too largerelative to the exit area as the compressor speed andhence pressure ratio is reduced. The axial velocity ofthe inlet air in the front stages thus becomes lowrelative to the blade speed, this changes theincidence of the air onto the blades and a conditionis reached where the flow separates and thecompressor flow breaks down. Where high pressureratios are required from a single compressor thisproblem can be overcome by introducing variablestator vanes in the front stages of the system. Thiscorrects the incidence of air onto the rotor blades toangles which they can tolerate. An alternative is theincorporation of interstage bleeds, where aproportion of air after entering the compressor is



Fig. 3-9 Pressure and velocity changesthrough an axial compressor.

removed at an intermediate stage and .dumped intothe bypass flow. While this method corrects the axialvelocity through the preceding stages, energy iswasted and incorporation of variable stators ispreferred.

22. The fan of the high by-pass ratio turbo-fan is anexample of an axial compressor which has beenoptimized to meet the specific requirements of thiscycle. While similar in principle to the corecompressor stage, the proportions of design aresuch that the inner gas path is similar to that of thecore compressor that follows it, while the tip diameteris considerably larger. The mass flow passed by thefan is typically six times that required by the core, theremaining five sixths by-pass the core and isexpanded through its own coaxial nozzle, or may bemixed with the flow at exit from the core in a commonnozzle. To optimize the cycle the by-pass flow has tobe raised to a pressure of approximately 1.6 timesthe inlet pressure. This is achieved in the fan byutilizing very high tip speeds (1500 ft. per sec.) andairflow such that the by-pass section of the bladesoperate with a supersonic inlet air velocity of up toMach 1.5 at the tip. The pressure that results isgraded from a high value at the tip where relativevelocities are highest to the more normal values of1.3 to 1.4 at the inner radius which supercharges thecore where aerodynamic design is more akin to thatof a conventional compressor stage. The capabilityof this type of compressor stage achieves the cyclerequirement of high flow per unit of frontal area, highefficiency and high pressure ratio in a single rotatingblade row without inlet guide vanes within anacceptable engine diameter. Thus keeping weightand mechanical complexity at an acceptable level.

Construction23. The construction of the compressor centresaround the rotor assembly and casings. The rotorshaft is supported in ball and roller bearings andcoupled to the turbine shaft in a manner that allowsfor any slight variation of alignment. The cylindricalcasing assembly may consist of a number ofcylindrical casings with a bolted axial joint betweeneach stage or the casing may be in two halves with abolted centre line joint. One or other of these con-struction methods is required in order that the casingcan be assembled around the rotor.

Rotors24. In compressor designs (fig. 3-10) the rotationalspeed is such that a disc is required to support thecentrifugal blade load. Where a number of discs arefitted onto one shaft they may be coupled andsecured together by a mechanical fixing but

generally the discs are assembled and weldedtogether, close to their periphery, thus forming anintegral drum.

25. Typical methods of securing rotor blades to thedisc are shown in fig. 3-11, fixing may be circumfer-ential or axial to suit special requirements of thestage. In general the aim is to design a securingfeature that imparts the lightest possible load on thesupporting disc thus minimizing disc weight. Whilstmost compressor designs have separate blades formanufacturing and maintainability requirements, itbecomes more difficult on the smallest engines todesign a practical fixing. However this may beovercome by producing blades integral with the disc;the so called 'blisk'.



Fig. 3-10 Rotors of drum and discconstruction.

Rotor blades26. The rotor blades are of airfoil section (fig. 3-12)and usually designed to give a pressure gradientalong their length to ensure that the air maintains areasonably uniform axial velocity. The higherpressure towards the tip balances out the centrifugalaction of the rotor on the airstream. To obtain theseconditions, it is necessary to 'twist' the blade fromroot to tip to give the correct angle of incidence ateach point. Air flowing through a compressor createstwo boundary layers of slow to stagnant air on theinner and outer walls. In order to compensate for theslow air in the boundary layer a localized increase inblade camber both at the blade tip and root has been

introduced. The blade extremities appear as ifformed by bending over each corner, hence the term'end-bend'.

Stator vanes27. The stator vanes are again of airfoil section andare secured into the compressor casing or into statorvane retaining rings, which are themselves securedto the casing (fig. 3-13). The vanes are oftenassembled in segments in the front stages and maybe shrouded at their inner ends to minimize thevibrational effect of flow variations on the longervanes. It is also necessary to lock the stator vanes insuch a manner that they will not rotate around thecasing.



Fig. 3-11 Methods of securing blades to disc.


28. Each stage of a multi-stage compressorpossesses certain airflow characteristics that aredissimilar from those of its neighbour; thus to designa workable and efficient compressor, the characteris-tics of each stage must be carefully matched. This isa relatively simple process to implement for one setof conditions (design mass flow, pressure ratio androtational speed), but is much more difficult whenreasonable matching is to be retained with thecompressor operating over a wide range ofconditions such as an aircraft engine encounters.

29. If the operating conditions imposed upon thecompressor blade departs too far from the designintention, breakdown of airflow and/or aerodynami-cally induced vibration will occur. These phenomenamay take one of two forms; the blades may stallbecause the angle of incidence of the air relative tothe blade is too high (positive incidence stall) or toolow (negative incidence stall). The former is a frontstage problem at low speeds and the latter usuallyaffects the rear stages at high speed, either can leadto blade vibration which can induce rapid destruction.If the engine demands a pressure rise from thecompressor, which is higher than the blading cansustain, 'surge' occurs. In this case there is an instan-taneous breakdown of flow through the machine andthe high pressure air in the combustion system isexpelled forward through the compressor with a loud'bang' and a resultant loss of engine thrust.



Fig. 3-12 A typical rotor blade showingtwisted contour.

Fig. 3-13 Methods of securing vanes to compressor casing.

Compressors are designed with adequate margin toensure that this area of instability (fig. 3-14) isavoided.


30. Where high pressure ratios on a single shaft arerequired it becomes necessary to introduce airflowcontrol into the compressor design. This may take

the form of variable inlet guide vanes for the firststage plus a number of stages incorporating variablestator vanes for the succeeding stages as the shaftpressure ratio is increased (fig. 3-15). As thecompressor speed is reduced from its design valuethese static vanes are progressively closed in orderto maintain an acceptable air angle value onto thefollowing rotor blades. Additionally interstage bleedmay be provided but its use in design is now usuallylimited to the provision of extra margin while theengine is being accelerated, because use at steadyoperating conditions is inefficient and wasteful offuel. Three types of air bleed systems are illustratedas follows: fig. 3-16 hydraulic, fig. 3-17 pneumaticand fig. 3-18 electronic.


31. Materials are chosen to achieve the most costeffective design for the components in question, inpractice for aero engine design this need is usuallybest satisfied by the lightest design that technologyallows for the given loads and temperaturesprevailing.



Fig. 3-14 Limits of stable airflow.

Fig. 3-15 Typical variable stator vanes.



Fig. 3-16 A hydraulically operated bleed valve and inlet guide vane airflow control system.

32. For casing designs the need is for a light butrigid construction enabling blade tip clearances to beaccurately maintained ensuring the highest possibleefficiency. These needs are achieved by usingaluminium at the front of the compression systemfollowed by .alloy steel as compression temperatureincreases. Whilst for the final stages of thecompression system, where temperature require-ments possibly exceed the capability of the beststeel, nickel based alloys may be required. The useof titanium in .preference to aluminium and steel isnow more common; particularly in military engineswhere its high rigidity to density ratio can result insignificant weight reduction. With the development ofnew manufacturing methods component costs cannow be maintained at a more acceptable level inspite of high initial material costs.

33. Stator vanes are normally produced from steelor nickel based alloys, a prime requirement being ahigh fatigue strength when "notched" by ingestion

damage. Earlier designs specified aluminium alloysbut because of its inferior ability to withstand damageits use has declined. Titanium may be used for statorvanes in the low pressure area but is unsuitable forthe smaller stator vanes further rearwards in thecompression system because of the higherpressures and temperatures encountered. Anyexcessive rub which may occur between rotating andstatic components as a result of other mechanicalfailures, can generate sufficient heat from friction toignite the titanium. This in turn can lead to expensiverepair costs and a possible airworthiness hazard.

34. In the design of rotor discs, drums and blades,centrifugal forces dominate and the requirement isfor metal with the highest ratio of strength to density.This results in the lightest possible rotor assemblywhich in turn reduces the forces on the enginestructure enabling a further reduction in weight to beobtained. For this reason, titanium even with its highinitial cost is the preferred material and has replaced



Fig. 3-17 A pneumatically operated bleed valve system.

the steel alloys that were favoured in earlier designs.As higher temperature titanium alloys are developedand produced they are progressively displacing thenickel alloys for the disc and blades at the rear of thesystem.

35. The high by-pass ratio fan blade (fig. 3-19) onlybecame a design possibility with the availability oftitanium, conventional designs being machined fromsolid forgings. A low weight fan blade is necessarybecause the front structure of the engine must beable to withstand the large out of balance forces thatwould result from a fan blade failure. To achieve asufficiently light solid fan blade, even with titanium,requires a short axial length (or chord). However,with this design, the special feature of a mid-spansupport ('snubber' or 'clapper') is required to preventaerodynamic instability. This design concept has thedisadvantage of the snubber being situated in thesupersonic flow where pressure losses are greatest,resulting in inefficiency and a reduction in airflow.This disadvantage has been overcome with theintroduction of the Rolls-Royce designed wide chordfan blade; stability is provided by the increased chordof the blade thus avoiding the need for snubbers.The weight is maintained at a low level by fabricating



Fig. 3-18 An electronically operated bleed valve system.

Fig. 3-19 Typical types of fan blades.

the blade from skins of titanium incorporating ahoneycomb core.

36. Centrifugal impeller material requirements aresimilar to those for the axial compressor rotors.Titanium is thus normally specified though aluminiummay still be employed on the largest low pressureratio designs where robust sections give adequateingestion capability and temperatures are acceptablylow.


37. The balancing of a compressor rotor or impelleris an extremely important operation in itsmanufacture. In view of the high rotational speedsand the mass of materials any unbalance wouldaffect the rotating assembly bearings and engineoperation. Balancing on these parts is effected on aspecial balancing machine, the principles of whichare briefly described in Part 25.



Rolls-Royce RB211 Trent

Rolls-Royce RB41 Nene

On 17 March 1944 Rolls-Royce commencedwork on the RB40 as the result of aGovernment request for a turbo-jet of 4200 lbthrust. After discussions with Supermarine,the airframe designers, the engine was scaleddown to produce 3400 lb. The resulting Nenewas eventually rated at 5000 lb and poweredthe Hawker Sea Hawk and SupermarineAttacker.

4: Combustion chambers

Contents Page

Introduction 35Combustion process 36Fuel supply 38Types of combustion chamber 38

Multiple combustion chamber Tubo-annular combustion chamber Annular combustion chamber

Combustion chamberperformance 41

Combustion intensity Combustion efficiency Combustion stability Emissions

Materials 43


1. The combustion chamber (fig. 4-1) has thedifficult task of burning large quantities of fuel,supplied through the fuel spray nozzles (Part 10),with extensive volumes of air, supplied by thecompressor (Part 3), and releasing the heat in sucha manner that the air is expanded and accelerated togive a smooth stream of uniformly heated gas at allconditions required by the turbine (Part 5). This taskmust be accomplished with the minimum loss inpressure and with the maximum heat release for thelimited space available.

2. The amount of fuel added to the air will dependupon the temperature rise required. However, themaximum temperature is limited to within the rangeof 850 to 1700 deg. C. by the materials from which

the turbine blades and nozzles are made. The air hasalready been heated to between 200 and 550 deg. the work done during compression, giving atemperature rise requirement of 650 to 1150 deg. C.from the combustion process. Since the gastemperature required at the turbine varies withengine thrust, and in the case of the turbo-propellerengine upon the power required, the combustionchamber must also be capable of maintaining stableand efficient combustion over a wide range of engineoperating conditions.

3. Efficient combustion has become increasinglyimportant because of the rapid rise in commercialaircraft traffic and the consequent increase inatmospheric pollution, which is seen by the generalpublic as exhaust smoke.



4. Air from the engine compressor enters thecombustion chamber at a velocity up to 500 feet persecond, but because at this velocity the air speed isfar too high for combustion, the first thing that thechamber must do is to diffuse it, i.e. decelerate it andraise its static pressure. Since the speed of burningkerosine at normal mixture ratios is only a few feetper second, any fuel lit even in the diffused airstream, which now has a velocity of about 80 feet persecond, would be blown away. A region of low axialvelocity has therefore to be created in the chamber,so that the flame will remain alight throughout therange of engine operating conditions.

5. In normal operation, the overall air/fuel ratio of acombustion chamber can vary between 45:1 and130:1, However, kerosine will only burn efficiently at,or close to, a ratio of 15:1, so the fuel must be burnedwith only part of the air entering the chamber, in whatis called a primary combustion zone. This is achievedby means of a flame tube (combustion liner) that has

various devices for metering the airflow distributionalong the chamber.

6. Approximately 20 per cent of the air mass flow istaken in by the snout or entry section (fig. 4-2).Immediately downstream of the snout are swirl vanesand a perforated flare, through which air passes intothe primary combustion zone. The swirling airinduces a flow upstream of the centre of the flametube and promotes the desired recirculation. The airnot picked up by the snout flows into the annularspace between the flame tube and the air casing.

7. Through the wall of the flame tube body, adjacentto the combustion zone, are a selected number ofsecondary holes through which a further 20 per centof the main flow of air passes into the primary zone.The air from the swirl vanes and that from thesecondary air holes interacts and creates a region oflow velocity recirculation. This takes the form of atoroidal vortex, similar to a smoke ring, which has theeffect of stabilizing and anchoring the flame (fig, 4-3).The recirculating gases hasten the burning of freshly

Combustion chambers


Fig. 4-1 An early combustion chamber.

injected fuel droplets by rapidly bringing them toignition temperature.

8. It is arranged that the conical fuel spray from thenozzle intersects the recirculation vortex at its centre.This action, together with the general turbulence inthe primary zone, greatly assists in breaking up thefuel and mixing it with the incoming air.

9. The temperature of the gases released bycombustion is about 1,800 to 2,000 deg. C., which isfar too hot for entry to the nozzle guide vanes of theturbine. The air not used for combustion, whichamounts to about 60 per cent of the total airflow, istherefore introduced progressively into the flametube. Approximately a third of this is used to lower thegas temperature in the dilution zone before it enters

the turbine and the remainder is used for cooling thewalls of the flame tube. This is achieved by a film ofcooling air flowing along the inside surface of theflame tube wall, insulating it from the hot combustiongases (fig. 4-4). A recent development allows coolingair to enter a network of passages within the flametube wall before exiting to form an insulating film ofair, this can reduce the required wall cooling airflowby up to 50 per cent. Combustion should becompleted before the dilution air enters the flametube, otherwise the incoming air will cool the flameand incomplete combustion will result.

10. An electric spark from an igniter plug (Part 11)initiates combustion and the flame is then self-sustained.

Combustion chambers


Fig. 4-2 Apportioning the airflow.

Fig. 4-3 Flame stabilizing and general airflow pattern.

11. The design of a combustion chamber and themethod of adding the fuel may vary considerably, butthe airflow distribution used to effect and maintaincombustion is always very similar to that described.


12. Fuel is supplied to the airstream by one of twodistinct methods. The most common is the injectionof a fine atomized spray into the recirculatingairstream through spray nozzles (Part 10). Thesecond method is based on the pre-vaporization ofthe fuel before it enters the combustion zone.

13. In the vaporizing method (fig.4-5) the fuel issprayed from feed tubes into vaporizing tubes whichare positioned inside the flame tube. These tubesturn the fuel through 180 degrees and, as they areheated by combustion, the fuel vaporizes beforepassing into the flame tube. The primary airflowpasses down the vaporizing tubes with the fuel andalso through holes in the flame tube entry sectionwhich provide 'fans' of air to sweep the flamerearwards. Cooling and dilution air is metered into

the flame tube in a manner similar to the atomizerflame tube.


14. There are three main types of combustionchamber in use for gas turbine engines. These arethe multiple chamber, the tubo-annular chamber andthe annular chamber.

Multiple combustion chamber15. This type of combustion chamber is used oncentrifugal compressor engines and the earlier typesof axial flow compressor engines. It is a directdevelopment of the early type of Whittle combustionchamber. The major difference is that the Whittlechamber had a reverse flow as illustrated in fig. 4-6but, as this created a considerable pressure loss, thestraight-through multiple chamber was developed byJoseph Lucas Limited.

16. The chambers are disposed around the engine(fig. 4-7) and compressor delivery air is directed byducts to pass into the individual chambers. Each

Combustion chambers


Fig. 4-4 Flame tube cooling methods.

chamber has an inner flame tube around which thereis an air casing. The air passes through the flametube snout and also between the tube and the outercasing as already described in para. 6.

17. The separate flame tubes are all interconnect-ed. This allows each tube to operate at the samepressure and also allows combustion to propagatearound the flame tubes during engine starting.

Combustion chambers


Fig. 4-5 A vaporizer combustion chamber.

Fig. 4-6 An early Whittle combustion chamber.

Tubo-annular combustion chamber18. The tubo-annular combustion chamber bridgesthe evolutionary gap between the multiple andannular types. A number of flame tubes are fittedinside a common air casing (fig. 4-8). The airflow issimilar to that already described. This arrangementcombines the ease of overhaul and testing of themultiple system with the compactness of the annularsystem.

Annular combustion chamber19. This type of combustion chamber consists of asingle flame tube, completely annular in form, whichis contained in an inner and outer casing (fig. 4-9).The airflow through the flame tube is similar to thatalready described, the chamber being open at thefront to the compressor and at the rear to the turbinenozzles.

20. The main advantage of the annular chamber isthat, for the same power output, the length of thechamber is only 75 per cent of that of a tubo-annularsystem of the same diameter, resulting in consider-able saving of weight and production cost. Anotheradvantage is the elimination of combustionpropagation problems from chamber to chamber.

21. In comparison with a tubo-annular combustionsystem, the wall area of a comparable annularchamber is much less; consequently the amount ofcooling air required to prevent the burning of theflame tube wall is less, by approximately 15 per cent,This reduction in cooling air raises the combustionefficiency (para. 27) to virtually eliminate unburntfuel, and oxidizes the carbon monoxide to non-toxiccarbon dioxide, thus reducing air pollution.

22. The introduction of the air spray type fuel spraynozzle (Part 10) to this type of combustion chamber

Combustion chambers


Fig. 4-7 Multiple combustion chambers.

also greatly improves the preparation of fuel forcombustion by aerating the over-rich pockets of fuelvapours close to the spray nozzle; this results in alarge reduction in initial carbon formation.


23. A combustion chamber must be capable ofallowing fuel to burn efficiently over a wide range ofoperating conditions without incurring a largepressure loss. In addition, if flame extinction occurs,then it must be possible to relight. In performingthese functions, the flame tube and spray nozzleatomizer components must be mechanically reliable.

24. The gas turbine engine operates on a constantpressure cycle, therefore any loss of pressure duringthe process of combustion must be kept to aminimum. In providing adequate turbulence andmixing, a total pressure loss varying from about 3 to8 per cent of the air pressure at entry to the chamberis incurred.

Combustion intensity25. The heat released by a combustion chamber orany other heat generating unit is dependent on thevolume of the combustion area. Thus, to obtain therequired high power output, a comparatively small

Combustion chambers


Fig. 4-8 Tubo-annular combustion chamber.

and compact gas turbine combustion chamber mustrelease heat at exceptionally high rates.

26. For example, at take-off conditions a Rolls-Royce RB211-524 engine will consume 20,635 lb. offuel per hour. The fuel has a calorific value of approx-imately 18,550 British thermal units per lb., thereforethe combustion chamber releases nearly 106,300British thermal units per second. Expressed in

another way this is an expenditure of potential heatat a rate equivalent to approximately 150,000 horse-power.

Combustion efficiency27. The combustion efficiency of most gas turbineengines at sea-level take-off conditions is almost 100per cent, reducing to 98 per cent at altitude cruiseconditions, as shown in fig. 4-10.

Combustion chambers


Fig. 4-9 Annular combustion chamber.

Combustion stability28. Combustion stability means smooth burningand the ability of the flame to remain alight over awide operating range.

29. For any particular type of combustion chamberthere is both a rich and weak limit to the air/fuel ratio,beyond which the flame is extinguished. Anextinction is most likely to occur in flight during aglide or dive with the engine idling, when there is ahigh airflow and only a small fuel flow, i.e. a veryweak mixture strength.

30. The range of air/fuel ratio between the rich andweak limits is reduced with an increase of air velocity,and if the air mass flow is increased beyond a certainvalue, flame extinction occurs. A typical stability loopis illustrated in fig. 4-11. The operating range definedby the stability loop must obviously cover the air/fuelratios and mass flow of the combustion chamber.

31. The ignition process has weak and rich limitssimilar to those shown for stability in fig. 4-11. Theignition loop, however, lies within the stability loopsince it is more difficult to establish combustion under'cold' conditions than to maintain normal burning.

Emissions32. The unwanted pollutants which are found in theexhaust gases are created within the combustionchamber. There are four main pollutants which arelegislatively controlled; unburnt hydrocarbons(unburnt fuel), smoke (carbon particles), carbonmonoxide and oxides of nitrogen. The principalconditions which affect the formation of pollutants arepressure, temperature and time.

33. In the fuel rich regions of the primary zone, thehydrocarbons are converted into carbon monoxideand smoke, Fresh dilution air can be used to oxidizethe carbon monoxide and smoke into non-toxiccarbon dioxide within the dilution zone. Unburnthydrocarbons can also be reduced in this zone bycontinuing the combustion process to ensurecomplete combustion.

34. Oxides of nitrogen are formed under the sameconditions as those required for the suppression ofthe other pollutants, Therefore it is desirable to coolthe flame as quickly as possible and to reduce thetime available for combustion. This conflict ofconditions requires a compromise to be made, butcontinuing improvements in combustor design andperformance has led to a substantially 'cleaner'combustion process.


35. The containing walls and internal parts of thecombustion chamber must be capable of resistingthe very high gas temperature in the primary zone. Inpractice, this is achieved by using the best heat-resisting materials available, the use of high heatresistant coatings and by cooling the inner wall of theflame tube as an insulation from the flame.

36. The combustion chamber must also withstandcorrosion due to the products of the combustion,creep failure due to temperature gradients andfatigue due to vibrational stresses.

Combustion chambers


Fig. 4-10 Combustion efficiency and air/fuelratio.

Fig. 4-11 Combustion stability limits.

Rolls-Royce Turbomeca Adour Mk102

Rolls-Royce RB37 Derwent V

Work commenced in January 1945 on a 0.855scale Nene, reduced to fit the engine nacelleof a Gloster Meteor. Known as the Derwent Vthe engine passed a 100 hr test at 2600 lbthrust in June 1945 and in September wentinto production with a service rating of 3500lb. Two world speed records were set byMeteor IV's powered by special Derwent V'sin November 1945 and September 1946.

5: Turbines

Contents Page

Introduction 45 Energy transfer from gas flowto turbine 49Construction 51

Nozzle guide vanes Turbine discs Turbine blades Contra-rotating turbines Dual alloy discs

Compressor-turbine matching 53Materials 53

Nozzle guide vanes Turbine discs Turbine blades

Balancing 57


1. The turbine has the task of providing the power todrive the compressor and accessories and, in thecase of engines which do not make use solely of a jetfor propulsion, of providing shaft power for apropeller or rotor. It does this by extracting energyfrom the hot gases released from the combustionsystem and expanding them to a lower pressure and

temperature. High stresses are involved in thisprocess, and for efficient operation, the turbine bladetips may rotate at speeds over 1,500 feet per second,The continuous flow of gas to which the turbine isexposed may have an entry temperature between850 and 1,700 deg. C. and may reach a velocity ofover 2,500 feet per second in parts of the turbine.

2. To produce the driving torque, the turbine mayconsist of several stages each employing one row ofstationary nozzle guide vanes and one row of movingblades (fig. 5-1). The number of stages dependsupon the relationship between the power required


from the gas flow, the rotational speed at which itmust be produced and the diameter of turbinepermitted.

3. The number of shafts, and therefore turbines,varies with the type of engine; high compression ratioengines usually have two shafts, driving high and lowpressure compressors (fig, 5-2). On high by-pass

ratio fan engines that feature an intermediatepressure system, another turbine may be interposedbetween the high and low pressure turbines, thusforming a triple-spool system (fig, 5-3). On someengines, driving torque is derived from a free-powerturbine (fig. 5-4). This method allows the turbine torun at its optimum speed because it is mechanicallyindependent of other turbine and compressor shafts.


Fig. 5-1 A triple-stage turbine with single shaft system.

4. The mean blade speed of a turbine has consid-erable effect on the maximum efficiency possible fora given stage output. For a given output the gasvelocities, deflections, and hence losses, arereduced in proportion to the square of higher meanblade speeds. Stress in the turbine disc increases asthe square of the speed, therefore to maintain thesame stress level at higher speed the sectionalthickness, hence the weight, must be increased dis-proportionately. For this reason, the final design is acompromise between efficiency and weight. Engines

operating at higher turbine inlet temperatures arethermally more efficient and have an improved powerto weight ratio. By-pass engines have a betterpropulsive efficiency and thus can have a smallerturbine for a given thrust.

5. The design of the nozzle guide vane and turbineblade passages is based broadly on aerodynamicconsiderations, and to obtain optimum efficiency,compatible with compressor and combustion design,the nozzle guide vanes and turbine blades are of a


Fig. 5-2 A twin turbine and shaft arrangement.


Fig. 5-3 A triple turbine and shaft arrangement.

basic aerofoil shape. There are three types ofturbine; impulse, reaction and a combination of thetwo known as impulse-reaction. In the impulse typethe total pressure drop across each stage occurs inthe fixed nozzle guide vanes which, because of theirconvergent shape, increase the gas velocity whilstreducing the pressure. The gas is directed onto theturbine blades which experience an impulse forcecaused by the impact of the gas on the blades. In thereaction type the fixed nozzle guide vanes aredesigned to alter the gas flow direction withoutchanging the pressure. The converging bladepassages experience a reaction force resulting fromthe expansion and acceleration of the gas. Normallygas turbine engines do not use pure impulse or pure

reaction turbine blades but the impulse-reactioncombination (fig. 5-5). The proportion of eachprinciple incorporated in the design of a turbine islargely dependent on the type of engine in which theturbine is to operate, but in general it is about 50 percent impulse and 50 per cent reaction. Impulse-typeturbines are used for cartridge and air starters (Part11).


6. From the description contained in para. 1, it willbe seen that the turbine depends for its operation onthe transfer of energy between the combustion


Fig. 5-4 A typical free power turbine.

gases and the turbine. This transfer is never 100 percent because of thermodynamic and mechanicallosses, (para. 11).

7.when the gas is expanded by the combustionprocess (Part 4), it forces its way into the dischargenozzles of the turbine where, because of theirconvergent shape, it is accelerated to about thespeed of sound which, at the gas temperature, isabout 2,500 feet per second. At the same time thegas flow is given a 'spin' or 'whirl' in the direction ofrotation of the turbine blades by the nozzle guidevanes. On impact with the blades and during thesubsequent reaction through the blades, energy isabsorbed, causing the turbine to rotate at high speedand so provide the power for driving the turbine shaftand compressor.

8. The torque or turning power applied to theturbine is governed by the rate of gas flow and theenergy change of the gas between the inlet and theoutlet of the turbine blades, The design of the turbineis such that the whirl will be removed from the gasstream so that the flow at exit from the turbine will be

substantially 'straightened out' to give an axial flowinto the exhaust system (Part 6). Excessive residualwhirl reduces the efficiency of the exhaust systemand also tends to produce jet pipe vibration whichhas a detrimental effect on the exhaust conesupports and struts.

9. It will be seen that the nozzle guide vanes andblades of the turbine are 'twisted', the blades havinga stagger angle that is greater at the tip than at theroot (fig. 5-6). The reason for the twist is to make thegas flow from the combustion system do equal workat all positions along the length of the blade and toensure that the flow enters the exhaust system witha uniform axial velocity. This results in certainchanges in velocity, pressure and temperatureoccurring through the turbine, as shown diagram-matically in fig. 5-7.

10. The 'degree of reaction' varies from root to tip,being least at the root and highest at the tip, with themean section having the chosen value of about 50per cent.


Fig. 5-5 Comparison between a pure Impulse turbine and an impulse/reaction turbine.

11. The losses which prevent the turbine from being100 per cent efficient are due to a number ofreasons. A typical uncooled three-stage turbinewould suffer a 3.5 per cent loss because ofaerodynamic losses in the turbine blades. A further4.5 per cent loss would be incurred by aerodynamiclosses in the nozzle guide vanes, gas leakage overthe turbine blade tips and exhaust system losses;these losses are of approximately equal proportions.The total losses result in an overall efficiency ofapproximately 92 per cent.


12. The basic components of the turbine are thecombustion discharge nozzles, the nozzle guidevanes, the turbine discs and the turbine blades. Therotating assembly is carried on bearings mounted inthe turbine casing and the turbine shaft may becommon to the compressor shaft or connected to itby a self-aligning coupling.

Nozzle guide vanes13. The nozzle guide vanes are of an aerofoil shapewith the passage between adjacent vanes forming aconvergent duct. The vanes are located (fig. 5-8) inthe turbine casing in a manner that allows forexpansion.


Fig. 5-6 A typical turbine blade showingtwisted contour.

Fig. 5-7 Gas flow pattern through nozzle and blade.

14. The nozzle guide vanes are usually of hollowform and may be cooled by passing compressordelivery air through them to reduce the effects of highthermal stresses and gas loads. For details of turbinecooling, reference should be made to Part 9.

15. Turbine discs are usually manufactured from amachined forging with an integral shaft or with aflange onto which the shaft may be bolted. The discalso has, around its perimeter, provision for theattachment of the turbine blades.

16. To limit the effect of heat conduction from theturbine blades to the disc a flow of cooling air ispassed across both sides of each disc (Part 9).

Turbine blades17. The turbine blades are of an aerofoil shape,designed to provide passages between adjacentblades that give a steady acceleration of the flow upto the 'throat', where the area is smallest and the

velocity reaches that required at exit to produce therequired degree of reaction (para. 5).

18. The actual area of each blade cross-section isfixed by the permitted stress in the material used andby the size of any holes which may be required forcooling purposes (Part 9). High efficiency demandsthin trailing edges to the sections, but a compromisehas to be made so as to prevent the blades crackingdue to the temperature changes during engineoperation.

19. The method of attaching the blades to theturbine disc is of considerable importance, since thestress in the disc around the fixing or in the bladeroot has an important bearing on the limiting rimspeed. The blades on the early Whittle engine wereattached by the de Laval bulb root fixing, but thisdesign was soon superseded by the 'fir-tree' fixingthat is now used in the majority of gas turbineengines. This type of fixing involves very accuratemachining to ensure that the loading is shared by all


Fig. 5-8 Typical nozzle guide vanes showing their shape and location.

the serrations. The blade is free in the serrationswhen the turbine is stationary and is stiffened in theroot by centrifugal loading when the turbine isrotating. Various methods of blade attachment areshown in fig. 5-9; however, the B.M.W. hollow bladeand the de Laval bulb root types are not nowgenerally used on gas turbine engines.

20. A gap exists between the blade tips and casing,which varies in size due to the different rates ofexpansion and contraction. To reduce the loss ofefficiency through gas leakage across the blade tips,a shroud is often fitted as shown in fig. 5-1. This ismade up by a small segment at the tip of each bladewhich forms a peripheral ring around the blade tips.An abradable lining in the casing may also be usedto reduce gas leakage as discussed in Part 9. ActiveClearance Control (A.C.C.) is a more effectivemethod of maintaining minimum tip clearancethroughout the flight cycle. Air from the compressor isused to cool the turbine casing and when used withshroudless turbine blades, enables higher tempera-tures and speeds to be used.

Contra-rotating turbine21. Fig. 5-10 shows a twelve stage contra-rotatingfree power turbine driving a contra-rotating rear fan.This design has only one row of static nozzle guidevanes. The remaining nozzle guide vanes are, ineffect, turbine blades attached to a rotating casingwhich revolves in the opposite direction to a rotatingdrum. Since all but one aerofoil row extracts energyfrom the gas stream, contra-rotating turbines are

capable of operating at much higher stage loadingsthan conventional turbines, making them attractivefor direct drive applications.

Dual alloy discs22. Very high stresses are imposed on the bladeroot fixing of high work rate turbines, which makeconventional methods of blade attachmentimpractical. A dual alloy disc, or 'blisk' as shown infig. 5-11, has a ring of cast turbine blades bonded tothe disc. This type of turbine is suitable for small highpower helicopter engines.


23. The flow characteristics of the turbine must bevery carefully matched with those of the compressorto obtain the maximum efficiency and performance ofthe engine. If, for example, the nozzle guide vanesallowed too low a maximum flow, then a backpressure would build up causing the compressor tosurge (Part 3); too high a flow would cause thecompressor to choke. In either condition a loss ofefficiency would very rapidly occur.


24. Among the obstacles in the way of using higherturbine entry temperatures have always been theeffects of these temperatures on the nozzle guidevanes and turbine blades, The high speed of rotationwhich imparts tensile stress to the turbine disc andblades is also a limiting factor.


Fig. 5-9 Various methods of attaching blades to turbine discs.

Nozzle guide vanes25. Due to their static condition. the nozzle guidevanes do not endure the same rotational stresses asthe turbine blades. Therefore, heat resistance is theproperty most required. Nickel alloys are used,although cooling is required to prevent melting.Ceramic coatings can enhance the heat resistingproperties and, for the same set of conditions,reduce the amount of cooling air required, thusimproving engine efficiency.

Turbine discs26. A turbine disc has to rotate at high speed in arelatively cool environment and is subjected to largerotational stresses. The limiting factor which affectsthe useful disc life is its resistance to fatiguecracking.


Fig. 5-10 Free power contra-rotating turbine.

Fig. 5-11 Section through a dual alloy disc.


Fig. 5-11 Section through a dual alloy disc.

27. In the past, turbine discs have been made inferritic and austenitic steels but nickel based alloysare currently used. Increasing the alloying elementsin nickel extend the life limits of a disc by increasingfatigue resistance. Alternatively, expensive powdermetallurgy discs, which offer an additional 10% instrength, allow faster rotational speeds to beachieved.

Turbine blades28. A brief mention of some of the points to beconsidered in connection with turbine blade designwill give an idea of the importance of the correctchoice of blade material. The blades, while glowingred-hot, must be strong enough to carry thecentrifugal loads due to rotation at high speed. Asmall turbine blade weighing only two ounces mayexert a load of over two tons at top speed and it mustwithstand the high bending loads applied by the gasto produce the many thousands of turbine horse-power necessary to drive the compressor. Turbineblades must also be resistant to fatigue and thermalshock, so that they will not fail under the influence ofhigh frequency fluctuations in the gas conditions, andthey must also be resistant to corrosion andoxidization. In spite of all these demands, the bladesmust be made in a material that can be accuratelyformed and machined by current manufacturingmethods.

29. From the foregoing, it follows that for aparticular blade material and an acceptable safe lifethere is an associated maximum permissible turbineentry temperature and a corresponding maximumengine power. It is not surprising, therefore, that met-allurgists and designers are constantly searching forbetter turbine blade materials and improved methodsof blade cooling.

30. Over a period of operational time the turbineblades slowly grow in length. This phenomenon isknown as 'creep' and there is a finite useful life limitbefore failure occurs.

31. The early materials used were high temperaturesteel forgings, but these were rapidly replaced bycast nickel base alloys which give better creep andfatigue properties.

32. Close examination of a conventional turbineblade reveals a myriad of crystals that lie in alldirections (equi-axed). Improved service life can beobtained by aligning the crystals to form columnsalong the blade length, produced by a method knownas 'Directional Solidification'. A further advance ofthis technique is to make the blade out of a single


Fig. 5-13 Comparison of turbine blade lifeproperties.

Fig. 5-14 Ceramic turbine blades.

crystal, Examples of these structures are shown infig. 5-12. Each method extends the useful creep lifeof the blade (fig. 5-13) and in the case of the singlecrystal blade, the operating temperature can be sub-stantially increased.

33. A non-metal based turbine blade can be manu-factured from reinforced ceramics. Their initialproduction application is likely to be for small highspeed turbines which have very high turbine entry

temperatures. An example of a ceramic blade isshown in fig. 5-14.


34. The balancing of a turbine is an extremelyimportant operation in its assembly. In view of thehigh rotational speeds and the mass of materials,any unbalance could seriously affect the rotatingassembly bearings and engine operation. Balancingis effected on a special balancing machine, theprinciples of which are briefly described in Part 25.


Rolls-Royce RB50 Trent

Late in 1943 the decision was taken at Rolls-Royce to build a turbo-prop for aircraft speedsof around 400 mph. The resulting engine,known as the RB50 Trent, was basically aDerwent II with a flexible quillshaft toreduction gear and propeller. On 20September 1945 a Gloster Meteor, fitted withtwo Trents, became the world's first turbo-prop powered aircraft to fly.

Rolls-Royce RB211-535E4

6: Exhaust system

Contents Page

Introduction 59Exhaust gas flow 61Construction and materials 63


1. Aero gas turbine engines have an exhaustsystem which passes the turbine discharge gases toatmosphere at a velocity, and in the requireddirection, to provide the resultant thrust. The velocityand pressure of the exhaust gases create the thrustin the turbo-jet engine (para. 5) but in the turbo-propeller engine only a small amount of thrust iscontributed by the exhaust gases, because most ofthe energy has been absorbed by the turbine fordriving the propeller. The design of the exhaustsystem therefore, exerts a considerable influence onthe performance of the engine. The areas of the jetpipe and propelling or outlet nozzle affect the turbineentry temperature, the mass airflow and the velocityand pressure of the exhaust jet.

2. The temperature of the gas entering the exhaustsystem is between 550 and 850 deg. C. according tothe type of engine and with the use of afterburning(Part 16) can be 1,500 deg. C. or higher. Therefore,it is necessary to use materials and a form of con-struction that will resist distortion and cracking, andprevent heat conduction to the aircraft structure.

3. A basic exhaust system is shown in fig. 6-1. Theuse of a thrust reverser (Part 15), noise suppressor(Part 19) and a two position propelling nozzle entailsa more complicated system as shown in fig. 6-2. Thelow by-pass engine may also include a mixer unit(fig. 6-4) to encourage a thorough mixing of the hotand cold gas streams.


Exhaust system


Fig. 6-1 A basic exhaust system.

Fig. 6-2 Exhaust system with thrust reverser, noise suppressor and two position propelling nozzle.


4. Gas from the engine turbine enters the exhaustsystem at velocities from 750 to 1,200 feet persecond, but, because velocities of this order producehigh friction losses, the speed of flow is decreased bydiffusion. This is accomplished by having anincreasing passage area between the exhaust coneand the outer wall as shown in fig. 6-1. The cone alsoprevents the exhaust gases from flowing across therear face of the turbine disc. It is usual to hold thevelocity at the exhaust unit outlet to a Mach numberof about 0.5, i.e. approximately 950 feet per second.Additional losses occur due to the residual whirlvelocity in the gas stream from the turbine. To reducethese losses, the turbine rear struts in the exhaustunit are designed to straighten out the flow before thegases pass into the jet pipe.

5. The exhaust gases pass to atmosphere throughthe propelling nozzle, which is a convergent duct,thus increasing the gas velocity (Part 2). In a turbo-jet engine, the exit velocity of the exhaust gases issubsonic at low thrust conditions only. During mostoperating conditions, the exit velocity reaches thespeed of sound in relation to the exhaust gas

temperature and the propelling nozzle is then said tobe 'choked'; that is, no further increase in velocitycan be obtained unless the temperature is increased.As the upstream total pressure is increased abovethe value at which the propelling nozzle becomes'choked', the static pressure of the gases at exitincreases above atmospheric pressure. Thispressure difference across the propelling nozzlegives what is known as 'pressure thrust' and iseffective over the nozzle exit area. This is additionalthrust to that obtained due to the momentum changeof the gas stream (Part 20).

Exhaust system


Fig. 6-3 Gas flow through a convergent-divergent nozzle.

Fig. 6-4 A low by-pass air mixer unit.

6. With the convergent type of nozzle a wastage ofenergy occurs, since the gases leaving the exit donot expand rapidly enough to immediately achieveoutside air pressure. Depending on the aircraft flightplan, some high pressure ratio engines can withadvantage use a convergent-divergent nozzle torecover some of the wasted energy This nozzleutilizes the pressure energy to obtain a furtherincrease in gas velocity and, consequently, anincrease in thrust.

7. From the illustration (fig. 6-3), it will be seen thatthe convergent section exit now becomes the throat,

with the exit proper now being at the end of the flareddivergent section. When the gas enters theconvergent section of the nozzle, the gas velocityincreases with a corresponding fall in static pressure.The gas velocity at the throat corresponds to thelocal sonic velocity. As the gas leaves the restrictionof the throat and flows into the divergent section, itprogressively increases in velocity towards the exit.The reaction to this further increase in momentum isa pressure force acting on the inner wall of thenozzle. A component of this force acting parallel tothe longitudinal axis of the nozzle produces thefurther increase in thrust.

Exhaust system


Fig. 6-5 High by-pass ratio engine exhaust systems.

8. The propelling nozzle size is extremely importantand must be designed to obtain the correct balanceof pressure, temperature and thrust. With a smallnozzle these values increase, but there is apossibility of the engine surging (Part 3), whereaswith a large nozzle the values obtained are too low,

9. A fixed area propelling nozzle is only efficientover a narrow range of engine operating conditions.To increase this range, a variable area nozzle maybe used. This type of nozzle is usually automaticallycontrolled and is designed to maintain the correctbalance of pressure and temperature at all operatingconditions. In practice, this system is seldom used asthe performance gain is offset by the increase inweight. However, with afterburning a variable areanozzle is necessary and is described in Part 16.

10. The by-pass engine has two gas streams toeject to atmosphere, the cool by-pass airflow and thehot turbine discharge gases.

11. In a low by-pass ratio engine, the two flows arecombined by a mixer unit (fig. 6-4) which allows theby-pass air to flow into the turbine exhaust gas flowin a manner that ensures thorough mixing of the twostreams.

12. In high by-pass ratio engines, the two streamsare usually exhausted separately. The hot and coldnozzles are co-axial and the area of each nozzle isdesigned to obtain maximum efficiency. However, animprovement can be made by combining the two gasflows within a common, or integrated, nozzleassembly. This partially mixes the gas flows prior toejection to atmosphere. An example of both types ofhigh by-pass exhaust system is shown in fig, 6-5.


13. The exhaust system must be capable of with-standing the high gas temperatures and is thereforemanufactured from nickel or titanium. It is alsonecessary to prevent any heat being transferred tothe surrounding aircraft structure. This is achieved bypassing ventilating air around the jet pipe, or bylagging the section of the exhaust system with aninsulating blanket (fig. 6-6). Each blanket has aninner layer of fibrous insulating material contained by

an outer skin of thin stainless steel, which is dimpledto increase its strength. In addition, acousticallyabsorbent materials are sometimes applied to theexhaust system to reduce engine noise (Part 19).

14. When the gas temperature is very high (forexample, when afterburning is employed), thecomplete jet pipe is usually of double-wall construc-tion (Part 16) with an annular space between the twowalls. The hot gases leaving the propelling nozzleinduce, by ejector action, a flow of air through theannular space of the engine nacelle. This flow of aircools the inner wall of the jet pipe and acts as aninsulating blanket by reducing the transfer of heatfrom the inner to the outer wall.

15. The cone and streamline fairings in the exhaustunit are subjected to the pressure of the exhaustgases; therefore, to prevent any distortion, ventholes are provided to obtain a pressure balance.

16. The mixer unit used in low by-pass ratioengines consists of a number of chutes throughwhich the bypass air flows into the exhaust gases. Abonded honeycomb structure is used for theintegrated nozzle assembly of high by-pass ratioengines to give lightweight strength to this largecomponent.

17. Due to the wide variations of temperature towhich the exhaust system is subjected, it must bemounted and have its sections joined together insuch a manner as to allow for expansion andcontraction without distortion or damage.

Exhaust system


Fig. 6-6 An insulating blanket.

Rolls-Royce Gnome

De Havilland H2 Ghost

The Ghost was designed as a larger and morepowerful version of the Goblin. After runningfor the first time on 2 September 1945 theengine was cleared for flight in the outernacelles of an Avro Lancastrian at 4000 lbthrust. The Ghost later went into production at5000 lb thrust to power the de HavillandComet 1 airliner and Venom fighter.

7: Accessory drives

Contents Page

Introduction 65Gearboxes and drives 65

Internal gearbox Radial driveshaft Direct drive Gear train drive Intermediate gearbox External gearbox Auxiliary gearbox

Construction and materials 69GearsGearbox sealing Materials


1. Accessory units provide the power for aircrafthydraulic, pneumatic and electrical systems inaddition to providing various pumps and controlsystems for efficient engine operation. The high levelof dependence upon these units requires anextremely reliable drive system.

2. The drive for the accessory units is typicallytaken from a rotating engine shaft, via an internalgearbox, to an external gearbox which provides amount for the accessories and distributes theappropriate geared drive to each accessory unit. Astarter may also be fitted to provide an input torqueto the engine. An accessory drive system on a highby-pass engine takes between 400 and 500horsepower from the engine.


Internal gearbox3. The location of the internal gearbox within thecore of an engine is dictated by the difficulties ofbringing a driveshaft radially outwards and the spaceavailable within the engine core.

4. Thermal fatigue and a reduction in engineperformance, due to the radial driveshaft disturbingthe gasflow, create greater problems within theturbine area than the compressor area. For anygiven engine, which incorporates an axial-flowcompressor, the turbine area is smaller than thatcontaining the compressor and therefore makes itphysically easier to mount the gearbox within thecompressor section. Centrifugal compressor enginescan have limited available space, so the internalgearbox may be located within a static nose cone or,in the case of a turbo-propeller engine, behind thepropeller reduction gear as shown in fig. 7-1.

5. On multi-shaft engines, the choice of whichcompressor shaft is used to drive the internalgearbox is primarily dependent upon the ease of


Accessory drives


Fig. 7-1 Mechanical arrangement of accessory drives.

engine starting. This is achieved by rotating thecompressor shaft, usually via an input torque fromthe external gearbox (Part 11). In practice the highpressure system is invariably rotated in order togenerate an airflow through the engine and the highpressure compressor shaft is therefore coupled tothe internal gearbox.

6. To minimize unwanted movement between thecompressor shaft bevel gear and radial driveshaftbevel gear, caused by axial movement of thecompressor shaft, the drive is taken by one of threebasic methods (fig. 7-2). The least number ofcomponents is used when the compressor shaftbevel gear is mounted as close to the compressorshaft location bearing as possible, but a smallamount of movement has to be accommodatedwithin the meshing of the bevel gears. Alternatively,the compressor shaft bevel gear may be mounted ona stub shaft which has its own location bearing. Thestub shaft is splined onto the compressor shaft whichallows axial movement without affecting the bevelgear mesh. A more complex system utilizes an idlergear which meshes with the compressor shaft viastraight spur gears, accommodating the axialmovement, and drives the radial driveshaft via abevel gear arrangement. The latter method waswidely employed on early engines to overcome gearengagement difficulties at high speed.

7. To spread the load of driving accessory units,some engines take a second drive from the slowerrotating low pressure shaft to a second externalgearbox (fig. 7-1). This also has the advantage oflocating the accessory units in two groups, thusovercoming the possibility of limited external spaceon the engine. When this method is used, an attemptis made to group the accessory units specific to theengine onto the high pressure system, since that isthe first shaft to rotate, and the aircraft accessoryunits are driven by the low pressure system. A typicalinternal gearbox showing how both drives are takenis shown in fig. 7-3.

Radial driveshaft8. The purpose of a radial driveshaft is to transmitthe drive from the internal gearbox to an accessoryunit or the external gearbox. It also serves to transmitthe high torque from the starter to rotate the highpressure system for engine starting purposes. Thedriveshaft may be direct drive or via an intermediategearbox (para. 14).

9. To minimize the effect of the driveshaft passingthrough the compressor duct and disrupting theairflow, it is housed within the compressor supportstructure. On by-pass engines, the driveshaft iseither housed in the outlet guide vanes or in a hollowstreamlined radial fairing across the low pressurecompressor duct.

10. To reduce airflow disruption it is desirable tohave the smallest driveshaft diameter as possible.The smaller the diameter, the faster the shaft must

Accessory drives


Fig. 7-2 Mechanical arrangement of internalgearboxes.

rotate to provide the same power. However, thisraises the internal stress and gives greater dynamicproblems which result in vibration. A long radialdriveshaft usually requires a roller bearing situatedhalfway along its length to give smooth running. Thisallows a rotational speed of approximately 25,000r.p.m. to be achieved with a shaft diameter of lessthan 1.5 inch without encountering serious vibrationproblems.

Direct drive11. In some early engines, a radial driveshaft wasused to drive each, or in some instances a pair, ofaccessory units. Although this allowed each

accessory unit to be located in any desirable locationaround the engine and decreased the powertransmitted through individual gears, it necessitateda large internal gearbox. Additionally, numerousradial driveshafts had to be incorporated within thedesign. This led to an excessive amount of timerequired for disassembly and assembly of the enginefor maintenance purposes.

12. In some instances the direct drive method maybe used in conjunction with the external gearboxsystem when it is impractical to take a drive from aparticular area of the engine to the external gearbox.For example, fig. 7-1 shows a turbo-propeller engine

Accessory drives


Fig. 7-3 An internal gearbox.

which requires accessories specific to the propellerreduction drive, but has the external gearbox locatedaway from this area to receive the drive from thecompressor shaft.

Gear train drive13. When space permits, the drive may be taken tothe external gearbox via a gear train (fig. 7-1). Thisinvolves the use of spur gears, sometimes incorpo-rating a centrifugal breather (Part 8). However, it israre to find this type of drive system in current use.

Intermediate gearbox14. Intermediate gearboxes are employed when it isnot possible to directly align the radial driveshaft withthe external gearbox. To overcome this problem anintermediate gearbox is mounted on the highpressure compressor case and re-directs the drive,through bevel gears, to the external gearbox. Anexample of this layout is shown in fig. 7-1.

External gearbox15. The external gearbox contains the drives for theaccessories, the drive from the starter and providesa mounting face for each accessory unit. Provision isalso made for hand turning the engine, via thegearbox, for maintenance purposes. Fig. 7-4 showsthe accessory units that are typically found on anexternal gearbox.

16. The overall layout of an external gearbox isdictated by a number of factors. To reduce dragwhilst the aircraft is flying it is important to present alow frontal area to the airflow. Therefore the gearboxis 'wrapped' around the engine and may look, fromthe front, similar to a banana in shape. Formaintenance purposes the gearbox is generallylocated on the underside of the engine to allowground crew to gain access. However, helicopterinstallation design usually requires the gearbox to belocated on the top of the engine for ease of access.

17. The starter/driven gearshaft (fig. 7-4) roughlydivides the external gearbox into two sections. Onesection provides the drive for the accessories whichrequire low power whilst the other drives the highpower accessories. This allows the small and largegears to be grouped together independently and isan efficient method of distributing the drive for theminimum weight.

18. If any accessory unit fails, and is preventedfrom rotating, it could cause further failure in theexternal gearbox by shearing the teeth of the geartrain. To prevent secondary failure occurring a weaksection is machined into the driveshafts, known as a

'shear-neck', which is designed to fail and thusprotect the other drives. This feature is not includedfor primary engine accessory units, such as the oilpumps, because these units are vital to the runningof the engine and any failure would necessitateimmediate shutdown of the engine.

19. Since the starter provides the highest torquethat the drive system encounters, it is the basis ofdesign. The starter is usually positioned to give theshortest drive line to the engine core. This eliminatesthe necessity of strengthening the entire gear trainwhich would increase the gearbox weight. However,when an auxiliary gearbox is fitted (para, 21) thestarter is moved along the gear train to allow theheavily loaded auxiliary gearbox drive to passthrough the external gearbox. This requires the spurgears between the starter and starter/drivengearshaft to have a larger face width to carry the loadapplied by the starter (fig. 7-5).

20. When a drive is taken from two compressorshafts, as discussed in para. 7, two separategearboxes are required. These are mounted eitherside of the compressor case and are generallyknown as the 'low speed' and 'high speed' externalgearboxes.

Auxiliary gearbox21. An auxiliary gearbox is a convenient method ofproviding additional accessory drives when the con-figuration of an engine and airframe does not allowenough space to mount all of the accessory units ona single external gearbox.

22. A drive is taken from the external gearbox (fig.7-5) to power the auxiliary gearbox which distributesthe appropriate gear ratio drive to the accessories inthe same manner as the external gearbox.


Gears23. The spur gears of the external or auxiliarygearbox gear train (fig. 7-4 and 7-5) are mountedbetween bearings supported by the front and rearcasings which are bolted together. They transmit thedrive to each accessory unit, which is normallybetween 5000 and 6000 r.p.m. for the accessoryunits and approximately 20,000 r.p.m. for thecentrifugal breather,

24. All gear meshes are designed with 'huntingtooth' ratios which ensure that each tooth of a geardoes not engage between the same set of opposingteeth on each revolution. This spreads any wearevenly across all teeth.

Accessory drives


25. Spiral bevel gears are used for the connectionof shafts whose axes are at an angle to one anotherbut in the same plane. The majority of gears within agear train are of the straight spur gear type, thosewith the widest face carry the greatest loads. Forsmoother running, helical gears are used but theresultant end thrust caused by this gear tooth patternmust be catered for within the mounting of the gear.

Gearbox sealing26. Sealing of the accessory drive system isprimarily concerned with preventing oil loss. Theinternal gearbox has labyrinth seals where the staticcasing mates with the rotating compressor shaft. Forsome o! the accessories mounted on the externalgearbox, an air blown pressurized labyrinth seal is

Accessory drives


Fig. 7-4 An external gearbox and accessory units.

employed. This prevents oil from the gearboxentering the accessory unit and also prevents con-tamination of the gearbox, and hence engine, in theevent of an accessory failure. The use of an air blownseal results in a gearbox pressure of about 3 lbs. persq. in. above atmospheric pressure. To supplement alabyrinth seal, an 'oil thrower ring' may be used. Thisinvolves the leakage oil running down the drivingshaft and being flung outwards by a flange on therotating shaft. The oil is then collected and returnedto the gearbox.

Materials27. To reduce weight, the lightest materials possibleare used. The internal gearbox casing is cast fromaluminium but the low environmental temperaturesthat an external gearbox is subjected to allows theuse of magnesium castings which are lighter still.The gears are manufactured from non-corrosionresistant steels for strength and toughness. They arecase hardened to give a very hard wear resistantskin and feature accurately ground teeth for smoothgear meshing.

Accessory drives


Fig. 7-5 An external gearbox with auxiliary gearbox drive.

Rolls-Royce Tay

Bristol Theseus

This engine was conceived in 1940 as a 4000hp turbo-prop but was later scaled down to2000 hp. Named the Theseus the engine wastype tested in December 1946. the world's firstturbo-prop to reach this stage of development.After extensive flight testing in an AvroLincoln, four Theseus engines were installedin a Handley-Page Hermes 5.


1. The lubrication system is required to providelubrication and cooling for all gears, bearings andsplines. It must also be capable of collecting foreignmatter which, if left in a bearing housing or gearbox,can cause rapid failure. Additionally, the oil mustprotect the lubricated components which are manu-factured from non-corrosion resistant materials. Theoil must accomplish these tasks without significantdeterioration.

2. The requirements of a turbo-propeller engine aresomewhat different to any other types of aero gasturbine. This is due to the additional lubrication of theheavily loaded propeller reduction gears and theneed for a high pressure oil supply to operate thepropeller pitch control mechanism.

3. Most gas turbine engines use a self-containedrecirculatory lubrication system in which the oil isdistributed around the engine and returned to the oiltank by pumps. However, some engines use asystem known as the total loss or expendable systemin which the oil is spilled overboard after the enginehas been lubricated.


4. There are two basic recirculatory systems,known as the 'pressure relief valve1 system and the'full flow' system. The major difference between themis in the control of the oil flow to the bearings. In bothsystems the temperature and pressure of the oil arecritical to the correct and safe running of the engine.Provision is therefore made for these parameters tobe indicated in the cockpit.

8: Lubrication

Contents Page

Introduction 73 Lubricating systems 73

Pressure relief valve systemFull flow systemTotal loss (expendable) system

Oil system components 77Lubricating oils 83


Pressure relief valve system5. In the pressure relief valve system the oil flow tothe bearing chambers is controlled by limiting thepressure in the feed line to a given design value. Thisis accomplished by the use of a spring loaded valvewhich allows oil to be directly returned from thepressure pump outlet to the oil tank, or pressurepump inlet, when the design value is exceeded. Thevalve opens at a pressure which corresponds to theidling speed of the engine, thus giving a constantfeed pressure over normal engine operating speeds.However, increasing engine speed causes thebearing chamber pressure to rise sharply. Thisreduces the pressure difference between the bearingchamber and feed jet, thus decreasing the oil flowrate to the bearings as engine speed increases. Toalleviate this problem, some pressure relief valvesystems use the increasing bearing chamberpressure to augment the relief valve spring load, Thismaintains a constant flow rate at the higher enginespeeds by increasing the pressure in the feed line asthe bearing chamber pressure increases.

6 Fig. 8-1 shows the pressure relief valve systemfor a turbo-propeller engine and indicates the basiccomponents that comprise an engine lubricationsystem. The oil pressure pump draws oil from thetank through a strainer which protects the pumpgears from debris which may have entered the tank,Oil is then delivered through a pressure filter to thepressure relief valve which maintains a constant oildelivery pressure to the feed jets in the bearingchambers. Some engines may have an additionalrelief valve (pressure limiting valve) which is fitted atthe oil pressure pump outlet. This valve is set to openat a much higher value than the pressure relief valveto return the oil to the inlet side of the oil pressurepump in the event of the system becoming blocked.A similar valve may also be fitted across the pressurefilter to prevent oil starvation of the bearing chambersshould the filter become partially blocked or the oilhaving a high viscosity under cold starting conditionspreventing sufficient flow through the filter. Provisionis also made to supply oil to the propeller pitchcontrol system, reduction gear and torquemeter



Fig. 8-1 A pressure relief valve type oil system.

system. Scavenge pumps return the oil to the tankvia the oil cooler. On entering the tank, the oil is de-aerated ready for recirculation.

Full flow system7. Although the pressure relief valve systemoperates satisfactorily for engines which have a lowbearing chamber pressure, which does not undulyincrease with engine speed, it becomes anundesirable system for engines which have highchamber pressures. For example, if a bearingchamber has a maximum pressure of 90 lb. per It would require a pressure relief valve setting of130 lb. per sq. in. to produce a pressure drop of 40lb. per sq. in. at the oil feed jet. This results in the

need for large pumps and difficulty in matching therequired oil flow at slower speeds.

8. The full flow system achieves the desired oil flowrates throughout the complete engine speed rangeby dispensing with the pressure relief valve andallowing the pressure pump delivery pressure tosupply directly the oil feed jets. Fig. 8-2 shows anexample of this system which may be found on aturbo-fan engine. The pressure pump size isdetermined by the flow required at maximum enginespeed. The use of this system allows smallerpressure and scavenge pumps to be used since thelarge volume of oil which is spilled by the pressurerelief valve system at maximum engine speed isobviated.



Fig. 8-2 A full flow type oil system.

9. To prevent high oil pressures from damagingfilters or coolers, pressure limiting valves are fitted toby-pass these units. These valves normally onlyoperate under cold starting conditions or in the eventof a blockage. Advance warning of a blocked filtermay be indicated in the cockpit by a differentialpressure switch which senses an increase in thepressure difference between the inlet and outlet ofthe filter.

Total loss (expendable) system10. For engines which run for periods of shortduration, such as booster and vertical lift engines,

the total loss oil system is generally used. Thesystem is simple and incurs low weight penaltiesbecause it requires no oil cooler, scavenge pump orfilters. On some engines oil is delivered in acontinuous flow to the bearings by a plunger-typepump, indirectly driven from the compressor shaft; onothers it is delivered by a piston-type pump operatedby fuel pressure (fig. 8-3). In the latter, the oil supplyis automatically selected by the high pressure fuelshut-off valve (cock) during engine starting and isdelivered as a single shot to the front and rearbearings. On some engines provision is made for a



Fig. 8-3 A total loss (expendable) oil system.

second shot to be delivered to the rear bearing only,after a predetermined period.

11. After lubricating the fuel unit and front bearings,the oil from the front bearing drains into a collectortray and is then ejected into the main gas streamthrough an ejector nozzle. The oil that has passedthrough the rear bearing, drains into a reservoir atthe rear of the bearing where it is retained bycentrifugal force until the engine is shut down. This

oil then drains overboard through a central tube inthe exhaust unit inner cone.


12. The oil tank (fig. 8-4) is usually mounted on theengine and is normally a separate unit although itmay also be an integral part of the external gearbox.It must have provision to allow the lubrication systemto be drained and replenished. A sight glass or adipstick must also be incorporated to allow the oil



Fig. 8-4 An oil tank.

system contents to be checked. The filler can beeither the gravity or pressure filling type; on someengines both types are fitted. Provision is also madefor a continuous supply of oil to be made available inaircraft which are designed to operate duringinverted flight conditions. Since air is mixed with theoil in the bearing chambers, a de-aerating device isincorporated within the oil tank which removes the airfrom the returning oil.

13. The oil pumps are vital to the efficient operationof the engine. Failure of the pumps will necessitate arapid shutdown of the engine. For this reason, the oilpump driveshafts do not incorporate a weak shear-neck (Part 7) because they must continue to supplyoil for as long as possible, regardless of damage.

14. As the feed oil is distributed to all the lubricatedparts of the engine a substantial amount of sealingair (Part 9) mixes with it and increases its volume.Additionally the bearing chambers operate underdiffering pressures. Therefore, to prevent flooding itis usually necessary to have $. scavenge pump foreach chamber.

15. Gear type pumps are normally used in recircu-latory oil systems but vane and gerotor pumps areemployed in some engines. The simplicity of single-shot pumps (para. 19) make them ideal for engineswhich run for a short duration and use the total losstype of oil system.

16. Gear pumps (fig. 8-5) consist of a pair of inter-meshing steel gears which are housed in a closefitting aluminium casing. When the gears are rotated,oil is drawn into the pump, carried round between theteeth and casing and delivered at the outlet.

17. Since a small quantity of incompressible oilbecomes trapped in the gear mesh, which can causea hydraulic lock and possible pump damage, a reliefslot is machined into the end faces of the casing toprovide an escape route for the oil.

18. Gear pumps are used both as pressure (feed)pumps and scavenge (return) pumps and are incor-porated within a common casing. The oil pumps packis driven by the accessory drive system (Part 7).

19. Single-shot pumps (fig. 8-6) have a quantity ofoil contained within a cylinder. When the piston isforced up the cylinder bore, under the control of thethrottle unit, the oil forces the outlet valves to openallowing a flow of oil to the parts required to belubricated. When the piston reaches the top of thecylinder bore the outlet valves close due to thereduced oil pressure. Recharging of the oil pump

cylinder is achieved by a spring forcing the piston toits original position. This reduces the pressurebetween the cylinder and the oil tank which allowsthe oil replenshing valves to open until the cylinder isrecharged.

20. The most common type of oil distribution deviceis a simple orifice which directs a metered amount ofoil onto its target. These jet orifices are positioned asclose to the target area as possible to overcome thepossibility of the local turbulent environmentdeflecting the jet of oil. The smallest diameter of a jetorifice is 0.04 inch which allows a flow of 12 gallonsper hour when operating at a pressure of 40 lb. persq. in. The use of restrictors upstream can reduce theflow rate if required.

21. All engines transfer heat to the oil by friction,churning and windage within a bearing chamber orgearbox. It is therefore common practice to fit an oilcooler in recirculatory oil systems. The coolingmedium may be fuel or air and, in some instances,both fuel-cooled and air-cooled coolers are used.



Fig. 8-5 Principle of a gear pump.

22. Some engines which utilize both types of coolermay incorporate an electronic monitoring systemwhich switches in the air-cooled cooler only when itis necessary. This maintains the ideal oil temperatureand improves the overall thermal efficiency.

23. The fuel-cooled oil cooler (fig. 8-7) has a matrixwhich is divided into sections by baffle plates. A largenumber of tubes convey the fuel through the matrix,the oil being directed by the baffle plates in a seriesof passes across the tubes. Heat is transferred fromthe oil to the fuel, thus lowering the oil temperature.

24. The fuel-cooled oil cooler incorporates a bypassvalve fitted across the oil inlet and outlet. The valveoperates at a pre-set pressure difference across the

cooler and thus prevents engine oil starvation in theevent of a blockage. A pressure maintaining valve isusually located in the feed line of the cooler whichensures that the oil pressure is always higher thanthe fuel pressure. In the event of a cooler internalfault developing, the oil will leak into the fuel systemrather than the potentially dangerous leakage of fuelinto the oil system.

25. The air-cooled oil cooler is similar to the fuel-cooled type in both construction and operation; themain difference is that air is used as the coolingmedium.

26. Magnetic plugs, or chip detectors (fig. 8-8), arefitted on the scavenge (return) side to collect ferritic



Fig. 8-6 A single-shot oil pump.



Fig. 8-7 A low pressure fuel-cooled oil cooler.

Fig. 8-8 A magnetic chip detector.

debris from each bearing chamber. They arebasically permanent magnets inserted in the oil flowand are retained in self-sealing valve housings.Safety features incorporated in the design ensurecorrect retention within the housing. Uponexamination they can provide a warning ofimpending failure without having to remove andinspect the filters. They are designed to be removedduring maintenance inspection, for condition,monitoring purposes (Part 24), without oil lossoccurring. Additionally they may be connected to acockpit warning system to give an in-flight indication.

27. In some engines, to minimize the effect of thedynamic loads transmitted from the rotatingassemblies to the bearing housings, a 'squeeze film'type of bearing is used (fig. 8-9). They have a smallclearance between the outer race of the bearing andhousing with the clearance being filled with oil. Theoil film dampens the radial motion of the rotatingassembly and the dynamic loads transmitted to thebearing housing thus reducing the vibration level ofthe engine and the possibility of damage by fatigue.

28. To prevent excessive air pressure within the oiltank, gearboxes and bearing chambers, a vent toatmosphere is incorporated within the lubrication

system. Any oil droplets in the air are separated outby a centrifugal breather prior to the air being ventedoverboard. Some breathers may incorporate aporous media, forming de-aerator segments, whichimproves the efficiency of the oil separation (fig, 8-10).



Fig. 8-9 A squeeze film bearing.

Fig. 8-10 A centrifugal breather.

Fig. 8-11 A thread-type oil filter.

29. To prevent foreign matter from continuouslycirculating around the lubricating system, a numberof filters and strainers are positioned within thesystem.

30. Coarse strainers are usually fitted at the outletof the oil tank or immediately prior to the inlet of theoil pumps to prevent debris from damaging thepumps. A fine pressure filter is fitted at the pressure

pump outlet which retains any small particles whichcould block the oil feed jets. Thread-type filters (fig.8-11) are often fitted as a 'last chance' filterimmediately upstream of the oil jets. Sometimesperforated plates or gauze filters are used for thisapplication, Scavenge filters are fitted in each oilreturn line to collect any debris from the lubricatedcomponents. An example of a pressure andscavenge filter is shown in fig. 8-12. They are



Fig. 8-12 A typical pressure and scavenge filter.

invariably of tubular construction with a pleatedwoven wire cloth, or a resin impregnated with fibres,as the filtering medium. Some filters comprise one ormore wire wound elements but these tend to beinsufficient for fine filtration. A 'pop up indicator' maybe fitted to the filter housing to give a visual warningof a partially blocked filter.


31. Early gas turbines used thinner oils than thoseused in piston engines but were produced from thesame mineral crude oil. As gas turbines weredeveloped to operate at higher speeds and tempera-tures these mineral oils oxidized and blocked thefilters and oilways. The development of low viscosity(thin) synthetic oils overcame the major problemsencountered with the early mineral oils.

32. The choice of a lubricating oil is initially decidedby the need to start the engine at very low tempera-tures, when the viscosity of the oil is high, whilst

being able to survive in an engine environment whichexhibits very high temperatures. Having met thesefundamental requirements, the need to provideimproved lubrication characteristics using additivesmust also be investigated. Special laboratory andengine tests are done to prove the suitability of aparticular oil for a specific type of engine.Assessments are made as the extent to which itdeteriorates and the corrosive effects it may have onthe engine.

33. Most gas turbines use a low viscosity oil due tothe absence of reciprocating parts and heavy dutygearing. This reduces the power required for starting,particularly at low temperatures. In fact normal startscan be made in temperatures as low as -40 deg. C.without having to pre-heat the oil.

34. Turbo-propeller engines use a slightly higherviscosity oil due to the additional requirements of thereduction gear and propeller pitch changemechanism.



Rolls-Royce RB162-86

Armstrong Siddeley Mamba

The Mamba axial-flow turbo-prop wasconceived in 1945 as a 1000 hp engine. Firstrun in April 1946, the single Mamba eventuallywent into service with the Short Seamew at1770 ehp. A further development was theDouble Mamba, a combination of two singleMambas in one power unit. Providing up to3875 ehp, the Double Mamba saw servicewith the Fairey Gannet.

9: Internal air system

Contents Page

Introduction 85Cooling 86

Turbine coolingBearing chamber coolingAccessory cooling

Sealing 89Labyrinth seals Ring seals Hydraulic seals Carbon seals Brush seals Hot gas ingestion

Control of bearing loads 91Aircraft services 93


1. The engine internal air system is defined asthose airflows which do not directly contribute to theengine thrust. The system has several importantfunctions to perform for the safe and efficientoperation of the engine. These functions include

internal engine and accessory unit cooling, bearingchamber sealing prevention of hot gas ingestion intothe turbine disc cavities, control of bearing axialloads, control of turbine blade tip clearances (Part 5)and engine anti-icing (Part 13). The system alsosupplies air for the aircraft services. Up to one fifth ofthe total engine core mass airflow may be used forthese various functions.

2. An increasing amount of work is done on the air,as it progresses through the compressor, to raise its


pressure and temperature. Therefore, to reduceengine performance losses, the air is taken as earlyas possible from the compressor commensurate withthe requirement of each particular function. Thecooling air is expelled overboard via a vent system orinto the engine main gas stream, at the highestpossible pressure, where a small performancerecovery is achieved.


3. An important consideration at the design stage ofa gas turbine engine is the need to ensure thatcertain parts of the engine, and in some instancescertain accessories, do not absorb heat to the extentthat is detrimental to their safe operation. Theprincipal areas which require air cooling are thecombustor and turbine. Refer to Part 4 for combustorcooling techniques.

4. Cooling air is used to control the temperature ofthe compressor shafts and discs by either cooling orheating them. This ensures an even temperature dis-tribution and therefore improves engine efficiency bycontrolling thermal growth and thus maintainingminimum blade tip and seal clearances. Typicalcooling and sealing airflows are shown in fig. 9-1.

Turbine cooling5. High thermal efficiency is dependent upon highturbine entry temperature, which is limited by theturbine blade and nozzle guide vane materials.Continuous cooling of these components allows theirenvironmental operating temperature to exceed thematerial's melting point without affecting the bladeand vane integrity. Heat conduction from the turbineblades to the turbine disc requires the discs to becooled and thus prevent thermal fatigue and uncon-trolled expansion and contraction rates.

Internal air system


Fig. 9-1 General internal airflow pattern.

Internal air system


Fig. 9-2 Nozzle guide vane and turbine blade cooling arrangement.

6. An air cooled high pressure nozzle guide vaneand turbine blade arrangement illustrating thecooling airflow is shown in fig. 9-2. Turbine vane andturbine blade life depends not only on their form butalso on the method of cooling, therefore the flowdesign of the internal passages is important. Therehave been numerous methods of turbine vane andturbine blade cooling which have been usedthroughout the history of gas turbines. Generally,single pass internal (convection) cooling was of greatpractical benefit but development has lead to multi-pass internal cooling of blades, impingement coolingof vanes with external air film cooling of both vanesand blades, these are shown in fig. 9-3. and fig. 9-4.

7. The 'pre-swirl nozzles' (fig. 9-2) reduce thetemperature and pressure of the cooling air fed to thedisc for blade cooling. The nozzles also impart a

substantial whirl velocity to assist efficient entry ofthe air into the rotating cooling passages.

8. Cooling air for the turbine discs enters theannular spaces between the discs and flowsoutwards over the disc faces. Flow is controlled byinterstage seals and, on completion of the coolingfunction, the air is expelled into the main gas stream(fig. 9-5); see para. 23., Hot gas ingestion.

Bearing chamber cooling9. Air cooling of the engine bearing chambers is notnormally necessary since the lubrication system(Part 8) is adequate for cooling purposes.Additionally, bearing chambers are located, wherepossible, in the cooler regions of the engine. Ininstances where additional cooling is required, it isgood practice to have a double skinned bearinghousing with cooling air fed into the intermediatespace.

Internal air system


Fig. 9-3 Development of high pressure turbine blade cooling.

Accessory cooling10. A considerable amount of heat is produced bysome of the engine accessories, of which theelectrical generator is an example, and these mayoften require their own cooling circuit. When air isused for cooling, the source may be the compressoror atmospheric air ducted from intake louvres in theengine cowlings.

11. When an accessory unit is cooled during flightby atmospheric air it is usually necessary to providean induced circuit for use during static groundrunning when there would be no external airflow. Thisis achieved by allowing compressor delivery air topass through nozzles situated in the cooling air outletduct of the accessory. The air velocity through thenozzles create a low pressure area which forms anejector, so inducing a flow of atmospheric air throughthe intake louvres. To ensure that the ejector systemonly operates during ground running, the flow of airfrom the compressor is controlled by a valve. Agenerator cooling system with an ejector is shown infig. 9-6.


12. Seals are used to prevent oil leakage from theengine bearing chambers, to control cooling airflows

and to prevent ingress of the mainstream gas into theturbine disc cavities.

13. Various sealing methods are used on gasturbine engines. The choice of which method isdependent upon the surrounding temperature andpressure, wearability, heat generation, weight, spaceavailable, ease of manufacture and ease of installa-tion and removal. Some of the sealing methods aredescribed in the following paragraphs. A hypotheticalturbine showing the usage of these seals is shown infig. 9-5.

Labyrinth seals14. This type of seal is widely used to retain oil inbearing chambers and as a metering device tocontrol internal airflows. Several variations oflabyrinth seal design are shown in fig. 9-7.

15. A labyrinth seal comprises a finned rotatingmember with a static bore which is lined with a softabradable material, or a high temperaturehoneycomb structure. On initial running of the enginethe fins lightly rub against the lining, cutting into it togive a minimum clearance. The clearance variesthroughout the flight cycle, dependent upon thethermal growth of the parts and the natural flexing ofthe rotating members. Across each seal fin there is apressure drop which results in a restricted flow of

Internal air system


Fig. 9-4 High pressure nozzle guide vane construction and cooling.

Internal air system


Fig. 9-5 A hypothetical turbine cooling and sealing arrangement.

sealing air from one side of the seal to the other.When this seal is used for bearing chamber sealing,it prevents oil leakage by allowing the air to flow fromthe outside to the inside of the chamber. This flowalso induces a positive pressure which assists the oilreturn system.

16. Seals between two rotating shafts are morelikely to be subject to rubs between the fins andabradable material due to the two shafts deflectingsimultaneously. This will create excessive heat whichmay result in shaft failure. To prevent this, a non-heatproducing seal is used where the abradable lining isreplaced by a rotating annulus of oil. When the shaftsdeflect, the fins enter the oil and maintain the sealwithout generating heat (fig. 9-7).

Ring seals17. A ring seal (fig. 9-7) comprises a metal ringwhich is housed in a close fitting groove in the statichousing. The normal running clearance between thering and rotating shaft is smaller than that which canbe obtained with the labyrinth seal. This is becausethe ring is allowed to move in its housing wheneverthe shaft comes into contact with it.

18. Ring seals are used for bearing chambersealing, except in the hot areas where oildegradation due to heat would lead to ring seizurewithin its housing.

Hydraulic seals19. This method of sealing is often used betweentwo rotating members to sea a bearing chamber.Unlike the labyrinth or ring seal, it does not allow acontrolled flow of air to traverse across the seal,

20. Hydraulic seals (fig. 9-7) are formed by a sealfin immersed in an annulus of oil which has beencreated by centrifugal forces. Any difference in airpressure inside and outside of the bearing chamberis compensated by a difference in oil level either sideof the fin.

Carbon seals21. Carbon seals (fig. 9-7) consist of a static ring ofcarbon which constantly rubs against a collar on arotating shaft. Several springs are used to maintaincontact between the carbon and the collar. This typeof seal relies upon a high degree of contact and doesnot allow oil or air leakage across it. The heat causedby friction is dissipated by the oil system.

Brush seals22. Brush seals (fig. 9-7) comprise a static ring offine wire bristles. They are in continuous contact witha rotating shaft, rubbing against a hard ceramiccoating. This type of seal has the advantage of with-standing radial rubs without increasing leakage.

Hot gas ingestion23. It is important to prevent the ingestion of hotmainstream gas into the turbine disc cavities as thiswould cause overheating and result in unwantedthermal expansion and fatigue. The pressure in theturbine annulus forces the hot gas, between therotating discs and the adjacent static parts, into theturbine disc rim spaces. In addition, air near the faceof the rotating discs is accelerated by friction causingit to be pumped outwards. This induces a comple-mentary inward flow of hot gas.

24. Prevention of hot gas ingestion is achieved bycontinuously supplying the required quantity ofcooling and sealing air into the disc cavities tooppose the inward flow of hot gas. The flow andpressure of the cooling and sealing air is controlledby interstage seals (fig. 9-5),


25. Engine shafts experience varying axial gasloads (Part 20) which act in a forward direction on thecompressor and in a rearward direction on theturbine. The shaft between them is therefore alwaysunder tension and the difference between the loadsis carried by the location bearing which is fixed in astatic casing (fig. 9-8). The internal air pressure acts

Internal air system


Fig. 9-6 A generator cooling system.

Internal air system


Fig. 9-7 Typical seals.

upon a fixed diameter pressure balance seal toensure the location bearing is adequately loadedthroughout the engine thrust range.


26. To provide cabin pressurization, airframe anti-icing and cabin heat, substantial quantities of air are

bled from the compressor. It is desirable to bleed theair as early as possible from the compressor tominimize the effect on engine performance.However, during some phases of the flight cycle itmay be necessary to switch the bleed source to alater compressor stage to maintain adequatepressure and temperature.

Internal air system


Fig. 9-8 Control of axial bearing load.

Rolls-Royce Gem 60

Rolls-Royce AJ65 Avon

Work commenced early in 1945 on the AJ65axial flow turbo-jet with a design thrust of 6500lb. This figure was reached in 1951 with the100 series RA3. In 1953 the considerablyredesigned 200 series RA14 was type testedat 9500 lb thrust. Development culminated inthe 300 series RB146 which produced 17.110lb thrust with afterburning.

10: Fuel system

Contents Page

Introduction 95Manual and automatic control 96Fuel control systems 99

Pressure control (turbo-propeller engine)Pressure control (turbo-jet engine)Flow controlCombined acceleration and speedcontrolPressure ratio control

Electronic engine control 111Speed and temperature control amplifiersEngine supervisory control

Low pressure fuel system 112Fuel pumps 112

Plunger-type fuel pump Gear-type fuel pump

Fuel spray nozzles 114Fuel heating 116Effect of a change of fuel 116Gas turbine fuels 117

Fuel requirements Vapour locking and boiling Fuel contamination control


1. The functions of the fuel system are to providethe engine with fuel in a form suitable for combustionand to control the flow to the required quantitynecessary for easy starting, acceleration and stablerunning, at all engine operating conditions. To dothis, one or more fuel pumps are used to deliver the

fuel to the fuel spray nozzles, which inject it into thecombustion system (Part 4) in the form of anatomized spray. Because the flow rate must varyaccording to the amount of air passing through theengine to maintain a constant selected engine speedor pressure ratio, the controlling devices are fullyautomatic with the exception of engine powerselection, which is achieved by a manual throttle or


power lever. A fuel shut-off valve (cock) control leveris also used to stop the engine, although in someinstances these two manual controls are combinedfor single-lever operation.

2. It is also necessary to have automatic safetycontrols that prevent the engine gas temperature,compressor delivery pressure, and the rotatingassembly speed, from exceeding their maximumlimitations.

3. With the turbo-propeller engine, changes inpropeller speed and pitch have to be taken intoaccount due to their effect on the power output of theengine. Thus, it is usual to interconnect the throttlelever and propeller controller unit, for by so doing thecorrect relationship between fuel flow and airflow ismaintained at all engine speeds and the pilot is givensingle-lever control of the engine. Although themaximum speed of the engine is normallydetermined by the propeller speed controller, over-speeding is ultimately prevented by a governor in thefuel system.

4. The fuel system often provides for ancillaryfunctions, such as oil cooling (Part 8) and thehydraulic control of various engine control systems;for example, compressor airflow control (Part 3).


5. The control of power or thrust of the gas turbineengine is effected by regulating the quantity of fuelinjected into the combustion system. When a higherthrust is required, the throttle is opened and thepressure to the fuel spray nozzles increases due tothe greater fuel flow. This has the effect of increasingthe gas temperature, which in turn increases theacceleration of the gases through the turbine to givea higher engine speed and a correspondingly greaterairflow, consequently producing an increase inengine thrust.

6. This relationship between the airflow inducedthrough the engine and the fuel supplied is, however,complicated by changes in altitude, air temperatureand aircraft speed. These variables change thedensity of the air at the engine intake and conse-quently the mass of air induced through the engine.A typical change of airflow with altitude is shown infig. 10-1. To meet this change in airflow a similarchange in fuel flow (fig. 10-2) must occur, otherwisethe ratio of airflow to fuel flow will change and willincrease or decrease the engine speed from thatoriginally selected by the throttle lever position.

7. Described in this Part are five representativesystems of automatic fuel control; these are thepressure control and flow control systems, which are

Fuel system


Fig. 10-1 Airflow changing with altitude.

Fig. 10-2 Fuel flow changing with altitude.

Fuel system


Fig. 10-3 Simplified fuel systems for turbo-propeller and turbo-jet engines.

Fuel system


Fig. 10-4 A pressure control system (turbo-propeller engine).

hydro-mechanical, and the acceleration and speedcontrol and pressure ratio control systems, which aremechanical. With the exception of the pressure ratiocontrol system, which uses a gear-type pump, all thesystems use a variable-stroke, multi-plunger typefuel pump to supply the fuel to the spray nozzles.

8. Some engines are fitted with an electronicsystem of control and this generally involves the useof electronic circuits to measure and translatechanging engine conditions to automatically adjustthe fuel pump output. On helicopters powered by gasturbine engines using the free-power turbineprinciple (Part 5), additional manual and automaticcontrols on the engine govern the free-power turbineand, consequently, aircraft rotor speed.


9. Typical high pressure (H.P.) fuel control systemsfor a turbo-propeller engine and a turbo-jet engineare shown in simplified form in fig. 10-3, eachbasically consisting of an H.P. pump, a throttlecontrol and a number of fuel spray nozzles. Inaddition, certain sensing devices are incorporated toprovide automatic control of the fuel flow in responseto engine requirements. On the turbo-propellerengine, the fuel and propeller systems are co-ordinated to produce the appropriate fuel/r.p.m.combination.

10. The usual method of varying the fuel flow to thespray nozzles is by adjusting the output of the H.P.fuel pump. This is effected through a servo system inresponse to some or all of the following:

(1) Throttle movement.(2) Air temperature and pressure.(3) Rapid acceleration and deceleration.(4) Signals of engine speed, engine gas

temperature and compressor deliverypressure.

Pressure control (turbo-propeller engine)11. The pressure control system (fig. 10-4) is atypical system as fitted to a turbo-propeller enginewhere the rate of engine acceleration is restricted bya propeller speed controller. The fuel pump output isautomatically controlled by spill valves in the flowcontrol unit (F.C.U.) and the engine speed governor.These valves, by varying the fuel pump servopressure, adjust the pump stroke to give the correctfuel flow to the engine.

12. At steady running conditions, at a given airintake pressure and below governed speed, the spillvalve in the F.C.U. is in a sensitive position, creating

a balance of forces across the fuel pump servopiston and ensuring a steady pressure to the throttlevalve.

13. When the throttle is slowly opened, thepressure to the throttle valve falls and allows theF.C.U. spill valve to close, so increasing the servopressure and pump delivery. As the pressure to thethrottle is restored, the spill valve returns to itssensitive or controlling position, and the fuel pumpstabilizes its output to give the engine speed for theselected throttle position. The reverse sequenceoccurs as the throttle is closed.

14. A reduction of air intake pressure, due to areduction of aircraft forward speed or increase inaltitude, causes the F.C.U. capsule to expand, thusincreasing the bleed from the F.C.U. spill valve. Thisreduces fuel pump delivery until the fuel flowmatches the airflow and the reduced H.P. pumpdelivery (throttle inlet pressure), allows the spill valveto return to its sensitive position. Conversely, anincrease in air intake pressure reduces the bleedfrom the spill valve and increases the fuel flow. Thecompensation for changes in air intake pressure issuch that fuel flow cannot be increased beyond thepre-determined maximum permissible for staticInternational Standard Atmosphere (I.S.A.) sea-levelconditions.

15. The engine speed governor prevents the enginefrom exceeding its maximum speed limitation. Withincreasing engine speed, the centrifugal pressurefrom the fuel pump rotor radial drillings increases andthis is sensed by the engine speed governordiaphragm. When the engine reaches its speedlimitation, the diaphragm is deflected to open thegovernor spill valve, thus overriding the F.C.U. andpreventing any further increase in fuel flow. Somepressure control systems employ a hydro-mechanical governor (para. 23).

16. The governor spill valve also acts as a safetyrelief valve. If the fuel pump delivery pressureexceeds its maximum controlling value, the servopressure acting on the orifice area of the spill valveforces the valve open regardless of the enginespeed, so preventing any further increase in fueldelivery pressure.

Pressure control (turbo-jet engine)17. In the pressure control system illustrated in fig.10-5, the rate of engine acceleration is controlled bya dashpot throttle unit. The unit forms part of the fuelcontrol unit and consists of a servo-operated throttle,which moves in a ported sleeve, and a control valve.

Fuel system


Fuel system


Fig. 10-5 A pressure control system (turbo-jet engine).

The control valve slides freely within the bore of thethrottle valve and is linked to the pilot's throttle by arack and pinion mechanism. Movement of the throttlelever causes the throttle valve to progressivelyuncover ports in the sleeve and thus increase thefuel flow. Fig. 10-6 shows the throttle valve andcontrol valve in their various controlling positions.

18. At steady running conditions, the dashpotthrottle valve is held in equilibrium by throttle servopressure opposed by throttle control pressure plusspring force. The pressures across the pressure dropcontrol diaphragm are in balance and the pumpservo pressure adjusts the fuel pump to give aconstant fuel flow.

19. When the throttle is opened, the control valvecloses the low pressure (L.P.) fuel port in the sleeveand the throttle servo pressure increases. Thethrottle valve moves towards the selected throttleposition until the L.P. port opens and the pressurebalance across the throttle valve is restored. Thedecreasing fuel pressure difference across thethrottle valve is sensed by the pressure drop controldiaphragm, which closes the spill valve to increasethe pump servo pressure and therefore the pumpoutput. The spill valve moves into the sensitiveposition, controlling the pump servo mechanism sothat the correct fuel flow is maintained for theselected throttle position.

20. During initial acceleration, fuel control is asdescribed in para. 19; however, at a predeterminedthrottle position the engine can accept more fuel andat this point the throttle valve uncovers an annulus,so introducing extra fuel at a higher pressure (pumpdelivery through one restrictor). This extra fuel furtherincreases the throttle servo pressure, whichincreases the speed of throttle valve travel and therate of fuel supply to the spray nozzle.

21. On deceleration, movement of the control valveacts directly on the throttle valve through the servospring. Control valve movement opens the flow portsthrough the control valve and throttle valve, to bleedservo fuel through the L.P. port. Throttle controlpressure then moves the throttle valve towards theclosed position, thus reducing the fuel flow to thespray nozzles.

22. Changes in air intake pressure, due to a changein aircraft altitude or forward speed, are sensed bythe capsule assembly in the fuel control unit. Withincreased altitude and a corresponding decrease inair intake pressure, the evacuated capsule opens thespill valve, so causing a reduction in pump stroke

Fuel system


Fig. 10-6 Acceleration control by dashpotthrottle.

Fuel system


Fig. 10-7 A proportional flow control system.

until the fuel flow matches the airflow. Conversely, anincrease in air intake pressure closes the spill valveto increase the fuel flow.

23. H.P. compressor shaft r.p.m. is governed by ahydro-mechanical governor which uses hydraulicpressure proportional to engine speed as itscontrolling parameter. A rotating spill valve sensesthe engine speed and the controlling pressure isused to limit the pump stroke and so prevent over-speeding of the H.P. shaft rotating assembly. Thecontrolling pressure is unaffected by changes in fuelspecific gravity.

24. At low H.P. shaft speeds, the rotating spill valveis held open, but as engine speed increases,centrifugal loading moves the valve towards theclosed position against the diaphragm loads. Thisrestricts the bleed of fuel to the L.P. side of the valveuntil, at governed speed, the governor pressuredeflects the servo control diaphragm and opens theservo spill valve to control the fuel flow and therebythe H.P. shaft speed.

25. If the engine gas temperature attempts toexceed the maximum limitation, the current in theL.P. speed limiter and temperature control solenoid isreduced. This opens the spill valve to reduce thepressure on the pressure drop control diaphragm.The flow control spill valve then opens to reduce thepump servo pressure and fuel pump output.

26. To prevent the L.P. compressor from over-speeding, multi-spool engines usually have an L.P.compressor shaft speed governor. A signal of L.P.shaft speed and intake temperature is fed to anamplifier and solenoid valve, the valve limiting thefuel flow in the same way as the gas temperaturecontrol (para. 25).

27. The system described uses main and startingspray nozzles under the control of an H.P. shut-offvalve. Two starting nozzles are fitted in thecombustion chamber, each being forward of anigniter plug. When the engine has started, the fuelflow to these nozzles is cut off by the H.P. shut-offvalve.

28. To ensure that a satisfactory fuel pressure to thespray nozzles is maintained at high altitudes, a backpressure valve, located downstream of the throttlevalve, raises the pressure levels sufficiently toensure satisfactory operation of the fuel pump servosystem.

Flow control29. A flow control fuel system is generally morecompact than a pressure control system and is notsensitive to flow effect of variations downstream ofthe throttle. The fuel pump delivery pressure isrelated to engine speed; thus, at low engine speedspump delivery pressure is quite low. The fuel pumpoutput is controlled to give a constant pressuredifference across the throttle valve at a constant airintake condition. Various devices are also used toadjust the fuel flow for air intake pressure variations,idling and acceleration control, gas temperature andcompressor delivery pressure control.

30. A variation of the flow control system is the pro-portional flow control system (fig 10-7), which is moresuitable for engines requiring large fuel flows andwhich also enables the fuel trimming devices toadjust the fuel flow more accurately. A smallcontrolling flow is created that has the same charac-teristics as the main flow, and this controlling or pro-portional flow is used to adjust the main flow.

31. A different type of spill valve, referred to as akinetic valve, is used in this system. This valveconsists of two opposing jets, one subjected to pumpdelivery pressure and the other to pump servopressure, and an interrupter blade that can be movedbetween the jets (fig. 10-8). When the blade is clearof the jets, the kinetic force of the H.P. fuel jet causesthe servo pressure to rise (spill valve closed) and thefuel pump moves to maximum stroke to increase thefuel flow. When the blade is lowered between thejets, the pressure jet is deflected and the servopressure falls, so reducing the pump stroke and thefuel flow, When the engine is steadily running, theblade is in an intermediate position allowing a slowbleed from servo and thus balancing the fuel pumpoutput.

32. All the controlling devices, except for the enginespeed governor, are contained in one combined fuelcontrol unit. The main parts of the control unit are thealtitude sensing unit (A.S.U.), the accelerationcontrol unit (A.C.U.), the throttle and pressurizingvalve unit, and the proportioning valve unit.

33. At any steady running condition below governedspeed, the fuel pump delivery is controlled to a fixedvalue by the A.S.U. The spill valve in this unit is heldin the controlling position by a balance of forces,spring force and the piston force. The piston issensitive to the pressure difference across thesensing valve, the pressure difference being createdby fuel flowing from the proportioning valve back tothe fuel pump inlet.

Fuel system


34. The proportioning valve diaphragm is held openin a balanced condition allowing fuel to pass to theA.S.U. This means that the restrictor outlet pressureis equal to the throttle outlet pressure and, as theirinlet pressures are equal, it follows that the pressuredifference across the restrictors and the throttle areequal; therefore, a constant fuel flow is obtained.

35. When the throttle is slowly opened, thepressure difference across the throttle valve and theproportioning flow restrictors decreases and the pro-portioning valve diaphragm adjusts its position. Thisreduces the proportional flow, which closes theA.S.U. spill valve and increases the servo pressure.The fuel pump increases its delivery and this restoresthe pressure difference across the throttle valve and

equalizes the pressure difference across therestrictors. The proportional flow is restored to itsoriginal value and the balance of forces in the A.S.U.returns the spill valve to the controlling position.

36. A variation of air intake pressure, due to achange of aircraft forward speed or altitude, issensed by the capsule in the A.S.U. A pressurereduction causes the A.S.U. capsule to expand, thusincreasing the bleed from the spill valve. Thisreduces fuel pump delivery until the fuel flowmatches the airflow and results in a lower pressuredifference across the throttle valve and the propor-tioning valve restrictors. The reduced proportionalflow restores the balance in the A.S.U. which returnsthe spill valve to its controlling position. Conversely,an increase in aircraft forward speed increases theair Intake pressure, which reduces the bleed from thespill valve and increases the fuel flow.

37. During a rapid acceleration, the suddendecrease in throttle pressure difference is sensed bythe A.S.U., causing the spill valve to close, Such arapid increase in fuel supply would, however, createan excessive gas temperature and also cause thecompressor to surge (Part 3). This occurs becausethe inertia of the rotating assembly results in anappreciable time lag in the rate of airflow increase. Itis essential therefore, to have an acceleration controlto override the A.S.U. to give a corresponding lag inthe rate of fuel flow increase.

38. The rapid initial increase of fuel flow causes arise in the pressure difference across the fuelmetering plunger and this is sensed by a diaphragmin the pressure drop control section. At a fixed valueof over fuelling, the pressure drop control diaphragmopens its servo spill valve to override the A.S.U, andmaintains a constant pressure difference across themetering plunger.

39. The increased fuel supply causes the engine toaccelerate and the fuel metering plunger gives themaximum permissible fuel flow to match theincreasing compressor delivery pressure. This itachieves through the A.C.U. servo system, which isunder the control of a spill valve operated bycompressor delivery air pressure acting on acapsule.

40. As the compressor delivery pressure continuesto rise, the capsule is compressed to open the spillvalve and to bleed pressure from above the meteringplunger. Pump delivery pressure acting underneaththe plunger causes it to lift, this increases the area ofthe main fuel flow passage.

Fuel system


Fig. 10-8 Servo pressure control by kineticvalve.

41. The pressure drop control spill valve closes toincrease the fuel pump delivery and maintains thecontrolling pressure difference across the plunger.The fuel flow, therefore, progressively rises as airflowthrough the compressor increases. The degree ofoverfuelling can be automatically changed by the airswitch, which increases the pressure signal on to thecapsule. The full value of compressor deliverypressure is now passed on to the A.C.U. capsuleassembly, thus increasing the opening rate of themetering plunger.

42. As the controlled overfuelling continues, thepressure difference across the throttle valveincreases. When it reaches the controlling value, theA.S.U. takes over due to the increasing proportionalflow and again gives a steady fuel flow to the spraynozzles.

43. The engine speed governor can be of thepressure control type described in para. 15, or ahydro-mechanical governor as described in para. 23.

44. The control of servo pressure by the hydro-mechanical governor is very similar to that of thepressure control governor, except that the governorpressure is obtained from pump delivery fuel passingthrough a restrictor and the restricted pressure iscontrolled by a rotating spill valve; this type ofgovernor is unaffected by changes in fuel specificgravity.

45. At low engine speeds, the rotating spill valve isheld open; however, as engine speed increases,centrifugal loading moves the valve towards theclosed position against the diaphragm loads. Thisrestricts the bleed of H.P. fuel to the L.P. side of thedrum until, at governed speed, the governorpressure deflects the diaphragm and opens the fuelpump servo pressure spill valve to control themaximum fuel flow and engine speed.

46. If the engine gas temperature exceeds itsmaximum limitation, the solenoid on the proportion-ing valve unit is progressively energized. This causesa movement of the rocker arm to increase theeffective flow area of one restrictor, thus increasingthe proportional flow and opening the A.S.U. spillvalve to reduce servo pressure. The fuel flow is thusreduced and any further increase of gas temperatureis prevented.

47. To prevent the L.P. compressor from over-speeding, some twin-spool engines have an L.P.shaft r.p.m. governor. A signal of L.P. shaft speed isfed to an amplifier and solenoid valve, which limits

the fuel output in the same way as the gastemperature control.

48. An idling speed governor is often fitted toensure that the idling r.p.m. does not vary withchanging engine loads. A variation of idling r.p.m.causes the rocker arm to move and alter the propor-tional flow, and the A.S.U. adjusts the pump deliveryuntil the correct idling r.p.m. is restored.

49. On some engines, a power limiter is used toprevent overstressing of the engine. To achieve this,compressor delivery pressure acts on the powerlimiter capsule. Excess pressure opens the powerlimiter atmospheric bleed to limit the pressure on theA.C.U. capsule and this controls the fuel flow throughthe metering plunger.

50. To enable the engine to be relit and to preventflame-out at altitude, the engine idling r.p.m. is madeto increase with altitude. To achieve this, someengines incorporate a minimum flow valve that addsa constant minimum fuel flow to that passing throughthe throttle valve.

Combined acceleration and speed control51. The combined acceleration and speed controlsystem (fig. 10-9) is a mechanical system withoutsmall restrictors or spill valves. It is also an all-speedgovernor system and therefore needs no separategovernor unit for controlling the maximum r.p.m. Thecontrolling mechanism is contained in one unit,usually referred to as the fuel flow regulator (F.F.R.).An H.P. fuel pump (para. 85) is used and the fuelpump servo piston is operated by H.P. fuel on oneside and main spray nozzle (servo) pressure on thespring side.

52. The F.F.R. is driven by the engine through agear train and has two centrifugal governors, knownas the speed control governor and the pressure dropcontrol governor. Two sliding valves are also rotatedby the gear train. One valve, known as the variablemetering sleeve, has a triangular orifice, known asthe variable metering orifice (V.M.O.), and this sleeveis given axial movement by a capsule assembly. TheV.M.O. sleeve moves inside a non-rotating governorsleeve that is moved axially by the speed controlgovernor. The other valve, known as the pressuredrop control valve, is provided with axial movementby the pressure drop control governor and has atriangular orifice, known as the pressure drop controlorifice, and a fixed-area rectangular orifice. Thespeed control governor is set by the throttle leverthrough a cam, a spring and a stirrup arm inside theregulator.

Fuel system


Fuel system


Fig. 10-9 A combined acceleration and speed control system.

53. At any steady running condition, the enginespeed is governed by the regulator controlling thefuel flow. The fuel pump delivery is fixed at a constantvalue by applying the system pressure difference tothe fuel pump servo piston. This is arranged tobalance the servo piston spring forces.

54. When the air intake pressure is at a constantvalue, the rotating V.M.O. sleeve is held in a fixedaxial position by the capsule loading. The fixedthrottle setting maintains a set load on the speedcontrol governor and, as the r.p.m. is constant, thegovernor sleeve is held in a fixed position.

55. The fuel pump delivery is passed to the annulussurrounding the V.M.O.; the annulus area iscontrolled by the governor sleeve, and the exposedarea of the orifice is set by the axial position of theV.M.O. sleeve. Consequently, fuel passes to theinside of the sleeve at a constant flow and thereforeat a constant pressure difference.

56. The pressure drop control valve, which alsoforms a piston, senses the pressure differenceacross the V.M.O. and maintains the fuel flow at afixed value in relation to a function of engine speed,by controlling the exposed area of the pressure dropcontrol orifice.

57. When the throttle is slowly opened, the load onthe speed control governor is increased, so movingthe governor sleeve to increase the V.M.O. annulusarea. The effect of opening the V.M.O. is to reducethe pressure difference and this is sensed by thepressure drop control governor, which opens thepressure drop valve. The reduced system pressuredifference is immediately sensed by the fuel pumpservo piston, which increases the pump stroke andconsequently the fuel output. The increasedcompressor delivery pressure acts on the capsuleassembly, which gradually opens the V.M.O. so thatthe fuel flow and engine speed continue to increase.At the speed selected, centrifugal forces acting onthe speed control governor move the governorsleeve to reduce the V.M.O. annulus area. Theresultant increased pressure difference is sensed bythe pressure drop control governor, which adjusts thepressure drop valve to a point at which the pumpservo system gives an output to match the enginerequirements. The function of the governors and thecontrol of the fuel flow is shown diagrammatically infig. 10-10.

58. During a rapid acceleration, the initial degree ofoverselling is mechanically controlled by a stop thatlimits the opening movement of the speed control

governor sleeve. A similar stop also prevents the fuelsupply from being completely cut off by the governorsleeve during a rapid deceleration.

59. Changes in altitude or forward speed of theaircraft vary the fuel flow required to maintain aconstant engine speed. To provide this control, thecapsule assembly senses changes in H.P.compressor inlet and delivery pressures and adjuststhe V.M.O. accordingly. For instance, as the aircraftaltitude increases, the compressor delivery pressurefalls and the capsule assembly expands to reducethe V.M.O. The increased system pressure drop issensed by the fuel pump servo piston, which adjuststhe pump output to match the reduced airflow and somaintain a constant engine speed. Conversely, anincrease in aircraft forward speed causes thecapsule assembly to be compressed and increasethe V.M.O. The reduced system pressure dropcauses the fuel pump to increase its output to matchthe increased airflow.

60. To prevent the maximum gas temperature frombeing exceeded, fuel flow is reduced in response tosignals from thermocouples sensing the temperature(Part 12). When the maximum temperature isreached, the signals are amplified and passed to arotary actuator which adjusts the throttle mechanism.This movement has the same effect on fuel flow asmanual operation of the throttle.

61. To ensure that the engine is not overstressed,the H.P. compressor delivery pressure is controlledto a predetermined value. At this value, a pressurelimiting device, known as a power limiter, reduces thepressure in the capsule chamber, thus allowing thecapsule assembly to expand and reduce the preventing any further increase in fuel flow.

62. A governor prevents the L.P. compressor shaftfrom exceeding its operating limitations and also actsas a maximum speed governor in an event of afailure of the F.F.R. The governor provides a variablerestrictor between the regulator and the main fuelspray nozzle manifold. Should the L.P. compressorreach its speed limitation, flyweights in the governormove a sleeve valve to reduce the flow area, Theincreased system pressure drop is sensed by thefuel pump servo piston, which reduces the fuel flowto the spray nozzles.

63. This fuel system has no pressurizing valve todivide the flow from the fuel pump into main andprimary fuel flows. Primary fuel pressure is takenfrom the fixed-area orifice of the pressure dropcontrol valve. This pressure is always higher than the

Fuel system


Fuel system


Fig. 10-10 Governor movement and fuel flow control.

main fuel pressure and it is not shut off by thepressure drop control piston. It therefore gives a sat-isfactory idling fuel flow at all altitudes.

64. On engines featuring water injection (Part 17), areset device (fig. 10-11), operated by a piston andreset cam, increases the loading on the throttlecontrol spring and stirrup arm, thus selecting a higherengine speed during water injection. To prevent thepower limiter (fig, 10-9) cancelling the effect of waterinjection, a capsule in the limiter is subjected to waterpressure to raise the compressor delivery pressureat which the power limiter operates.

Pressure ratio control65. The pressure ratio control (fig. 10-12) is amechanical system similar to the combined acceler-ation and speed control system, but uses the ratio ofH.P. compressor delivery pressure to air intakepressure (P4/P1) as the main controlling parameter.It needs no separate governor unit for controlling themaximum r.p.m. The controlling mechanism iscontained in one unit, which is usually referred to asa fuel flow regulator (F.F.R.). A gear-type pump isused, as described in para. 88, and the pump outputto the F.F.R. is controlled by a pressure drop spillvalve.

66. The F.F.R. is driven by the engine through agear train and has two rotating valves. One valve,

known as a variable metering sleeve, has atriangular orifice, known as the variable meteringorifice (V.M.O.), and this sleeve is given axialmovement by a capsule assembly. The other valve,known as the pressure drop control valve, isprovided with axial movement by a centrifugalgovernor, known as a pressure drop controlgovernor, Both valves form variable restrictors whichcontrol the fuel flow to the spray nozzles.

67. Control of the V.M.O. area is a function of apressure ratio control unit housed in the F.F.R. Apressure ratio control valve, subjected to P4 and P1,pressures, regulates the movement of the F.F.R.capsule and thus controls the V.M.O. area to producethe pressure ratio dictated by the throttle or powerlever.

68. At any steady running condition, the output ofthe fuel pump is greater than the engine requirement.The pressure drop spill valve is open to allow surplusfuel to return to the inlet side of the pump. This actioncontrols the fuel delivery to that demanded by theF.F.R.

69. When the throttle is slowly opened, the throttle-controlled orifice is increased and the controlpressure falls, thus allowing the pressure ratiocontrol valve to move towards the closed position(acceleration stop). F.F.R. capsule chamber pressure

Fuel system


Fig. 10-11 Effect of water reset on speed control governor.

Fuel system


Fig. 10-12 A pressure ratio control system.

increases and the capsule moves the meteringsleeve to increase the V.M.O. area. The effect ofopening the V.M.O. is to reduce the pressuredifference and this is sensed by the pressure dropgovernor, which opens the pressure drop controlorifice. The reduced system pressure difference isimmediately sensed by the pressure drop spill valve,which moves towards the closed position and conse-quently increases the fuel output. The increased fuelflow accelerates the engine with a subsequentincrease in pressure ratio (P4/P1). When therequired pressure ratio is reached, the pressure ratiocontrol valve opens and the F.F.R. capsule chamberpressure reduces. The capsule assembly expands,moving the V.M.O. sleeve to reduce the orifice area.The resultant increased pressure difference issensed by the pressure drop control governor, whichadjusts the pressure drop control orifice to a point atwhich the pressure drop spill valve gives a fueloutput consistent with steady running requirements.

70. During a rapid acceleration, the degree ofoverselling is mechanically controlled by the acceler-ation stop, which limits the movement of the pressureratio control valve. A similar stop prevents the fuelsupply from being completely cut off during a rapiddeceleration.

71. When accelerating to a higher P4/P1 ratio, thethrottle control orifice is increased. The reducedpressure allows the pressure ratio control capsule tocontract so that the valve contacts the accelerationstop. F.F.R. capsule chamber pressure increasesand the capsule moves to increase the V.M.O. area.This action continues until the required P4/P1 ratio isreached. The increased P4 pressure allows thepressure ratio control capsule to re-expand and thevalve to return to the steady running position.

72. A change in altitude of the aircraft requires avariation in fuel flow to match the engine thrust andaircraft climb requirement. The normal effect of analtitude increase is to decrease the P1 and P4pressures, thus opening the pressure ratio controlvalve and allowing the F.F.R. capsule to expand toreduce the V.M.O. area and, in consequence, thefuel flow. However, to match the engine thrust andaircraft climb requirement it is necessary to increasethe P4/P1 ratio with increasing altitude. This is doneby a trimmer valve and a capsule that is subjected toP1 pressure. As P1 pressure decreases, the trimmervalve moves across the P1 controlled orifice toreduce the control pressure. This is sensed by thecontrol capsule, which, by acting on the pressureratio control valve, slows the closure of the V.M.O. as

altitude is increased. This maintains the thrustrequirement with the throttle at a fixed position.

73. To prevent the maximum L.P. compressor r.p.m.and engine gas temperature from being exceeded, avalve, known as the auxiliary throttling valve, is fittedin the outlet from the fuel pump, Under steadyrunning conditions, the valve is held open by springforce, When limiting conditions are reached, the fuelflow is reduced in response to speed andtemperature signals from the engine. The signals areamplified and passed to a rotary actuator thatreduces the area of a variable restrictor. The effect ofthis is to increase the fuel pressure, which partiallycloses the throttling valve. H.P. fuel pressure actingon the face of the pressure drop spill valve isincreased and the spill valve opens to reduce the fuelflow to the spray nozzles.

74. H.P. shaft speed is also governed by theauxiliary throttling valve. Should other controllingdevices fail and pump speed increases, the fuelpressure closes the throttling valve and opens thepressure drop spill valve to reduce the fuel flow.

75. With the throttle closed, idling condition isdetermined by controlling the amount of air beingvented through the idling adjuster and the groundidling solenoid valve, With both bleeds in operation,satisfactory flight idling for the air off-takes isensured. By closing the solenoid valve a lower powercondition for ground idling is obtained.

76. This fuel system, like the combined accelerationand speed control system, has no pressurizing valveto divide the flow from the fuel pump into main andprimary flows.


77. As stated in para. 8, some engines utilize asystem of electronic control to monitor engineperformance and make necessary control inputs tomaintain certain engine parameters within predeter-mined limits. The main areas of control are engineshaft speeds and exhaust gas temperature (E.G.T.)which are continuously monitored during engineoperation. Some types of electronic control functionas a limiter only, that is, should engine shaft speed orE.G.T. approach the limits of safe operation, then aninput is made to the fuel flow regulator (F.F.R.) toreduce the fuel flow thus maintaining shaft speed orE.G.T. at a safe level. Supervisory control systemsmay contain a limiter function but, basically, by usingaircraft generated data, the system enables a moreappropriate thrust setting to be selected quickly and

Fuel system


accurately by the pilot. The control system thenmakes small control adjustments to maintain enginethrust consistent with that pre-set by the pilot,regardless of changing atmospheric conditions. Fullauthority digital engine control (FAD.E.G.) takes overvirtually all of the steady state and transient controlintelligence and replaces most of the hydromechani-cal and pneumatic elements of the fuel system. Thefuel system is thus reduced to a pump and controlvalve, an independent shut-off cock and a minimumof additional features necessary to keep the enginesafe in the event of extensive electronic failure.

78. Full authority fuel control (F.A.F.C.) provides fullelectronic control of the engine fuel system in thesame way as F.A.D.E.C., but has none of thetransient control intelligence capability used tocontrol the compressor airflow system as the existingengine control system is used for these.

Speed and temperature control amplifiers79. The speed and temperature control amplifierreceives signals from thermocouples measuringE.G.T. and from speed probes sensing L.P. and insome cases, L.P. shaft speeds (N1 and N2). Theamplifier basically comprises speed and temperaturechannels which monitor the signals sensed. If eitherN1, N2 or E.G.T. exceed pre-set datums, theamplifier output stage is triggered to connect anelectrical supply to a solenoid valve (para. 47) or avariable restrictor (para. 73) which override the F.F.R.and cause a reduction in fuel flow. The limiter willonly relinquish control back to the F.F.R. if the inputconditions are altered (altitude, speed, ambienttemperature or throttle lever position). The limitersystem is designed to protect against parametersexceeding their design values under normaloperation and basic fuel system failures.

Engine supervisory control80. The engine supervisory control (E.S.C.) systemperforms a supervisory function by trimming the fuelflow scheduled by the fuel flow governor (F.F.G.) tomatch the actual engine power with a calculatedengine power for a given throttle angle. The E.S.C.provides supervisory and limiting functions by meansof a single control output signal to a torque motor inthe F.F.G. In order to perform its supervisory functionthe E.S.C. monitors inputs of throttle angle, enginebleed state, engine pressure ratio (E.P.R.) and airdata computer information (altitude, Mach numberand temperatures). From this data the supervisorychannel predicts the value of N1 required to achievethe command E.P.R. calculated for the throttle angleset by the pilot. Simultaneously a comparison ismade between the command E.P.R. and the actual

E.P.R. and the difference is compared with aprogrammed datum.

81. During acceleration the comparitor connects thepredicted value of N1 to the limiter channel until thedifference between the command and actual approximately 0.03 E.P.R. At this point thepredicted L.P. shaft speed is disconnected and theE.P.R. difference signal is connected to the limiterchannel.

82. The final output from the supervisory channel,in the form of an error signal, is supplied to a 'lowestwins' circuit along with the error signals from thelimiter channel. While the three error signals remainpositive (N1 and E.G.T. below datum level and actualE.P.R. below command E.P.R.) no output is signalledto the torque motor. If, however, the output stage ofthe E.S.C. predicts that E.G.T. will exceed datum orthat N1 will either exceed its datum or the predictedlevel for the command E.P.R., then a signal is passedto the torque motor to trim the fuel flow.


83. An L.P. system (fig.10-13) must be provided tosupply the fuel to the engine at a suitable pressure,rate of flow and temperature, to ensure satisfactoryengine operation. This system may include an L.P.pump to prevent vapour locking and cavitation of thefuel, and a fuel heater to prevent ice crystals forming.A fuel filter is always used in the system and in someinstances the flow passes through an oil cooler (Part8). Transmitters may also be used to signal fuelpressure, flow and temperature (Part 12).


84. There are two basic types of fuel pump, theplunger-type pump and the constant-delivery gear-type pump; both of these are positive displacementpumps. Where low pressures are required at the fuelspray nozzles, the gear-type pump is preferredbecause of its lightness.

Plunger-type fuel pump85. The pump shown in fig. 10-14 is of the single-unit, variable-stroke, plunger-type; similar pumpsmay be used as double units depending upon theengine fuel flow requirements.

86. The fuel pump is driven by the engine gear trainand its output depends upon its rotational speed andthe stroke of the plungers. A single-unit fuel pumpcan deliver fuel at the rate of 100 to 2,000 gallons perhour at a maximum pressure of about 2,000 lb. per

Fuel system


Fuel system


Fig. 10-14 A plunger-type fuel pump.

Fig. 10-14 A low pressure system.

square inch. To drive this pump, as much as 60horsepower may be required.

87. The fuel pump consists of a rotor assemblyfitted with several plungers, the ends of which projectfrom their bores and bear on to a non-rotatingcamplate. Due to the inclination of the camplate,movement of the rotor imparts a reciprocating motionto the plungers, thus producing a pumping action.The stroke of the plungers is determined by the angleof inclination of the camplate. The degree ofinclination is varied by the movement of a servopiston that is mechanically linked to the camplateand is biased by springs to give the full strokeposition of the plungers. The piston is subjected toservo pressure on the spring side and on the otherside to pump delivery pressure; thus variations in thepressure difference across the servo piston cause itto move with corresponding variations of thecamplate angle and, therefore, pump stroke.

Gear-type fuel pump88. The gear-type fuel pump (fig. 10-12) is drivenfrom the engine and its output is directly proportionalto its speed. The fuel flow to the spray nozzles iscontrolled by recirculating excess fuel delivery backto inlet. A spill valve, sensitive to the pressure dropacross the controlling units in the system, opens andcloses as necessary to increase or decrease thespill.


89. The final components of the fuel system are thefuel spray nozzles, which have as their essentialfunction the task of atomizing or vaporizing the fuel toensure its rapid burning. The difficulties involved inthis process can be readily appreciated when oneconsiders the velocity of the air stream from thecompressor and the short length of combustionsystem (Part 4) in which the burning must becompleted.

90. An early method of atomizing the fuel is to passit through a swirl chamber where tangentiallydisposed holes or slots imparted swirl to the fuel byconverting its pressure energy to kinetic energy. Inthis state, the fuel is passed through the dischargeorifice which removes the swirl motion as the fuel isatomized to form a cone-shaped spray. This is called'pressure jet atomization'. The rate of swirl andpressure of the fuel at the fuel spray nozzle areimportant factors in good atomization. The shape ofthe spray is an indication of the degree ofatomization as shown in fig. 10-15. Later fuel spraynozzles utilize the airspray principle which employs

high velocity air instead of high velocity fuel to causeatomization. This method allows atomization at lowfuel flow rates (provided sufficient air velocity exists)thus providing an advantage over the pressure jetatomizer by allowing fuel pumps of a lighter con-struction to be used.

91. The atomizing spray nozzle, as distinct from thevaporizing burner (Part 4), has been developed infive fairly distinct types; the Simplex, the variable port(Lubbock), the Duplex or Duple, the spill type and theairspray nozzle.

92. The Simplex spray nozzle shown in fig. 10-16was first used on early jet engines. It consists of achamber, which induces a swirl into the fuel, and afixed-area atomizing orifice. This fuel spray nozzlegave good atomization at the higher fuel flows, that

Fuel system


Fig. 10-15 Various stages of fuel atomization.

is, at the higher fuel pressures, but was very unsatis-factory at the low pressures required at low enginespeeds and especially at high altitudes. The reasonfor this is that the Simplex was, by the nature of itsdesign, a 'square law' spray nozzle; that is, the flowthrough the nozzle is proportional to the square rootof the pressure drop across it. This meant that if theminimum pressure for effective atomization was 30lb. per square inch, the pressure needed to givemaximum flow would be about 3,000 lb. per squareinch. The fuel pumps available at that time wereunable to cope with such high pressures so thevariable port spray nozzle was developed in an effortto overcome the square law effect.

93. Although now only of historical value, thevariable port or Lubbock fuel spray nozzle (fig. 10-17) made use of a spring-loaded piston to control thearea of the inlet ports to the swirl chamber. At low fuelflows, the ports were partly uncovered by themovement of the piston; at high flows, they were fullyopen. By this method, the square law pressure rela-tionship was mainly overcome and good atomizationwas maintained over a wide range of fuel flows. Thematching of sets of spray nozzles and the sticking ofthe sliding piston due to dirt particles were, however,difficulties inherent in the design, and this type waseventually superseded by the Duplex and the Duplefuel spray nozzles.

94. The Duplex and the Duple spray nozzlesrequire a primary and a main fuel manifold and havetwo independent orifices, one much smaller than theother. The smaller orifice handles the lower flows andthe larger orifice deals with the higher flows as thefuel pressure increases. A pressurizing valve may beemployed with this type of spray nozzle to apportionthe fuel to the manifolds (fig. 10-18). As the fuel flow

and pressure increases, the pressurizing valvemoves to progressively admit fuel to the mainmanifold and the main orifices. This gives acombined flow down both manifolds. In this way, theDuplex and Duple nozzles are able to give effectiveatomization over a wider flow range than the Simplexspray nozzle for the same maximum fuel pressure.Also, efficient atomization is obtained at the low flowsthat may be required at high altitude. In the combinedacceleration and speed control system (para. 51),the fuel flow to the spray nozzles is apportioned inthe F.F.R.

95. The spill type fuel spray nozzle can bedescribed as being a Simplex spray nozzle with apassage from the swirl chamber for spilling fuelaway. With this arrangement it is possible to supplyfuel to the swirl chamber at a high pressure all thetime, As the fuel demand decreases with altitude orreduction in engine speed, more fuel is spilled awayfrom the swirl Chamber, leaving less to pass throughthe atomizing orifice. The spill spray nozzles'constant use of a relatively high pressure means thateven at the extremely low fuel flows that occur athigh altitude there is adequate swirl to provideconstant and efficient atomization of the fuel.

96. The spill spray nozzle system, however,involves a somewhat modified type of fuel supplyand control system from that used with the previoustypes. A means has to be provided for removing the

Fuel system


Fig. 10-16 A Simplex fuel spray nozzle.

Fig. 10-17 A variable port or Lubbock fuelspray nozzle.

spill and also for controlling the amount of spill flowat various engine operating conditions. Adisadvatage of this system is that excess heat maybe generated when a large volume of fuel is beingrecirculated to inlet. Such heat may eventually leadto a deterioration of the fuel.

97. The airspray nozzle (fig. 10-19), carries aproportion of the primary combustion air (Part 4) withthe injected fuel. By aerating the spray, the local fuel-rich concentrations produced by other types of spraynozzle are avoided, thus giving a reduction in bothcarbon formation and exhaust smoke. An additionaladvantage of the airspray nozzle is that the low

pressures required for atomization of the fuel permitsthe use of the comparatively lighter gear-type pump.

98. A flow distributor (fig. 10-20) is often required tocompensate for the gravity head across the manifoldat low fuel pressures to ensure that all spray nozzlespass equal quantities of fuel.

99. Some combustion systems vaporize the fuel(Part 4) as it enters the combustion zone.


100. On many engines, a fuel-cooled oil cooler(Part 8) is located between the L.P. fuel pump andthe inlet to the fuel filter (fig. 10-13), and advantageis taken of this to transfer the heat from the oil to thefuel and thus prevent blockage of the filter elementby ice particles. When heat transference by thismeans is insufficient, the fuel is passed through asecond heat exchanger where it absorbs heat from athermostatically controlled airflow taken from thecompressor.


101. The main effect on the engine of a changefrom one grade of fuel to another arises from thevariation of specific gravity and the number of heatunits obtainable from a gallon of fuel. As the numberof heat units per pound is practically the same for allfuels approved for gas turbine engines, a comparisonof heat values per gallon can be obtained bycomparing specific gravities.

102. Changes in specific gravity have a definiteeffect on the centrifugal pressure type of enginespeed governor (para. 15), for with an increase inspecific gravity the centrifugal pressure acting on thegovernor diaphragm is greater. Thus the speed atwhich the governor controls is reduced, and inconsequence the governor must be reset.

103. With a decrease in specific gravity, thecentrifugal pressure on the diaphragm is less and thespeed at which the governor controls is increased; inconsequence, the pilot must control the maximumr.p.m. by manual operation of the throttle to preventoverspeeding the engine until the governor can bereset. The hydro-mechanical governor (para. 23) isless sensitive to changes of specific gravity than thecentrifugal governor and is therefore preferred onmany fuel systems.

104. The pressure drop governor in the combinedacceleration and speed control system (para. 51) isdensity compensated, by using a buoyant material

Fuel system


Fig. 10-18 A Duple fuel spray nozzle andpressurizing valve.

for the governor weights, resulting in fuel beingmetered on mass flow rather than volume flow.

105. Changes to a lower grade of fuel can lead toproduction of carbon, giving a greater flameluminosity and temperature, leading to highercombustor metal temperatures and reducedcombustor and turbine life.


106. Fuels for aircraft gas turbine engines mustconform to strict requirements to give optimumengine performance, economy, safety and overhaullife. Fuels are classed under two headings, kerosine-type fuel and wide-cut gasoline-type fuel.

Fuel system


Fig. 10-19 An airspray nozzle.

Fig. 10-20 Fuel flow distributor.

Fuel requirements107. In general, a gas turbine fuel should have thefollowing qualities:

(1) Be 'pumpable' and flow easily under alloperating conditions.

(2) Permit engine starting at all groundconditions and give satisfactory flightrelighting characteristics.

(3) Give efficient combustion at all conditions.(4) Have as high a calorific value as possible.(5) Produce minimal harmful effects on the

combustion system or the turbine blades.(6) Produce minimal corrosive effects on the

fuel system components.(7) Provide adequate lubrication for the moving

parts of the fuel system.(8) Reduce fire hazards to a minimum.

108. The pumping qualities of the fuel depend uponits viscosity or thickness, which is related to fueltemperature, Fuel must be satisfactory down toapproximately -50 deg. C. As the fuel temperaturefalls, ice crystals may form to cause blockage of thefuel filter or the orifices in the fuel system. Fuelheating and anti-icing additives are available toalleviate this problem.

109. For easy starting, the gas turbine enginedepends upon the satisfactory ignition of theatomized spray of fuel from the fuel spray nozzles,assuming that the engine is being motored at therequired speed. Satisfactory ignition depends uponthe quality of fuel in two ways:

(1) The volatility of the fuel; that is, its ability tovaporize easily, especially at lowtemperatures.

(2) The degree of atomization, which dependsupon the viscosity of the fuel, the fuelpressure applied, and the design of theatomizer.

110. The calorific value (fig. 10-21) of a fuel is anexpression of the heat or energy content per poundor gallon that is released during combustion. Thisvalue, which is usually expressed in British thermalunits, influences the range of an aircraft. Where thelimiting factor is the capacity of the aircraft tanks, thecalorific value per unit volume should be as high aspossible, thus enabling more energy, and hencemore aircraft range, to be obtained from a givenvolume of fuel. When the useful payload is thelimiting factor, the calorific value per unit of weightshould be as high as possible, because more energycan then be obtained from a minimum weight of fuel.

Other factors which affect the choice of heat per unitof volume or weight, must also be taken into consid-eration; these include the type of aircraft, theduration of flight, and the required balance betweenfuel weight and payload.

Fuel system


111. Turbine fuels tend to corrode the componentsof the fuel and combustion systems mainly as a resultof the sulphur and water content of the fuel. Sulphur,when burnt in air, forms sulphur dioxide; when mixedwith water this forms sulphurous acid and is verycorrosive, particularly on copper and lead. Because itis impracticable to completely remove the sulphurcontent, it is essential that the sulphur be kept to acontrolled minimum. Although free water is removedprior to use, dissolved water, i.e. water in solution,cannot be effectively removed, as the fuel would re-absorb moisture from the atmosphere when stored ina vented aircraft or storage tank (para. 118).

112. All gas turbine fuels are potentially dangerousand therefore handling and storage precautionsshould be strictly observed.

Vapour locking and boiling113. The main physical difference between kerosineand wide-cut fuels is their degree of volatility, the lattertype of fuel having a higher volatility, thus increasingthe problem of vapour locking and boiling. Withkerosine-type fuels, the volatility is controlled by distil-lation and flash point, but with the wide-cut fuels it iscontrolled by distillation and the Reid VapourPressure (R.V.P.) test. In this test, the absolutepressure of the fuel is recorded by special apparatuswith the fuel temperature at 37.8 deg. C. (100 deg. F.).

114. Kerosine has a low vapour pressure and willboil only at extremely high altitudes or high tempera-

Fig. 10-21 Relationship between calorificvalue and specific gravity.

tures, whereas a wide-cut fuel wilt boil at a muchlower altitude.

115. The fuel temperature during flight dependsupon altitude, rate of climb, duration at altitude andkinetic heating due to forward speed. When boilingdoes occur, the vapour loss can be very high,especially with wide-cut fuels, and this may causevapour locking with consequent malfunctions of theengine fuel system and fuel metering equipment.

116. To obviate or reduce the risk of boiling, it isusual to pressurize the fuel tanks. This involvesmaintaining an absolute pressure above the fuel inexcess of its vapour pressure at any specifictemperature. This may be accomplished by using aninert gas or by using the fuel vapour pressure with acontrolled venting system.

117. For sustained supersonic flight, some measureof tank insulation is necessary to reduce kineticheating effects, even when lower volatility fuels areused.

Fuel contamination control118. Fuel can be maintained in good condition bywell planned storage and by making routine aircrafttank drain checks. The use of suitable filters,fuel/water separators and selected additives willrestrict the contamination level, e.g. free water andsolid matter, to a practical minimum. Keeping the fuelfree of undissolved water will prevent serious icingproblems, reduce the microbiological growth andminimize corrosion. Reducing the solid matter willprevent undue wear in the fuel pumps, reducecorrosion and lessen the possibility of blockageoccurring within the fuel system.

Fuel system


Rolls-Royce RB211-535C

Metrovick G2

Following the successful operation at sea ofthe Metrovick F2-based 2500 hp Gatric marinegas turbine, the Royal Navy ordered fourlarger sets with a maximum operational ratingof 4500 shp. Developed from the MetrovickF2/4 Beryl axial-flow aircraft engine; the G2swere installed in the Motor Gunboats 'BoldPioneer1 and 'Bold Pathfinder; the formergoing to sea in 1951.


1. Two separate systems are required to ensurethat a gas turbine engine will start satisfactorily.Firstly, provision must be made for the compressorand turbine to be rotated up to a speed at whichadequate air passes into the combustion system tomix with fuel from the fuel spray nozzles (Part 10).Secondly, provision must be made for ignition of theair/fuel mixture in the combustion system. Duringengine starting the two systems must operate simul-taneously, yet it must also be possible to motor theengine over without ignition for maintenance checksand to operate only the ignition system for relightingduring flight (para. 28).

2. The functioning of both systems is co-ordinatedduring a starting cycle and their operation is auto-matically controlled after the initiation of the cycle byan electrical circuit. A typical sequence of eventsduring the start of a turbo-jet engine is shown in fig.11-1.

11: Starting and ignition

Contents Page

Introduction 121Methods of starting 122

ElectricCartridgeIso-propyl-nitrateAirGas turbineHydraulic

Ignition 127Relighting 131


Fig. 11-1 A typical starting sequence of aturbo-jet engine.


3. The starting procedure for all jet engines isbasically the same, but can be achieved by variousmethods. The type and power source for the startervaries in accordance with engine and aircraft require-ments. Some use electrical power, others use gas,air or hydraulic pressure, and each has its ownmerits. For example, a military aircraft requires theengine to be started in the minimum time and, whenpossible, to be completely independent of externalequipment. A commercial aircraft, however, requiresthe engine to be started with the minimumdisturbance to the passengers and by the mosteconomical means. Whichever system is used,reliability is of prime importance.

4. The starter motor must produce a high torqueand transmit it to the engine rotating assembly in amanner that provides smooth acceleration from restup to a speed at which the gas flow through the

engine provides sufficient power for the engineturbine to take over.

Electric5. The electric starter is usually a direct current(D.C.) electric motor coupled to the engine through areduction gear and ratchet mechanism, or clutch,which automatically disengages after the engine hasreached a self-sustaining speed (fig. 11-2).

6. The electrical supply may be of a high or lowvoltage and is passed through a system of relays andresistances to allow the full voltage to be progres-sively built up as the starter gains speed. It alsoprovides the power for the operation of the ignitionsystem. The electrical supply is automaticallycancelled when the starter load is reduced after theengine has satisfactorily started or when the timecycle is completed. A typical electrical startingsystem is shown in fig. 11-3.

Starting and ignition


Fig. 11-2 An electric starter.

Starting and ignition


Fig. 11-3 A low voltage electrical starting system.

Cartridge7. Cartridge starting is sometimes used on militaryengines and provides a quick independent method ofstarting. The starter motor is basically a smallimpulse-type turbine that is driven by high velocitygases from a burning cartridge. The power output ofthe turbine is passed through a reduction gear andan automatic disconnect mechanism to rotate theengine. An electrically fired detonator initiates theburning of the cartridge charge. As a cordite chargeprovides the power supply for this type of starter, thesize of the charge required may well limit the use ofthe cartridge starters. A triple-breech starter isillustrated in fig. 11-4.

Iso-propyl-nitrate8. This type of starter provides a high power outputand gives rapid starting characteristics. It has aturbine that transmits power through a reduction gearto the engine. In this instance, the turbine is rotatedby high pressure gases resulting from thecombustion of iso-propyl-nitrate. This fuel is sprayedinto a combustion chamber, which forms part of thestarter, where it is electrically ignited by a high-energy ignition system. A pump supplies the fuel tothe combustion chamber from a storage tank and an

air pump scavenges the starter combustion chamberof fumes before each start. Operation of the fuel andair pumps, ignition systems, and cycle cancellation,is electrically controlled by relays and time switches.An iso-propyl-nitrate starting system is shown in fig.11-5.

Air9. Air starting is used on most commercial andsome military jet engines. It has many advantagesover other starting systems, and is comparativelylight, simple and economical to operate.

10. An air starter motor transmits power through areduction gear and clutch to the starter output shaftwhich is connected to the engine. A typical air startermotor is shown in fig. 11-6.

11. The starter turbine is rotated by air taken froman external ground supply, an auxiliary power unit(A.P.U.) or as a cross-feed from a running engine.The air supply to the starter is controlled by an elec-trically operated control and pressure reducing valvethat is opened when an engine start is selected andis automatically closed at a predetermined starterspeed. The clutch also automatically disengages asthe engine accelerates up to idling r.p.m. and the

Starting and ignition


Fig. 11-4 A triple-breech cartridge starter.

Starting and ignition


Fig. 11-5 An iso-propyl-nitrate starting system.

rotation of the starter ceases. A typical air startingsystem is shown in fig. 11-7.

12. A combustor starter is sometimes fitted to anengine incorporating an air starter and is used tosupply power to the starter when an external supplyof air is not available. The starter unit has a smallcombustion chamber into which high pressure air,from an aircraft-mounted storage bottle, and fuel,from the engine fuel system, are introduced. Controlvalves regulate the air supply which pressurizes afuel accumulator to give sufficient fuel pressure foratomization and also activates the continuousignition system. The fuel/air mixture is ignited in thecombustion chamber and the resultant gas isdirected onto the turbine of the air starter. Anelectrical circuit is provided to shut off the air supplywhich in turn terminates the fuel and ignition systemson completion of the starting cycle.

13. Some turbo-jet engines are not fitted with startermotors, but use air impingement onto the turbineblades as a means of rotating the engine. The air isobtained from an external source, or from an enginethat is running, and is directed through non-returnvalves and nozzles onto the turbine blades. A typicalmethod of air impingement starting is shown in fig.11-8.

Gas turbine14. A gas turbine starter is used for some jetengines and is completely self-contained. It has itsown fuel and ignition system, starting system (usuallyelectric or hydraulic) and self-contained oil system.This type of starter is economical to operate andprovides a high power output for a comparatively lowweight.

15. The starter consists of a small, compact gasturbine engine, usually featuring a turbine-drivencentrifugal compressor, a reverse flow combustionsystem and a mechanically independent |free-powerturbine. The free-power turbine is connected to themain engine via a two-stage epicyclic reduction gear,automatic clutch and output shaft. A typical gasturbine starter is shown in fig. 11-9.

16. On initiation of the starting cycle, the gas turbinestarter is rotated by its own starter motor until itreaches self-sustaining speed, when the starting andignition systems are automatically switched off.Acceleration then continues up to a controlled speedof approximately 60,000 r.p.m. At the same time asthe gas turbine starter engine is accelerating, theexhaust gas is being directed, via nozzle guidevanes, onto the free-power turbine to provide thedrive to the main engine. Once the main enginereaches self-sustaining speed, a cut-out switch

Starting and ignition


Fig. 11-6 An air starter motor.

operates and shuts down the gas turbine starter. Asthe starter runs down, the clutch automaticallydisengages from the output shaft and the mainengine accelerates up to idling r.p.m. under its ownpower.

Hydraulic17. Hydraulic starting is used for starling somesmall jet engines. In most applications, one of theengine-mounted hydraulic pumps is utilized and isknown as a pump/starter, although other applicationsmay use a separate hydraulic motor. Methods oftransmitting the torque to the engine may vary, but atypical system would include a reduction gear andclutch assembly. Power to rotate the pump/starter isprovided by hydraulic pressure from a ground supplyunit and is transmitted to the engine through thereduction gear and clutch. The starting system iscontrolled by an electrical circuit that also operateshydraulic valves so that on completion of the starting

cycle the pump /starter functions as a normalhydraulic pump.


18. High-energy (H.E.) ignition is used for startingall jet engines and a dual system is always fitted.Each system has an ignition unit connected to itsown igniter plug, the two plugs being situated indifferent positions in the combustion system.

19. Each H.E. ignition unit receives a low voltagesupply, controlled by the starting system electricalcircuit, from the aircraft electrical system. Theelectrical energy is stored in the unit until, at a pre-determined value, the energy is dissipated as a highvoltage, high amperage discharge across the igniterplug.

20. Ignition units are rated in 'joules' (one jouleequals one watt per second). They are designed to

Starting and ignition


Fig. 11-7 An air starting system.

give outputs which may vary according to require-ments. A high value output (e.g. twelve joule) isnecessary to ensure that the engine will obtain a sat-isfactory relight at high altitudes and is sometimesnecessary for starting. However, under certain flightconditions, such as icing or take-off in heavy rain orsnow, it may be necessary to have the ignitionsystem continuously operating to give an automaticrelight should flame extinction occur. For thiscondition, a low value output (e.g. three to six joule)is preferred because it results in a longer life of theigniter plug and ignition unit. Consequently, to suit allengine operating conditions, a combined systemgiving a high and low value output is favoured. Sucha system would consist of one unit emitting a highoutput to one igniter plug, and a second unit giving alow output to a second igniter plug. However, someignition units are capable o! supplying both high andlow outputs, the value being pre-selected asrequired.

Starting and ignition


Fig. 11-8 Air impingement starting.

Fig. 11-9 A gas turbine starter.

21. An ignition unit may be supplied with directcurrent (D.C.) and operated by a tremblermechanism or a transistor chopper circuit, orsupplied with alternating current (A.C.) and operatedby a transformer. The operation of each type of unitis described in the subsequent paragraphs.

22. The ignition unit shown in fig. 11-10 is atypicalD.C. trembler-operated unit. An induction coil,operated by the trembler mechanism, charges thereservoir capacitor (condenser) through a highvoltage rectifier. When the voltage in the capacitor isequal to the breakdown value of a sealed dischargegap, the energy is discharged across the face of theigniter plug. A choke is fitted to extend the duration ofthe discharge and a discharge resistor is fitted to

ensure that any residual stored energy in thecapacitor is dissipated within one minute of thesystem being switched off. A safety resistor is fitted toenable the unit to operate safely, even when the hightension lead is disconnected and isolated.

23. Operation of the transistorized ignition unit issimilar to that of the D.C. trembler-operated unit,except that the trembler-unit is replaced by atransistor chopper circuit. A typical transistorized unitis shown in fig. 11-11; such a unit has manyadvantages over the trembler-operated unit becauseit has no moving parts and gives a much longeroperating life. The size of the transistorized unit isreduced and its weight is less than that of thetrembler-operated unit.

Starting and ignition


Fig. 11-10 A D.C. trembler-operated ignition unit.

Starting and ignition


Fig. 11-11 A transistorized ignition unit.

Fig. 11-12 An A.C. ignition unit.

24. The A.C. ignition unit, shown in fig, 11-12,receives an alternating current which is passedthrough a transformer and rectifier to charge acapacitor. When the voltage in the capacitor is equalto the breakdown value of a sealed discharge gap,the capacitor discharges the energy across the faceof the igniter plug. Safety and discharge resistors arefitted as in the trembler-operated unit.

25. There are two basic types of igniter plug; theconstricted or constrained air gap type and theshunted surface discharge type. The air gap type issimilar in operation to the conventional reciprocatingengine spark plug, but has a larger air gap betweenthe electrode and body for the spark to cross. Apotential difference of approximately 25,000 volts isrequired to ionize the gap before a spark will occur.This high voltage requires very good insulationthroughout the circuit. The surface discharge igniterplug (fig. 11-13) has the end of the insulator formedby a semi-conducting pellet which permits anelectrical leakage from the central high tensionelectrode to the body. This ionizes the surface of the

pellet to provide a low resistance path for the energystored in the capacitor. The discharge takes the formof a high intensity flashover from the electrode to thebody and only requires a potential difference ofapproximately 2000 volts for operation.

26. The normal spark rate of a typical ignitionsystem is between 60 and 100 sparks per minute.Periodic replacement of the igniter plug is necessarydue to the progressive erosion of the igniterelectrodes caused by each discharge.

27. The igniter plug tip protrudes approximately 0.1inch into the flame tube. During operation the sparkpenetrates a further 0.75 inch. The fuel mixture isignited in the relatively stable boundary layer whichthen propagates throughout the combustion system.


28. The jet engine requires facilities for relightingshould the flame in the combustion system be extin-guished during flight. However, the ability of theengine to relight will vary according to the altitudeand forward speed of the aircraft. A typical relightenvelope, showing the flight conditions under whichan engine will obtain a satisfactory relight, is shownin fig. 11-14. Within the limits of the envelope, theairflow through the engine will rotate the compressorat a speed satisfactory for relighting; all that isrequired therefore, provided that a fuel supply isavailable, is the operation of the ignition system. Thisis provided for by a separate switch that operatesonly the ignition system.

Starting and ignition


Fig. 11-13 An igniter plug.

Fig. 11-14 A typical flight relight envelope.

Armstrong Siddeley Sapphire

The Sapphire originated in 1946 with theMetrovick F9, which was handed over toArmstrong-Siddeley when Metropolitan-Vickers withdrew from aviation in 1947. TheSapphire first ran in October 1948 and theengine was flight tested in Meteor, Hastingsand Canberra aircraft; before going intoproduction for the Gloster Javelin and HawkerHunter F2.

Rolls-Royce contra-rotating fan (concept)

Contents Page

Introduction 133Controls 133Instrumentation 135

Engine thrustEngine torqueEngine speedTurbine gas temperatureOil temperature and pressureFuel temperature and pressureFuel flowVibrationWarning systemsAircraft integrated data systemElectronic indicating systems

Synchronizing and synchrophasing 144


1. The controls of the gas turbine engine aredesigned to remove, as far as possible, work loadfrom the pilot while still allowing him ultimate controlof the engine. To achieve this, the fuel flow is auto-matically controlled after the pilot has made the initialpower selection (Part 10).

2. All engine parameters require monitoring andinstrumentation is provided to inform the pilot of thecorrect functioning of the various engine systemsand to warn of any impending failure. Should any ofthe automatic governors fail, the engine can be

manually controlled by the pilot selecting the desiredthrust setting and monitoring the instruments tomaintain the engine within the relevant operatinglimitations.

3. The multitude of dials and gauges on the pilot'sinstrument panel may be replaced by one or anumber of cathode ray tubes to display engineparameters. These are small screens capable ofdisplaying all of the information necessary to operatethe engine safely.


4. The control of a gas turbine engine generallyrequires the use of only one control lever and themonitoring of certain indicators located on the pilot'sinstrument panel (fig. 12-1). Operation of the control(throttle/power) lever selects a thrust level which isthen maintained automatically by the fuel system(Part 10).


12: Controls and instrumentation

Controls and instrumentation


Fig. 12-1 Pilot's instrument panel - turbo-jet engines.

5. On engines fitted with afterburning, single levercontrol is maintained, although a further fuel systemis required to supply and control the fuel to theafterburner (Part 16).

6. On a turbo-propeller engine, the throttle lever isinterconnected with the propeller control unit(P.C.U.), thus maintaining single lever operation ofthe engine. Similarly, the throttle control lever of ahelicopter is interconnected with the collective pitchlever, so ensuring that an increase in pitch isaccompanied by an increase in engine power,

7. The fuel system (Part 10) incorporates a highpressure fuel shut-off cock to provide a means ofstopping the engine. This may be operated by aseparate lever, interconnected with the throttle lever,or electrically actuated and controlled by a switch onthe pilot's instrument panel.

8. A turbo-jet engine fitted with a thrust reverserusually has an additional control lever that allowsreverse thrust to be selected (Part 15). On a turbo-propeller engine, a separate control lever is notrequired because the interconnected throttle andP.C.U. lever is operated to reverse the pitch of thepropeller.

INSTRUMENTATION9. The performance of the engine and the operationof the engine systems are shown on gauges or bythe operation of flag or dolls-eye type indicators. Adiagrammatic arrangement of the control and instru-mentation for a turbo-jet engine is shown in fig. 12-2.

Controls and instrumentation


Fig. 12-2 Diagrammatic arrangement of engine control and instrumentation.

Controls and instrumentation


Fig. 12-3 Electro-mechanical E.P.R. transmitter.

Engine thrust10. The thrust of an engine is shown on a thrust-meter, which will be one of two basic types; the firstmeasures turbine discharge or jet pipe pressure, andthe second, known as an engine pressure ratio(E.P.R.) gauge, measures the ratio of two or threeparameters. When E.P.R. is measured, the ratio isusually that of jet pipe pressure to compressor inletpressure. However, on a fan engine the ratio may bethat of integrated turbine discharge and fan outletpressures to compressor inlet pressure.

11. In each instance, an indication of thrust output isgiven, although when only the turbine dischargepressure is measured, correction is necessary forvariation of inlet pressure; however, both types mayrequire correction for variation of ambient airtemperature. To compensate for ambientatmospheric conditions, it is possible to set acorrection figure to a sub-scale on the gauge; thus,

the minimum thrust output can be checked under alloperating conditions.

12. Suitably positioned pilot tubes sense thepressure or pressures appropriate to the type ofindication being taken from the engine. The pilottubes are either directly connected to the indicator orto a pressure transmitter that is electricallyconnected to the indicator.

13. An indicator that shows only the turbinedischarge pressure is basically a gauge, the dial ofwhich may be marked in pounds per square inch(p.s.i.), inches of mercury (in. Hg.), or a percentageof the maximum thrust.

14. E.P.R. can be indicated by either electro-mechanical or electronic transmitters. In both casesthe inputs to the transmitter are engine inlet pressure(P1) and an integrated pressure (PINT) comprised of

Controls and instrumentation


Fig. 12-4 A simple torquemeter system.

fan outlet and turbine exhaust pressures. In somecases either fan outlet pressure or turbine exhaustpressure are used alone in place of PINT.

15. The electro-mechanical system indicates achange in pressure by using transducer capsules(fig. 12-3) to deflect the centre shaft of the pressuretransducer causing the yoke to pivot about the axisA.A. This movement is sensed by the linear variabledifferential transformer (L.V.D.T.) and converted to ana.c. electrical signal which is amplified and applied tothe control winding of the servo motor.

16. The servo motor, through the gears, alters thepotentiometer output voltage signal to the E.P.R.indicator and simultaneously drives the gimbal in thesame direction as the initial yoke movement until theL.V.D.T. signal to the motor is cancelled and thesystem stabilizes at the new setting.

17. The electronic E.P.R. system utilizes twovibrating cylinder pressure transducers which sensethe engine air pressures and vibrate at frequenciesrelative to these pressures. From these vibrationfrequencies electrical signals of E.P.R. are computedand are supplied to the E.P.R. gauge and electronicengine control system (Part 10).

Engine torque18. Engine torque is used to indicate the power thatis developed by a turbo-propeller engine, and theindicator is known as a torquemeter. The enginetorque or turning moment is proportional to thehorse-power and is transmitted through the propellerreduction gear.

19. A torquemeter system is shown in fig. 12-4. Inthis system, the axial thrust produced by the helicalgears is opposed by oil pressure acting on a numberof pistons; the pressure required to resist the axialthrust is transmitted to the indicator.

20. In addition to providing an indication of enginepower; the torquemeter system may also be used toautomatically operate the propeller featheringsystem if the torquemeter oil pressure falls due to apower failure. It is also used, on some installations,to assist in the automatic operation of the waterinjection system to restore or boost the take-offpower at high ambient temperatures or at highaltitude airports (Part 17).

Engine speed21. All engines have their rotational speed (r.p.m.)indicated. On a twin or triple-spool engine, the highpressure assembly speed is always indicated; inmost instances, additional indicators show the speed

of the low pressure and intermediate pressureassemblies.

22. Engine speed indication is electricallytransmitted from a small generator, driven by theengine, to an indicator that shows the actualrevolutions per minute (r.p.m.), or a percentage ofthe maximum engine speed (fig. 12-5). The enginespeed is often used to assess engine thrust, but it

Controls and instrumentation


Fig. 12-5 Engine speed indicators andgenerator.

does not give an absolute indication of the thrustbeing produced because inlet temperature andpressure conditions affect the thrust at a givenengine speed.

23. The engine speed generator supplies a three-phase alternating current, the frequency of which isdependent upon engine speed. The generator outputfrequency controls the speed of a synchronousmotor in the indicator, and rotation of a magnetassembly housed in a drum or drag cup inducesmovement of the drum and consequent movement ofthe indicator pointer,

24. Where there is no provision for driving agenerator, a variable-reluctance speed probe, inconjunction with a phonic wheel, may be used toinduce an electric current that is amplified and thentransmitted to an indicator (fig. 12-6). This methodcan be used to provide an indication of r.p.m. withoutthe need for a separately driven generator, with itsassociated drives, thus reducing the number ofcomponents and moving parts in the engine.

25. The speed probe is positioned on thecompressor casing in line with the phonic wheel,which is a machined part of the compressor shaft.The teeth on the periphery of the wheel pass theprobe once each revolution and induce an electriccurrent by varying the magnetic flux across a coil inthe probe. The magnitude of the current is governedby the rate of change of the magnetic flux and is thusdirectly related to engine speed.

26. In addition to providing an indication of rotorspeed, the current induced at the speed probe canbe used to illuminate a warning lamp on theinstrument panel to indicate to the pilot that a rotorassembly is turning. This is particularly important atengine start, because it informs the pilot when toopen the fuel cock to allow fuel to the engine. Thelamp is connected into the slatting circuit and isilluminated during the starting cycle.

Turbine gas temperature27. The temperature of the exhaust gases is alwaysindicated to ensure that the temperature of theturbine assembly can be checked at any specificoperating condition. In addition, an automatic gastemperature control system is usually provided, toensure that the maximum gas temperature is notexceeded (Part 10).

28. Turbine gas temperature (T.G.T.) sometimesreferred to as exhaust gas temperature (E.G.T.) or jetpipe temperature (J.P.T.), is a critical variable ofengine operation and it is essential to provide anindication of this temperature. Ideally, turbine entrytemperature (T.E.T.) should be measured; however,because of the high temperatures involved this is notpractical, but, as the temperature drop across theturbine varies in a known manner, the temperature atthe outlet from the turbine is usually measured bysuitably positioned thermocouples. The temperaturemay alternatively be measured at an intermediatestage of the turbine assembly, as shown in fig. 12-7.

29. The thermocouple probes used to transmit thetemperature signal to the indicator consist of twowires of dissimilar metals that are joined togetherinside a metal guard tube. Transfer holes in the tubeallow the exhaust gas to flow across the junction.The materials from which the thermocouples wiresare made are usually nickel-chromium and nickel-aluminium alloys.

30. The probes are positioned in the gas stream soas to obtain a good average temperature readingand are normally connected to form a parallel circuit.An indicator, which is basically a millivoltmetercalibrated to read in degrees centigrade, isconnected into the circuit (fig. 12-8).

31. The junction of the two wires at the thermocou-ple probe is known as the 'hot' or 'measuring' junctionand that at the indicator as the 'cold' or 'reference'junction. If the cold junction is at a constanttemperature and the hot junction is sensing theexhaust gas temperature, an electromotive force(E.M.F.), proportional to the temperature difference

Controls and instrumentation


Fig. 12-6 Variable-reluctance speed probeand phonic wheel.

of the two junctions is created in the circuit and thiscauses the indicator pointer to move. To preventvariations of cold junction temperature affecting theindicated temperature, an automatic temperaturecompensating device is incorporated in the indicatoror in the circuit.

32. The thermocouple probes may be of single,double or triple element construction. Where multipleprobes are used they are of differing lengths in orderto obtain a temperature reading from different pointsin the gas stream to provide a better average readingthan can be obtained from a single probe (fig. 12-7).

33. The output to the temperature control systemcan also be used to provide a signal, in the form ofshort pulses, which, when coupled to an indicator,will digitally record the life of the engine. Duringengine operation in the higher temperature ranges,the pulse frequency increases progressively causingthe cyclic-type indicator to record at a higher rate,thus relating engine or unit life directly to operatingtemperatures.

34. Thermocouples may also be positioned totransmit a signal of air intake temperature into theexhaust gas temperature indicating and controlsystems, thus giving a reading of gas temperaturethat is compensated for variations of intaketemperature. A typical double-element thermocouplesystem with air intake probes is shown in fig. 12-8.

Oil temperature and pressure35. It is essential for correct and safe operation ofthe engine that accurate indication is obtained ofboth the temperature and pressure of the oil.Temperature and pressure transmitters andindicators are illustrated in fig 12-9.

36. Oil temperature is sensed by a temperature-sensitive element fitted in the oil system. A change intemperature causes a change in the resistance valueand, consequently, a corresponding change in thecurrent flow at the indicator. The indicator pointer isdeflected by an amount equivalent to thetemperature change and this is recorded on thegauge in degrees centigrade.

37. Oil pressure is electrically transmitted to anindicator on the instrument panel. Some installationsuse a flag-type indicator, which indicates if thepressure is high, normal or low; others use a dial-type gauge calibrated in pounds per square inch(p.s.i.).

38. Electrical operation of each type is similar; oilpressure, acting on the transmitter, causes a changein the electric current supplied to the indicator. Theamount of change is proportional to the pressureapplied at the transmitter.

39. The transmitter may be of either the direct or thedifferential pressure type. The latter senses thepressure difference between engine feed and return

Controls and instrumentation


Fig. 12-7 Turbine thermocouple installation.

oil pressures, the return oil being pressurized bycooling and sealing air (Part 9) from the bearings.

40. In addition to a pressure gauge operated by atransmitter, an oil low pressure warning switch maybe provided to indicate that a minimum pressure isavailable for continued safe running of the engine.The switch is connected to a warning lamp in theflight compartment and the lamp illuminates if thepressure falls below an acceptable minimum.

Fuel temperature and pressure41. The temperature and pressure of the lowpressure fuel supply are electrically transmitted totheir respective indicators and these show if the lowpressure system is providing an adequate supply offuel without cavitation and at a temperature to suitthe operating conditions. The fuel temperature and

pressure indicators are similar to those fortemperature and pressure indication.

42. On some engines, a fuel differential pressureswitch, fitted to the low pressure fuel filter, senses thepressure difference across the filter element. Theswitch is connected to a warning lamp that providesindication of partial filter blockage, with the possibilityof fuel starvation.

Fuel flow43. Although the amount of fuel consumed during agiven flight may vary slightly between engines of thesame type, fuel flow does provide a useful indicationof the satisfactory operation of the engine and of theamount of fuel being consumed during the flight. Atypical system consists of a fuel flow transmitter,which is fitted into the low pressure fuel system, andan indicator, which shows the rate of fuel flow and the

Controls and instrumentation


Fig. 12-8 A typical double element thermocouple system.

total fuel used in gallons, pounds or kilogrammes perhour (fig. 12-10). The transmitter measures the fuelflow electrically and an associated electronic unitgives a signal to the indicator proportional to the fuelflow.

Vibration44. A turbo-jet engine has an extremely lowvibration level and a change of vibration, due to animpending or partial failure, may pass without beingnoticed. Many engines are therefore fitted withvibration indicators that continually monitor thevibration level of the engine. The indicator is usuallya milliammeter that receives signals through anamplifier from engine mounted transmitters (fig. 12-11).

45. A vibration transmitter is mounted on the enginecasing and electrically connected to the amplifier andindicator. The vibration sensing element is usually anelectro-magnetic transducer that converts the rate ofvibration into electrical signals and these cause theindicator pointer to move proportional to the vibrationlevel. A warning lamp on the instrument panel isincorporated in the system to warn the pilot if anunacceptable level of vibration is approached,enabling the engine to be shut down and so reducethe risk of damage.

Controls and instrumentation


Fig. 12-9 Oil temperature and pressure transmitters and indicators.

Fig. 12-10 Fuel flow transmitter andindicator.

46. The vibration level recorded on the gauge is thesum total of vibration felt at the pick-up. A moreaccurate method differentiates between thefrequency ranges of each rotating assembly and soenables the source of vibration to be isolated. This isparticularly important on multi-spool engines.

47. A crystal-type vibration transmitter, giving amore reliable indication of vibration, has beendeveloped for use on multi-spool engines. A systemof filters in the electrical circuit to the gauge makes itpossible to compare the vibration obtained against aknown frequency range and so locate the vibrationsource. A multiple-selector switch enables the pilot toselect a specific area to obtain a reading of the levelof vibration.

Warning systems48. Warning systems are provided to give anindication of a possible failure or the existence of adangerous condition, so that action can be taken tosafeguard the engine or aircraft. Although the varioussystems of an aircraft engine are designed whereverpossible to 'fail safe1, additional safety devices aresometimes fitted. Automatic propeller featheringshould a power loss occur, and automatic closing ofthe high pressure fuel shut-off cock should a turbine

shaft failure occur, are but two examples. On someengine types, the fuel system is fitted with a controlto enable the engine to be operated by manualthrottling should a main fuel system failure occur.

49. In addition to a fire warning system (Part 14), anumber of other audible or visual warning systemscan be fitted to a gas turbine engine. These may befor low oil or fuel pressure, excessive vibration oroverheating. Indication of these may be by warninglight, bell or horn. A flashing light is used to attract thepilot's attention to a central warning panel (C.W.P.)where the actual fault is indicated.

50. Other instruments and lights warn the pilot ofthe selected position of the thrust reverser, the fanreverser or the afterburner variable nozzle, whenapplicable. Gauges also inform the pilot of suchthings as hydraulic pressure and flow and generatoroutput, which are vital to the correct operation of theaircraft systems.

Aircraft integrated data system51. The aircraft integrated data system (A.I.D.S.) isan extension of the 'black box' aircraft accident datarecorder. By monitoring and recording various engineparameters, either manually or automatically, it ispossible to detect an incipient failure and thusprevent a hazardous situation arising.

52. Selected performance parameters may berecorded for trend analysis or fault detection (Part24). Existing instruments are used, whereverpossible, to provide the signals to a magnetic tape.Further instrumentation, recording air pressure frompoints throughout the engine, oil contamination, tankcontents and scavenge oil temperature, may beprovided as required for flight recording,

53. After each flight the magnetic tape is processedby computer and the results are analyzed. Anydeviation from the normal condition will enable a faultto be identified and the necessary remedial action tobe taken.

Electronic indicating systems54. Electronic indicating systems consolidateengine indications, systems monitoring, and crewalerting functions onto one or more cathode raytubes (C.R.T.'s) mounted in the instrument panel.The information is displayed on the screen in theform of dials with digital readout and warnings,cautions and advisory messages shown as text.

Controls and instrumentation


Fig. 12-11 Vibration transmitter andindicator.

55. Only those parameters required by the crew toset and monitor engine thrust are permanentlydisplayed on the screen. The system monitors theremaining parameters and displays them only if oneor more exceed safe limitations. The pilot can,however, override the system and elect to have allmain parameters in view at any time (fig. 12-12).

56. Warnings, cautions and advisory messages aredisplayed only when necessary and are colour codedto communicate the urgency of the fault to the flightcrew. Provision is made to record any event or out oftolerance parameter in a non-volatile memory forlater evaluation by ground maintenance crews.

57. Electronic indicating systems offer improvedflight operations by reducing the pilot workloadthrough automatic monitoring of engine operationand a centralized caution and warning system.Reduced flight deck clutter is another feature as the

multitude of instruments traditionally present arereplaced by the C.R.T.'s.


58. Synchronizing and synchrophasing systems aresometimes used on turbo-propeller engined aircraftto achieve a reduction of noise during flight.

59. On a multi-engined aircraft, a synchronizingsystem ensures the propeller speeds are all thesame. This is achieved by an electrical system thatcompares speed signals from engine-mountedgenerators. Out-of-balance signals, using oneengine as a master signal, are automaticallycorrected by electrically trimming the engine speedsuntil all signals are equal.

60. A synchrophasing system ensures that anygiven blade of an engine propeller is in the same

Controls and instrumentation


Fig. 12-12 Typical electronic indicating display.

relative position as the corresponding, blade of thepropeller on the master engine. This again is auto-matically achieved by very fine trimming of enginespeeds resulting from phase signals from the syn-chrophasing generators.

61. On turbo-jet engines, synchronization can beachieved in a similar manner to that used for a turbo-

propeller engine. On multi-spool engines, only onespool is synchronized. Manual trimming of engine orshaft speed can be done with the assistance of asynchroscope. This visually indicates, in comparisonwith a master engine, if the other engines are runningat exactly the same speed; the normal engine speedindicator is, of course, not sufficiently sensitive to usefor synchronizing.

Controls and instrumentation


Rolls-Royce advanced turbo-propeller

De Havilland H6 Gyron Junior

When a change in government fighter require-ments halted development of the 20,000 lbthrust H4 Gyron in 1955, de Havilland decidedto build a 0.45 scale version known as the H6Gyron Junior. First run in August 1955 it waslater used to power the Blackburn BuccaneerS1 at 7100 lb thrust and the stainless steelBristol 188 at 14,000 lb with afterburner.

13: Ice protection

Contents Page

Introduction 147Hot air system 149Electrical system 150


1. Icing of the engine and the leading edges of theintake duct can occur during flight through cloudscontaining supercooled water droplets or duringground operation in freezing fog. Protection againstice formation may be required since icing of theseregions can considerably restrict the airflow throughthe engine, causing a loss in performance andpossible malfunction of the engine. Additionally,damage may result from ice breaking away andbeing ingested into the engine or hitting the acousticmaterial lining the intake duct.

2. An ice protection system must effectively preventice formation within the operational requirements ofthe particular aircraft. The system must be reliable,easy to maintain, present no excessive weightpenalty and cause no serious loss in engineperformance when in operation.

3. Analyses are carried out to determine whetherice protection is required and, if so, the heat inputrequired to limit ice build up to acceptable levels. Fig.13-1 illustrates the areas of a turbo-fan enginetypically considered for ice protection.

4. There are two basic systems of ice protection;turbo-jet engines generally use a hot air supply (fig.13-2), and turbo-propeller engines use electricalpower or a combination of electrical power and hot


Ice protection


Fig. 13-1 Areas typically considered for ice protection.

Fig. 13-2 Hot air ice protection.

air. Protection may be supplemented by thecirculation of hot oil around the air intake as shown infig. 13-3. The hot air system is generally used toprevent the formation of ice and is known as an anti-icing system. The electrical power system is used tobreak up ice that has formed on surfaces and isknown as a de-icing system.


5. The hot air system provides surface heating ofthe engine and/or powerplant where ice is likely toform. The protection of rotor blades is rarelynecessary, because any ice accretions are dispersedby centrifugal action. If stators are fitted upstream ofthe first rotating compressor stage these may require

protection. If the nose cone rotates it may not needanti-icing if its shape, construction and rotationalcharacteristics are such that likely icing isacceptable.

6. The hot air for the anti-icing system is usuallytaken from the high pressure compressor stages. It isducted through pressure regulating valves, to theparts requiring anti-icing. Spent air from the nosecowl anti-icing system may be exhausted into thecompressor intake or vented overboard.

7. If the nose cone is anti-iced its hot air supply maybe independent or integral with that of the nose cowland compressor stators. For an independent system,the nose cone is usually anti-iced by a continuous

Ice protection


Fig. 13-3 Combination of hot air, oil and electrical ice protection.

unregulated supply of hot air via internal ducting fromthe compressor.

8. The pressure regulating valves are electricallyactuated by manual selection, or automatically bysignals from the aircraft ice detection system. Thevalves prevent excessive pressures being developedin the system, and act also as an economy device atthe higher engine speeds by limiting the air offtakefrom the compressor, thus preventing an excessiveloss in performance. The main valve may bemanually locked in a pre-selected position prior totake-off in the event of a valve malfunction, prior toreplacement.


9. The electrical system of ice protection isgenerally used for turbo-propeller engine installa-tions, as this form of protection is necessary for thepropellers. The surfaces that require electricalheating are the air intake cowling of the engine, the

propeller blades and spinner and, when applicable,the oil cooler air intake cowling.

10. Electrical heating pads are bonded to the outerskin of the cowlings. They consist of strip conductorssandwiched between layers of neoprene, or glasscloth impregnated with epoxy resin. To protect thepads against rain erosion, they are coated with aspecial, polyurethane-based paint. When the de-icing system is operating, some of the areas are con-tinuously heated to prevent an ice cap forming on theleading edges and also to limit the size of the ice thatforms on the areas that are intermittently heated (fig.13-4).

11. Electrical power is supplied by a generator and,to keep the size and weight of the generator to aminimum, the de-icing electrical loads are cycledbetween the engine, propeller and, sometimes, theairframe.

12. When the ice protection system is in operation,the continuously heated areas prevent any ice

Ice protection


Fig. 13-4 Electrical ice protection.

forming, but the intermittently heated areas allow iceto form, during their 'heat-off period. During the 'heat-on' period, adhesion of the ice is broken and it is thenremoved by aerodynamic forces.

13. The cycling time of the intermittently heatedelements is arranged to ensure that the engine canaccept the amount of ice that collects during the'heat-off' period and yet ensure that the 'heat-on1period is long enough to give adequate shedding,

without causing any run-back icing to occur behindthe heated areas.

14. A two-speed cycling system is often used toaccommodate the propeller and spinner require-ments; a 'fast' cycle at the high air temperatureswhen the water concentration is usually greater anda 'slow' cycle in the lower temperature range. Atypical cycling sequence chart is shown in fig, 13-5.

Ice protection


Fig. 13-5 Typical ice protection cyclic sequence.

Rolls-Royce RB211-524D4D

Bristol Proteus

Work began in September 1944 on the 4000e.h.p. Proteus turbo-prop originally intendedto power the Bristol Brabazon 2 andSaunders-Poe Princess. The Proteus first ranin January 1947 and was later used to powerthe Bristol Britannia at 4445 e.h.p. Adevelopment of this engine, the MarineProteus, is used to power various patrol boats,hovercraft and hydrofoils.

14: Fire protection

Contents Page

Introduction 153Prevention of engine fire ignition 153

External cooling and ventilationFire detection 154Fire containment 156Fire extinguishing 157Engine overheat detection 157


1. All gas turbine engines and their associatedinstallation systems incorporate features thatminimize the possibility of an engine fire. It isessential, however, that if a failure does take placeand results in a fire, there is provision for theimmediate detection and rapid extinction of the fire,and for the prevention of it spreading. The detectionand extinguishing systems must add as little weightto the installation as possible.


2. An engine/powerplant is designed to ensure thatthe prevention of engine fire ignition is achieved asfar as possible. In most instances a dual failure isnecessary before a fire can occur.

3. Most of the potential sources of flammable fluidsare isolated from the 'hot end' of the engine. Externalfuel and oil system components and their associatedpipes are usually located around the compressorcasings, in a 'cool' zone, and are separated by a

fireproof bulkhead from the combustion, turbine andjet pipe area, or 'hot' zone. The zones may beventilated, as described in para 8, to prevent theaccumulation of flammable vapours.

4. All pipes that carry fuel, oil or hydraulic fluid, aremade fire resistant/proof to comply with fireregulations, and all electrical components andconnections are made explosion-proof. Sparkingcaused by discharge of static electricity is preventedby bonding all aircraft and engine components. Thisgives electrical continuity between all thecomponents and makes them incapable of ignitingflammable vapour.

5. On some engines, tubes carrying flammablefluids in 'hot areas' of the engine are constructed witha double skin. Should a fracture of the main fluidcarrying tube occur the outer skin will contain anyleakage, so preventing any possible fire ignition.

6. The power plant cowlings are provided with anadequate drainage system to remove flammablefluids from the nacelle, bay, or pod, and all sealleakages from components are drained overboard ata position such that fluid cannot re-enter the pod andcreate a fire hazard.

7. Spontaneous ignition can be minimized onaircraft flying at high Mach numbers by ductingboundary layer bleed air around the engine.


However, if ignition should occur, this high velocity airstream may have to be shut off, otherwise it wouldincrease the flame intensity and reduce the effective-ness of the extinguishing system by rapid dispersalof the extinguishant.

External cooling and ventilation8. The engine bay or pod is usually cooled andventilated by atmospheric air being passed aroundthe engine and then vented overboard (fig. 14-1).Convection cooling during ground running may beprovided by using an internal cooling outlet vent asan ejector system. An important function of theairflow is to purge any flammable vapours from theengine compartment. By keeping the airflow minimal,the power plant drag is minimized and, as therequired quantity of fire extinguishant is in proportionto the zonal airflow, any fire outbreak would be of lowintensity.

9. On some engines a fireproof bulkhead is alsoprovided to separate the 'cool' area or zone of theengine, which contains the fuel, oil, hydraulic andelectrical systems, from the 'hot' area surroundingthe combustion, turbine and exhaust sections of theengine. Differential pressures can be created in thetwo zones by calibration of the inlet and outletapertures to prevent the spread of fire from the hotzone.

10. Fig. 14-2 shows a more complex cooling andventilation system used on a turbo-fan engine. Air isinduced from the intake duct and also delivered fromthe fan to provide multi-zone cooling, each zonehaving its own calibrated cooling flow.


11. The rapid detection of a fire is essential tominimize the fire period before engine shut-down drilland release of extinguishant is effected. It is alsoextremely important that a fire detection system willnot give a false fire warning resulting from shortcircuiting caused by chafing or the ingress ofmoisture in the case of electrically operated systemsand chafes of the capillary resulting in loss of thecontained gas in the case of the gas filled continuouselement sensing type,

12. A detection system may consist of a number ofstrategically located detector units, or be of thecontinuous element (gas filled or electrical) sensingtype that can be shaped and attached to pre-formedtubes. The sensing element can be routed acrossoutlet orifices, such as a zone extractor ventilationduct, to give early detection of a fire (fig. 14-3).

13. In the case of electrical systems the presence ofa fire is signalled by a change in the electrical char-acteristics of the detector circuit, according to thetype of detector, be it thermistor, thermocouple orelectrical continuous element. In these cases thechange in temperature creates the signal which,through an amplifier, operates the warning indicator.

14. Both the thermocouple and thermistor detectorshave properties making them ideally suited to thisapplication. The thermocouple comprises twodissimilar metals which are joined together to formtwo junctions. As the temperature difference betweenthe two junctions increases an E.M.F. is produced inthe circuit and it is this E.M.F. that triggers the fire

Fire protection


Fig. 14-1 A typical cooling and ventilation system.

warning displays. The thermistor consists of a semi-conductor material whose resistance changes as

temperature increases, with a corresponding changein the current flowing in the circuit. It is this change in

Fire protection


Fig. 14-2 Cooling and ventilation - turbo-fan engine.

the current that operates the warning indicators. Athermistor may be used as a single point detector oras a continuous element sensor.

15. Another form of continuous element sensortakes the form of a capacitor consisting of a tubecontaining a dielectric material with a conductorrunning through the centre. A voltage difference isapplied between the tube and the centre conductor.As the temperature increases then the properties ofthe dielectric change with a corresponding change inthe value of capacitance. This change of capacitanceis displayed as a fire warning.

16. The gas filled detector consists of stainlesssteel tubing filled with gas absorbent material and inthe event of a fire or overheat condition thetemperature rise will cause the core of the sensingloop to expel the absorbed active gas into the sealedtube causing a rapid increase in pressure. This buildup of pressure is sensed by the detector alarmswitch. Should the sensing loop become damagedcausing a loss of the pressurized gas, an integrityswitch will indicate a detection loop fault on theappropriate engine. Fire indication is given by awarning light and bell.

17. At high Mach numbers, the considerably highertemperature levels may be such as to render thethermistor or thermocouple fire detection systemunsatisfactory. Thermal detectors that sense either atemperature rise, or a rate of temperature rise, maytherefore prove most suitable.

18. Alternatives to the above types are surveillancedetectors that respond to light radiation from a fire.These may be made so sensitive that they respondonly to the ultra-violet and infra-red rays emitted froma kerosine fire.


19. An engine fire must be contained within thepower plant and not be allowed to spread to otherparts of the aircraft. The cowlings that surround theengine are usually made of aluminium alloys, whichwould be unable to contain a fire when the aircraft isstatic. During flight, however, the airflow around thecowlings provides sufficient cooling to render themfireproof. Fireproof bulkheads and any cowlings thatare not affected by a cooling airflow, and sections ofcowlings around certain outlets that may act as'flame-holders', are usually manufactured from steelor titanium.

Fire protection


Fig. 14-3 A continuous element fire detecting system.


20. Before a fire extinguishing system is operated,the engine must be stopped to reduce the dischargeof flammable fluids and air into the fire area. Anyvalves, such as the low pressure fuel cock, thatcontrol the flow of flammable fluid must be situatedoutside the 'hot' zone to prevent fire damagerendering them inoperative.

21. After a fire has been extinguished, no attemptmust be made to start the engine again as this wouldprobably re-establish the fluid leak and the ignitionsource that were the original causes of the fire.Furthermore, the extinguishing system may beexhausted.

22. The extinguishant that is used for engine fires isusually one of the Freon compounds. Pressurizedcontainers are provided for the extinguishant andthese are located outside the fire risk zone. When therelevant electrical circuit is manually operated, the

extinguishant is discharged from the containersthrough a series of perforated spray pipes or nozzlesinto the fire (fig. 14-4). The discharge must besufficient to give a predetermined concentration ofextinguishant for a period that may vary between 0.5seconds and 2 seconds. The system is generally onethat enables two separate discharges to be made.


23. Turbine overheat does not constitute a seriousfire risk. Detection of an overheat condition, however,is essential to enable the pilot to stop the enginebefore mechanical or material damage results.

24. A warning system of a similar type to the firedetection system, or thermocouples suitablypositioned in the cooling airflow, may be used todetect excessive temperatures. Thermal switchespositioned in the engine overboard air vents, such asthe cooling air outlets, may also be included to givean additional warning.

Fire protection


Fig. 14-4 A typical fire extinguishing system.

Rolls - Royce Gem 2

Armstrong Siddeley Python

The Python was developed from the ASXaxial-flow turbo-jet which first ran in April 1943and was producing 2800 lb thrust by 1944.With the addition of a propeller gearbox theengine produced 3600 shp plus 1100 lb thrustand was known as the ASP. Renamed thePython it entered service as the power plantfor the Westland Wyvern S4 turbo-propfighter.

15: Thrust reversal

Contents Page

Introduction 159Principles of operation 160

Clamshell door systemBucket target systemCold stream reverser systemTurbo-propeller reverse pitch system

Construction and materials 166


1. Modern aircraft brakes are very efficient but onwet, icy or snow covered runways this efficiency maybe reduced by the loss of adhesion between theaircraft tyre and the runway thus creating a need foran additional method of bringing the aircraft to restwithin the required distance.

2. A simple and effective way to reduce the aircraftlanding run on both dry and slippery runways is toreverse the direction of the exhaust gas stream, thus

using engine power as a deceleration force. Thrustreversal has been used to reduce airspeed in flightbut it is not commonly used on modern aircraft. Thedifference in landing distances between an aircraftwithout reverse thrust and one using reverse thrust isillustrated in fig. 15-1.

3. On high by-pass ratio (fan) engines, reversethrust action is achieved by reversing the fan (coldstream) airflow. It is not necessary to reverse theexhaust gas flow (hot stream) as the majority of theengine thrust is derived from the fan.


4. On propeller-powered aircraft, reverse thrustaction is obtained by changing the pitch of thepropeller blades. This is usually achieved by a hydro-mechanical system, which changes the blade angleto give the braking action under the response of thepower or throttle lever in the aircraft.

5. Ideally, the gas should be directed in acompletely forward direction. It is not possible,however, to achieve this, mainly for aerodynamicreasons, and a discharge angle of approximately 45degrees is chosen. Therefore, the effective power inreverse thrust is proportionately less than the powerin forward thrust for the same throttle angle.


6. There are several methods of obtaining reversethrust on turbo-jet engines; three of these are shownin fig. 15-2 and explained in the followingparagraphs.

7. One method uses clamshell-type deflector doorsto reverse the exhaust gas stream and a seconduses a target system with external type doors to dothe same thing. The third method used on fanengines utilizes blocker doors to reverse the coldstream airflow.

8. Methods of reverse thrust selection and thesafety features incorporated in each systemdescribed are basically the same. A reverse thrustlever in the crew compartment is used to selectreverse thrust; the lever cannot be moved to thereverse thrust position unless the engine is runningat a low power setting, and the engine cannot beopened up to a high power setting if the reverser failsto move into the full reverse thrust position. Shouldthe operating pressure fall or fail, a mechanical lockholds the reverser in the forward thrust position; thislock cannot be removed until the pressure isrestored. Operation of the thrust reverser system isindicated in the crew compartment by a series oflights.

Clamshell door system9. The clamshell door system is a pneumaticallyoperated system, as shown in detail in fig. 15-3.Normal engine operation is not affected by thesystem, because the ducts through which theexhaust gases are deflected remain closed by thedoors until reverse thrust is selected by the pilot.

10. On the selection of reverse thrust, the doorsrotate to uncover the ducts and close the normal gasstream exit. Cascade vanes then direct the gasstream in a forward direction so that the jet thrustopposes the aircraft motion.

Thrust reversal


Fig. 15-1 Comparative landing runs with and without thrust reversal.

Thrust reversal


Fig. 15-2 Methods of thrust reversal.

Thrust reversal


Fig. 15-3 A typical thrust reverser system using clamshell doors.

Thrust reversal


Fig. 15-4 A typical fan cold stream thrust reversal system.

11. The clamshell doors are operated by pneumaticrams through levers that give the maximum load tothe doors in the forward thrust position; this ensureseffective sealing at the door edges, so preventinggas leakage. The door bearings and operatinglinkage operate without lubrication at temperatures ofup to 600 deg. C.

Bucket target system12. The bucket target system is hydraulicallyactuated and uses bucket-type doors to reverse thehot gas stream. The thrust reverser doors areactuated by means of a conventional pushrodsystem. A single hydraulic powered actuator isconnected to a drive idler, actuating the doorsthrough a pair of pushrods (one for each door).

13. The reverser doors are kept in synchronizationthrough the drive idler. The hydraulic actuator incor-porates a mechanical lock in the stowed (actuatorextended) position.

14. In the forward thrust mode (stowed) the thrustreverser doors form the convergent-divergent finalnozzle for the engine.

Cold stream reverser system15. The cold stream reverser system (fig. 15-4) canbe actuated by an air motor, the output of which isconverted to mechanical movement by a series offlexible drives, gearboxes and screwjacks, or by asystem incorporating hydraulic rams.

16. When the engine is operating in forward thrust,the cold stream final nozzle is 'open' because thecascade vanes are internally covered by the blockerdoors (flaps) and externally by the movable(translating) cowl; the latter item also serves toreduce drag.

17. On selection of reverse thrust, the actuationsystem moves the translating cowl rearwards and atthe same time folds the blocker doors to blank off thecold stream final nozzle, thus diverting the airflowthrough the cascade vanes.

Turbo-propeller reverse pitch system18. As mentioned in para. A, reverse thrust action isaffected on turbo-propeller powered aircraft bychanging the pitch of the propeller blades through ahydro-mechanical pitch control system (fig. 15-5).Movement of the throttle or power control lever

Thrust reversal


Fig. 15-5 A propeller pitch control system.

Thrust reversal


Fig. 15-6 Hot stream thrust reverser installations.

directs oil from the control system to the propellermechanism to reduce the blade angle to zero, andthen through to negative (reverse) pitch. Duringthrottle lever movement, the fuel to the engine istrimmed by the throttle valve, which is interconnect-ed to the pitch control unit, so that engine power andblade angle are co-ordinated to obtain the desiredamount of reverse thrust. Reverse thrust action mayalso be used to manoeuvre a turbo-propeller aircraftbackwards after it has been brought to rest.

19. Several safety factors are incorporated in thepropeller control system for use in the event ofpropeller malfunction, and these devices are usuallyhydro-mechanical pitch locking devices or stops.


20. The clamshell and bucket target doors (fig. 15-6) described in paras. 9 and 12 form part of the jet

pipe. The reverser casing is connected to the aircraftstructure or directly to the engine. The casingsupports the two reverser doors, the operatingmechanism and, in the case of the clamshell doorsystem, the outlet ducts that contain the cascadevanes. The angle and area of the gas stream arecontrolled by the number of vanes in each outletduct.

21. The clamshell and bucket target doors lie flushwith the casing during forward thrust operation andare hinged along the centre line of the jet pipe. Theyare, therefore, in line with the main gas load and thisensures that the minimum force is required to movethe doors.

22. Both the clamshell door system and the buckettarget system are subjected to high temperaturesand to high gas loads. The components of bothsystems, especially the doors, are therefore

Thrust reversal


Fig. 15-7 A cold stream thrust reverser installation.

constructed from heat-resisting materials and are ofparticularly robust construction.

23. The cold stream thrust reverser casing (fig. 15-7) is fitted between the low pressure compressorcasing and the cold stream final nozzle. Cascadevane assemblies are arranged in segments aroundthe circumference of the thrust reverser casing.Blocker doors are internally mounted and are

connected by linkages to the external movable(translating) cowl, which is mounted on rollers andtracks. Because the thrust reverser is not subjectedto high temperatures, the casing, blocker doors andcowl are constructed mainly of aluminium alloys orcomposite materials. The cowl is double-skinned,with the space between the skins containing noiseabsorbent material (Part 19).

Thrust reversal


Turbo-Union RB199

Metrovick F2/4 Beryl

Development of the F2, the first British axialflow turbo-jet, began in f 940. After initial flighttrials in the tail of an Avro Lancaster, two F2swere installed in a Gloster Meteor and firstflew on 13 November 1943. After earlyproblems the F2/4 Beryl was developed whichgave up to 4000 lb thrust and was used topower the Saunders Roe SR/A1 flying boatfighter.

16: Afterburning

Contents Page

Introduction 169Operation of afterburning 170Construction 173

Burners Jet pipe Propelling nozzle

Control system 173Thrust increase 175Fuel consumption 178


1. Afterburning (or reheat) is a method ofaugmenting the basic thrust of an engine to improvethe aircraft take-off, climb and (for military aircraft)combat performance. The increased power could beobtained by the use of a larger engine, but as thiswould increase the weight, frontal area and overallfuel consumption, afterburning provides the bestmethod of thrust augmentation for short periods.

2. Afterburning consists of the introduction andburning of fuel between the engine turbine and the jetpipe propelling nozzle, utilizing the unburned oxygenin the exhaust gas to support combustion (fig. 16-1).The resultant increase in the temperature of theexhaust gas gives an increased velocity of the jetleaving the propelling nozzle and therefore increasesthe engine thrust.

3. As the temperature of the afterburner flame canbe in excess of 1,700 deg. C., the burners are usuallyarranged so that the flame is concentrated aroundthe axis of the jet pipe. This allows a proportion of theturbine discharge gas to flow along the wall of the jetpipe and thus maintain the wall temperature at a safevalue.


4. The area of the afterburning jet pipe is larger thana normal jet pipe would be for the same engine, toobtain a reduced velocity gas stream. To provide foroperation under all conditions, an afterburning jetpipe is fitted with either a two-position or a variable-area propelling nozzle (fig. 16-2). The nozzle isclosed during non-afterburning operation, but whenafterburning is selected the gas temperatureincreases and the nozzle opens to give an exit areasuitable for the resultant increase in the volume ofthe gas stream. This prevents any increase inpressure occurring in the jet pipe which would affectthe functioning of the engine and enables afterburn-ing to be used over a wide range of engine speeds.

5. The thrust of an afterburning engine, withoutafterburning in operation, is slightly less than that ofa similar engine not fitted with afterburningequipment; this is due to the added restrictions in thejet pipe. The overall weight of the power plant is alsoincreased because of the heavier jet pipe and after-burning equipment.

6. Afterburning is achieved on low by-pass enginesby mixing the by-pass and turbine streams before theafterburner fuel injection and stabilizer system isreached so that the combustion takes place in the

mixed exhaust stream. An alternative method is toinject the fuel and stabilize the flame in the individualby-pass and turbine streams, burning the availablegases up to a common exit temperature at the finalnozzle. In this method, the fuel injection is scheduledseparately to the individual streams and it is normalto provide some form of interconnection between theflame stabilizers in the hot and cold streams to assistthe combustion processes in the cold by-pass air.


7. The gas stream from the engine turbine entersthe jet pipe at a velocity of 750 to 1,200 feet persecond, but as this velocity is far too high for a stableflame to be maintained, the flow is diffused before itenters the afterburner combustion zone, i.e. the flowvelocity is reduced and the pressure is increased.However, as the speed of burning kerosine at normalmixture ratios is only a few feet per second, any fuellit even in the diffused air stream would be blownaway. A form of flame stabilizer (vapour gutter) is,therefore, located downstream of the fuel burners toprovide a region in which turbulent eddies are formedto assist combustion and where the local gas velocityis further reduced to a figure at which flame stabi-lization occurs whilst combustion is in operation.


Fig. 16-1 Principle of afterburning


Fig. 16-2 Examples of afterburning jet pipes and propelling nozzles.

8. An atomized fuel spray is fed into the jet pipethrough a number of burners, which are so arrangedas to distribute the fuel evenly over the flame area.Combustion is then initiated by a catalytic igniter,which creates a flame as a result of the chemicalreaction of the fuel/air mixture being sprayed on to aplatinum-based element, by an igniter plug adjacentto the burner, or by a hot streak of flame thatoriginates in the engine combustion chamber (fig.

16-3): this latter method is known as 'hot-shot'ignition. Once combustion is initiated, the gastemperature increases and the expanding gasesaccelerate through the enlarged area propellingnozzle to provide the additional thrust.

9. In view of the high temperature of the gasesentering the jet pipe from the turbine, it might beassumed that the mixture would ignite spontaneous-ly. This is not so, for although cool flames form at


Fig. 16-3 Methods of afterburning ignition.

temperatures up to 700 deg. C., combustion will nottake place below 800 deg. C. If however, theconditions were such that spontaneous ignition couldbe effected at sea level, it is unlikely that it could beeffected at altitude where the atmospheric pressureis low. The spark or flame that initiates combustionmust be of such intensity that a light-up can beobtained at considerable altitudes.

10. For smooth functioning of the system, a stableflame that will burn steadily over a wide range ofmixture strengths and gas flows is required. Themixture must also be easy to ignite under allconditions of flight and combustion must bemaintained with the minimum loss of pressure.


Burners11. The burner system consists of several circularconcentric fuel manifolds supported by struts insidethe jet pipe. Fuel is supplied to the manifolds by feedpipes in the support struts and sprayed into the flamearea, between the flame stabilizers, from holes in thedownstream edge of the manifolds. The flamestabilizers are blunt nosed V-section annular ringslocated downstream of the fuel burners. Analternative system includes an additional segmentedfuel manifold mounted within the flame stabilizers.The typical burner and flame stabilizer shown in fig.16-4 is based on the latter system.

Jet pipe12. The afterburning jet pipe is made from a heat-resistant nickel alloy and requires more insulationthan the normal jet pipe to prevent the heat ofcombustion being transferred to the aircraft structure.The jet pipe may be of a double skin constructionwith the outer skin carrying the flight loads and theinner skin the thermal stresses; a flow of cooling airis often induced between the inner and outer skins.Provision is also made to accommodate expansionand contraction, and to prevent gas leaks at the jetpipe joints.

13. A circular heatshield of similar material to the jetpipe is often fitted to the inner wall of the jet pipe toimprove cooling at the rear of the burner section. Theheatshield comprises a number of bands, linked by

cooling corrugations, to form a single skin. The rearof the heatshield is a series of overlapping 'tiles'riveted to the surrounding skin (fig. 16-4). The shieldalso prevents combustion instability from creatingexcessive noise and vibration, which in turn wouldcause rapid physical deterioration of the afterburnerequipment.

Propelling nozzle14. The propelling nozzle is of similar material andconstruction as the jet pipe, to which it is secured asa separate assembly. A two-position propellingnozzle has two movable eyelids that are operated byactuators, or pneumatic rams, to give an open orclosed position (para. 4.). A variable-area propellingnozzle has a ring of interlocking flaps that are hingedto the outer casing and may be enclosed by an outershroud. The flaps are actuated by powered rams tothe closed position, and by gas loads to the interme-diate or the open positions; control of the flapposition is by a control unit and a pump provides thepower to the rams (para. 18).


15. It is apparent that two functions, fuel flow andpropelling nozzle area, must be co-ordinated for sat-isfactory operation of the afterburner system, Thesefunctions are related by making the nozzle areadependent upon the fuel flow at the burners or vice-versa. The pilot controls the afterburner fuel flow orthe nozzle area in conjunction with a compressordelivery/jet pipe pressure sensing device (a pressureratio control unit). When the afterburner fuel flow isincreased, the nozzle area increases; when theafterburner fuel flow decreases, the nozzle area isreduced. The pressure ratio control unit ensures thepressure ratio across the turbine remains unchangedand that the engine is unaffected by the operation ofafterburning, regardless of the nozzle area and fuelflow.

16. Since large fuel flows are required for afterburn-ing, an additional fuel pump is used. This pump isusually of the centrifugal flow or gear type and isenergized automatically when afterburning isselected. The system is fully automatic and incorpo-rates 'fail safe' features in the event of an afterburnermalfunction. The interconnection between the controlsystem and afterburner jet pipe is shown diagram-matically in fig. 16-5.


17. When afterburning is selected, a signal isrelayed to the afterburner fuel control unit. The unitdetermines the total fuel delivery of the pump andcontrols the distribution of fuel flow to the burnerassembly. Fuel from the burners is ignited, resultingin an increase in jet pipe pressure (P6). This altersthe pressure ratio across the turbine (P3/P6), and theexit area of the jet pipe nozzle is automaticallyincreased until the correct PS/PS ratio has beenrestored. With a further increase in the degree of

afterburning, the nozzle area is progressivelyincreased to maintain a satisfactory P3/P6 ratio. Fig.16-6 illustrates a typical afterburner fuel controlsystem.

18. To operate the propelling nozzle against thelarge 'drag' loads imposed by the gas stream, apump and either hydraulically or pneumaticallyoperated rams are incorporated in the controlsystem. The system shown in fig. 16-7 uses oil as the


Fig. 16-4 Typical afterburning jet pipe equipment.

hydraulic medium, but some systems use fuel.Nozzle movement is achieved by the hydraulicoperating rams which are pressurized by an oilpump, pump output being controlled by a linkagefrom the pressure ratio control unit. When anincrease in afterburning is selected, the afterburnerfuel control unit schedules an increase in fuel pumpoutput. The jet pipe pressure (P6) increases, alteringthe pressure ratio across the turbine (P3/P6). Thepressure ratio control unit alters oil pump output,causing an out-of-balance condition between thehydraulic ram load and the gas load on the nozzleflaps. The gas load opens the nozzle to increase itsexit area and, as the nozzle opens, the increase in

nozzle area restores the P3/P6 ratio and thepressure ratio control unit alters oil pump output untilbalance is restored between the hydraulic rams andthe gas loading on the nozzle flaps.


19. The increase in thrust due to afterburningdepends solely upon the ratio of the absolute jet pipetemperatures before and after the extra fuel is burnt.For example, neglecting small losses due to theafterburner equipment and gas flow momentumchanges, the thrust increase may be calculated asfollows.


Fig. 16-5 Simplified control system.


Fig. 16-6 A simplified typical afterburner fuel control system.

20. Assuming a gas temperature before afterburn-ing of 640 deg. C. (913 deg. K.) and with afterburn-ing of 1,269 deg. C. (1,542 deg. K.). then thetemperature ratio = 1,542 = 1.69. 913

The velocity of the jet stream increases as thesquare root of the temperature ratio. Therefore, thejet velocity = ^/T.69 = 1.3. Thus, the jet streamvelocity is increased by 30 per cent, and the increasein static thrust, in this instance, is also 30 per cent(fig. 16-8).

21. Static thrust increases of up to 70 per cent areobtainable from low by-pass engines fitted with after-burning equipment and at high forward speedsseveral times this amount of thrust boost can beobtained. High thrust boosts can be achieved on lowby-pass engines because of the large amount ofoxygen in the exhaust gas stream and the low initialtemperature of the exhaust gases.


Fig. 16-8 Thrust increase and temperatureratio.

Fig. 16-7 A simplified typical afterburner nozzle control system.

22. It is not possible to go on increasing the amountof fuel that is burnt in the jet pipe so that all theavailable oxygen is used, because the jet pipe wouldnot withstand the high temperatures that would beincurred and complete combustion cannot beassured.


23. Afterburning always incurs an increase inspecific fuel consumption and is, therefore, generallylimited to periods of short duration. Additional fuelmust be added to the gas stream to obtain therequired temperature ratio (para. 19). Since thetemperature rise does not occur at the peak ofcompression, the fuel is not burnt as efficiently as inthe engine combustion chamber and a higherspecific fuel consumption must result. For example,assuming a specific fuel consumption without after-burning of 1,15 lb./hr./lb. thrust at sea level and aspeed of Mach 0,9 as shown in fig. 16-9. then with70 per cent afterburning under the same conditionsof flight, the consumption will be increased to


Fig. 16-9 Specific fuel consumptioncomparison.

Fig. 16-10 Afterburning and its effect on the rate of climb.

approximately 2.53 lb./hr./lb. thrust. With an increasein height to 35,000 feet this latter figure of 2.53lb./hr./lb. thrust will fall slightly to about 2.34 lb./hr./lb.thrust due to the reduced intake temperature. When

this additional fuel consumption is combined with theimproved rate of take-off and climb (fig. 16-10), it isfound that the amount of fuel required to reduce thetime taken to reach operation height is not excessive.


Rolls-Royce Dart

Armstrong Siddeley Viper

The Viper was designed as a result ofexperience gained with the larger Sapphireturbojet. Originally built as a 1,640 lb thrustshort-life engine for target drones, it lateremerged as a long life engine for the JetProvost. Subsequently the engine wasdeveloped by Bristol Siddeley as the powerplant for civil executive jets, and Rolls-Roycefor present generation trainers and light strikeaircraft with a maximum thrust of 4,400 lb(5,000 lb with reheat).

17: Water injection

Contents Page

Introduction 181 Compressor inlet injection 183Combustion chamber injection 184


1. The maximum power output of a gas turbineengine depends to a large extent upon the density orweight of the airflow passing through the engine.There is, therefore, a reduction in thrust or shafthorsepower as the atmospheric pressure decreaseswith altitude, and/or the ambient air temperatureincreases. Under these conditions, the power outputcan be restored or, in some instances, boosted fortake-off by cooling the airflow with water or

water/methanol mixture (coolant). When methanol isadded to the water it gives anti-freezing propertiesand also provides an additional source of fuel. Atypical turbo-jet engine thrust restoration curve isshown in fig. 17-1 and a turbo-propeller enginepower restoration and boost curve is shown in fig.17-2.

2. There are two basic methods of injecting thecoolant into the airflow. Some engines have thecoolant sprayed directly into the compressor inlet,but the injection of coolant into the combustionchamber inlet is usually more suitable for axial flowcompressor engines. This is because a more evendistribution can be obtained and a greater quantity ofcoolant can be satisfactorily injected.

3. When water/methanol mixture is sprayed into thecompressor inlet, the temperature of the compressor


Water injection


Fig. 17-1 Turbo-jet thrust restoration.

Fig. 17-2 Turbo-propeller power boost.

inlet air is reduced and consequently the air densityand thrust are increased. If water only was injected,it would reduce the turbine inlet temperature, but withthe addition of methanol the turbine inlet temperatureis restored by the burning of methanol in thecombustion chamber. Thus the power is restoredwithout having to adjust the fuel flow.

4. The injection of coolant into the combustionchamber inlet increases the mass flow through theturbine, relative to that through the compressor. Thepressure and temperature drop across the turbine isthus reduced, and this results in an increased jet pipepressure, which in turn gives additional thrust. Theconsequent reduction in turbine inlet temperature,due to water injection, enables the fuel system toschedule an increase of fuel flow to a value that gives

an increase in the maximum rotational speed of theengine, thus providing further additional thrust,Where methanol is used with the water, the turbineinlet temperature is restored, or partially restored, bythe burning of the methanol in the combustionchamber.


5. The compressor inlet injection system shown infig. 17-3 is a typical system for a turbo-propellerengine. When the injection system is switched on,water/methanol mixture is pumped from an aircraft-mounted tank to a control unit. The control unitmeters the flow of mixture to the compressor inletthrough a metering valve that is operated by a servopiston. The servo system uses engine oil as anoperating medium, and a servo valve regulates the

Water injection


Fig. 17-3 A typical compressor inlet injection system.

supply of oil. The degree of servo valve opening isset by a control system that is sensitive to propellershaft torque oil pressure and to atmospheric airpressure acting on a capsule assembly.

6. The control unit high pressure oil cock controllever is interconnected to the throttle control systemin such a manner that, until the throttle is movedtowards the take-off position, the oil cock remainsclosed, and thus the metering valve remains closed,preventing any mixture flowing to the compressor

inlet Movement of the throttle control to the take-offposition opens the oil cock, and the oil pressurepasses through the servo valve to open the meteringvalve by means of the servo piston.


7. The combustion chamber injection systemshown in fig. 17-4 is a typical system for a turbo-jetengine. The coolant flows from an aircraft-mountedtank to an air-driven turbine pump that delivers it to awater flow sensing unit. The water passes from the

Water injection


Fig. 17-4 A typical combustion chamber injection system.

sensing unit to each fuel spray nozzle and is sprayedfrom two jets onto the flame tube swirl vanes, thuscooling the air passing into the combustion zone. Thewater pressure between the sensing unit and thedischarge jets is sensed by the fuel control system,which automatically resets the engine speedgovernor to give a higher maximum engine speed.

8. The water flow sensing unit opens only when thecorrect pressure difference is obtained between

compressor delivery air pressure and waterpressure. The system is brought into operation whenthe engine throttle lever is moved to the take-offposition, causing microswitches to operate andselect the air supply for the turbine pump.

9. The sensing unit also forms a non-return valve toprevent air pressure feeding back from the dischargejets and provides for the operation of an indicatorlight to show when water is flowing.

Water injection


Rolls-Royce Pegasus

Rolls-Royce RB 108

The RB108 was the first engine to bedesigned specifically as a direct VTOL engine.First running in July 1955 the engine was sub-sequently thrust rated at 2340 lb, giving athrust to weight ratio of 8.7:1. In addition topowering a variety of VTOL test rigs, theRB108 flew in a Gloster Meteor, the ShortSC1 and the Marcel Dassault Balzac.

18: Vertical/short take-offand landing

Contents Page

Introduction 187Methods of providing powered lift 189

Lift/propulsion engines Lift engines Remote lift systems Swivelling engines Bleed air for STOL

Lift thrust augmentation 194Special engine ratings Lift burning systems Ejectors

Aircraft control 197Reaction controls Differential engine throttling Automatic control systems


1. Vertical take-off and landing (VTOL) or shorttake-off and landing (STOL) are desirable character-istics for any type of aircraft, provided that the normalflight performance characteristics, includingpayload/range, are not unreasonably impaired. Untilthe introduction of the gas turbine engine, with itshigh power/weight ratio, the only powered lift systemcapable of VTOL was the low disc loading rotor, ason the helicopter.

2. Early in 1941, the late Dr A. A. Griffiths, the thenChief Scientist at Rolls-Royce, envisaged the use ofthe jet engine as a powered lift system. However, itwas not until 1947 that a light weight jet engine,designed by Rolls-Royce for missile propulsion,existed and had a high enough thrust/weight ratio forthe first pure lift-jet engine to be developed from it.

3. In 1956 the Bristol Aero-Engine Company wasapproached by Monsieur Michel Wibault with aproposal to use a turbo-shaft engine and a reductiongearbox to drive four centrifugal compressors whichwould be situated two on each side of the aircraft.The casing of these compressors could be rotated tochange direction of the thrust (fig. 18-1). The conceptincorporated two original ideas i.e. the ability todeflect the thrust over the complete range of anglesfrom the position for normal flight to that for verticallift and a system where the resultant thrust alwaysacted near to the centre of gravity of the aircraft.


Vertical/short take-off and landing


4. The principle proposed by M. Wibault wasdeveloped by using a pure jet engine with a freepower turbine to drive an axial flow fan whichexhausted into a pair of swivelling nozzles, one on

each side of the aircraft. A further development wasto use the fan to supercharge the engine, exhaustingthe by-pass air through one pair of swivelling nozzlesand adding a second pair of swivelling nozzles to the

Fig. 18-2 Lift/Propulsion engine.

Fig. 18-1 Michel Wibault's ground attack gyropter (concept) 1956

exhaust system from the engine turbine. In this waythe first ducted fan lift/propulsion engine (thePegasus) evolved (fig. 18-2).

5. Subsequent experience with the Pegasus enginein the Harrier V/STOL fighter aircraft (fig. 18-3), leadto the development of the short take-off and verticallanding (STOVL) operational technique. In this waythe additional lift generated by the aircraft wing, evenafter a short take-off run, provided a large increase inthe payload/range capability of the aircraft comparedto a pure vertical take-off. Vertical landing hadseveral operational advantages compared to a shortlanding and so was maintained.


6. Although the Pegasus engine is the only V/STOLengine in operational service in the Western Worldthere are several possible methods of providingpowered lift, such as;

(1) Deflecting (or vectoring) the exhaust gasesand hence the thrust of the engine.

(2) Using specially designed engines for lift only.

(3) Driving a lift system, which is remote from theengine, either from the engine or by aseparate power unit.

(4) Swivelling the engines.(5) For STOL aircraft, using bleed air from the

engines to increase circulation around thewing and hence increase lift.

In several of the projected V/STOL aircraft acombination of two or more of these methods hasbeen used.

Lift/Propulsion engines7. The lift/propulsion engine is capable of providingthrust for both normal wing borne flight and for lift.This is achieved by changing the direction of thethrust either by a deflector system consisting of one,two or four swivelling nozzles or by a device knownas a switch-in deflector which redirects the exhaustgases from a rearward facing propulsion nozzle toone or two downward facing lift nozzles (fig, 18-4).

8. Thrust deflection on a single nozzle is accom-plished by connecting together sections of the jet

Vertical/short take-off and landing


Fig. 18-4 Thrust deflector systems.

Fig. 18-3 V/STOL fighter aircraft.

Vertical/short take-off and landing


Fig. 18-5 Deflector nozzle.

Fig. 18-6 Side mounted swivelling nozzle.

pipe, the joint faces of which are so angled that,when the sections are counter-rotated, the nozzlemoves from the horizontal to the vertical position (fig.18-5). To avoid either a side component o! thrust or athrust line offset from the engine axis during themovement of the nozzle it is necessary that the firstjoint face is perpendicular to the axis of the jet pipe.If it is desired that the nozzle does not rotate, as maybe the case if it is a variable area nozzle, a third jointface which is perpendicular to the axis of the nozzleis required.

9. The two and four nozzle deflector systems useside mounted nozzles (fig. 18-6) which can rotate onsimple bearings through an angle of well over 90degrees so that reverse thrust can be provided ifrequired. A simple drive system, for example, asprocket and chain, can be used and by mechanicalconnections all the nozzles can be made to deflectsimultaneously. For forward flight, to avoid a highperformance loss and consequent increase in fuelconsumption, careful design of the exhaust unit andnozzle aerodynamic passages are essential tominimize the pressure losses due to turning theexhaust flow through two close coupled bends (fig.18-7).

10. The switch-in deflector consists of one or a pairof heavily reinforced doors which form part of the jetpipe wall when the engine is operating in the forwardthrust condition. To select lift thrust, the doors aremoved to blank off the conventional propelling nozzleand direct the exhaust flow into a lift nozzle (fig. 18-8). The lift nozzles may be designed so that they canbe mechanically rotated to vary the angle of thethrust and permit intermediate lift/thrust positions tobe selected.

11. A second type of switch-in deflector system isused on the tandem fan or hybrid fan vectored thrustengine (fig. 18-9). In this case the deflector system issituated between the stages of the fan of a mixedflow turbo-fan engine. In normal flight the valve ispositioned so that the engine operates in the samemanner as a mixed flow turbo-fan and for lift thrustthe valve is switched so that the exhaust flow fromthe front part of the fan exhausts through downwardfacing lift nozzles and a secondary inlet is opened toprovide the required airflow to the rear part of the fanand the main engine. On a purely subsonic V/STOLaircraft where fuel consumption is important thevalve may be dispensed with and the engineoperated permanently in the latter high by-passmode described above.

12. Thrust deflecting nozzles will create anupstream pressure distortion which may excitevibration of the fan or low pressure turbine blades ifthe nozzle system is close to these components.Snubbers (Part 3) may be used on the fan blades toresist vibration. On the low pressure turbine, shroudsat the blade tips (Part 5) or wire lacing may be usedto achieve the same result.

Lift engines13. The lift engine is designed to produce verticalthrust during the take-off and landing phases ofV/STOL aircraft. Because the engine is not used innormal flight it must be light and have a small volumeto avoid causing a large penalty on the aircraft. Thelift engine may be a turbo-jet which for a given thrustgives the lowest weight and volume. Should a low jetvelocity be necessary a lift fan may be employed.

14. Pure lift-jet engines have been developed withthrust/weight ratios of about 20:1 and still highervalues are projected for the future. Weight is reducedby keeping the engine design simple and also byextensive use of composite materials (fig. 18-10).Because the engine is operated for only limitedperiods during specific flight conditions i.e. duringtake-off and landing, the fuel system can besimplified and a total loss oil system (Part 8), in which

Vertical/short take-off and landing


Fig. 18-7 Nozzle duct configuration.

the used lubricating oil is ejected overboard, can beused.

15. Lift engines can be designed to operate in thevertical or horizontal position and a thrust deflectingnozzle fitted to provide some of the advantages ofthrust vectoring. Alternatively, the engine may bemounted so that it can swivel through a large angleto provide thrust vectoring. The lift-jet engine willhave an extremely hot, high velocity jet exhaust andto reduce ground erosion by the jet the normal

exhaust nozzle may be replaced by a multi-lobenozzle to increase the rate of mixing with thesurrounding air.

16. The lift-fan engine is designed to reduce the jetexhaust velocity, to reduce ground erosion and allowoperation from unprepared ground surfaces. It alsoreduces the jet noise significantly. A range of designoptions have been considered for this type of engineand some are shown on fig. 18-11.

Vertical/short take-off and landing


Fig. 18-9 Vectored thrust engine.

Fig. 18-8 Switch-In deflector system.

Remote lift systems17. Direct lift remote systems duct the by-pass air orengine exhaust air to downward facing lift nozzlesremote from the engine. These nozzles may be in thefront fuselage of the aircraft or in the wings. Theengine duct is blocked by means of a diverter similarto that described in para. 10.

18. The remote lift-fan (fig. 18-12) is mounted in theaircraft wing or fuselage, and is driven mechanicallyor by air or gas ducted into a tip turbine, The drivesystem is provided by the main propulsion powerplant or by a separate engine.

19. The advantage of the remote lift system is thatit gives some freedom to the aircraft to position the

Vertical/short take-off and landing


Fig. 18-11 Lift-fan engine configurations.

Fig. 18-10 A lift-jet engine.

propulsion system to the best advantage whilst stillmaintaining the resultant thrust near the aircraftcentre of gravity in the jet lift mode. This freedom isachieved at a cost of increased volume, particularlywith the gas driven systems, due to the size of theducts to feed the gas to the remote lift system.Although the mechanically driven remote lift-faneliminates the need for these large gas ducts, it isdone at the expense of long shafts and high powergearboxes and clutch systems.

Swivelling engines20. This method consists of having propulsionengines which can be mechanically swiveled closed

through at least 90 degrees to provide thrustvectoring (fig. 18-13). In addition to these propulsionengines, one or more lift engines may be installed toprovide supplementary lift during the take-off andlanding phase of flight.

21. The swivelling engine system can only be usedwith two or more engines. This then introduces theproblem of safety in the event of an engine failure.So, although there is only a small weight penalty andno increase in fuel consumption, safety considera-tions tend to offset these advantages compared tosome of the other powered lift systems. The normalmethod of providing aircraft control at low speeds isby differential throttling and vectoring of the engineswhich simplifies the basic engine design but makesthe control system more complex.

Bleed air for STOL22. Fig. 18-14 shows one method how STOL canbe achieved with a form of 'flap blowing'. The turbo-fan engine has a geared variable pitch fan and anoversized low pressure (L. P.) compressor from theexit of which air is bled and ducted to the flap systemin the wing trailing edge. The variable pitch fanenables high L.P. compressor speed and thus highbleed pressure to be maintained over a wide range ofthrusts. This gives excellent control at greatlydifferent aircraft flight conditions.


23. In many cases on V/STOL aircraft augmentationof the lift thrust is necessary to avoid an engine whichis oversized for normal flight with the consequenteffects of higher engine weight and fuel consumptionthan would be the case for a conventional aircraft-This lift thrust augmentation can be achieved in anumber of different ways:

(1) Using special engine ratings.(2) Burning in the lift nozzle gas flow.(3) By means of an ejector system.

Special engine ratings24. Experience has shown that an engine ratingstructure can be devised which provides high thrustlevels for short periods of time without reducingengine life. Operation in ground effect and the take-off and landing manoeuvres require maximum thrustfor less than 15 seconds so that use of a short liftrating for that time is feasible. Fig. 18-15 shows anexample of thrust permissible with a 15 second shortlift rating compared to that with a 2.5 minute normallift rating.

Vertical/short take-off and landing


Fig. 18-12 Remote lift fan.

Fig. 18-13 Jet lift with swivelling nozzles.

25. At high ambient temperatures, the engine mayrun into a turbine temperature limit before reachingits maximum r.p.m. and suffer a thrust loss as aresult. Restoration of the thrust can be achieved bymeans of water injection into the combustionchamber (Part 17) which allows operation at a higherturbine gas temperature for a given turbine bladetemperature. If desired, water injection can also beused to increase the thrust at low ambient tempera-tures.

Lift burning systems26. The thrust of the four nozzle lift/propulsionengine may be boosted by burning fuel in the bypassflow in the duct or plenum chamber supplying thefront nozzles. This is called plenum chamber burning(P.C.B.) (fig. 18-16) and thrust of the by-pass air maybe doubled by this process. This thrust capability isavailable for normal flight as well as take-off andlanding and so can be used to increase manoeuvra-bility and give supersonic flight.

27. The thrust of a remote lift jet can also beaugmented by burning fuel in a combustion chamberjust upstream of the lift nozzle (fig. 18-17). Thissystem is commonly known as a remote augmentedlift system (R.A.L.3.). The thrust boost available fromthe burner reduces the amount of airflow to besupplied to it and therefore reduces the size of theducting needed to direct the air from the engine tothe remote lift nozzle.

Ejectors28. The principle of the ejector is that a small, highenergy jet entrains large quantities of ambient air byviscous mixing and an increase in thrust over that ofthe high energy jet results. A number of projectedV/STOL aircraft have incorporated this concept usingeither all the engine exhaust air or just the bypassflow.

Vertical/short take-off and landing


Fig. 18-14 Flap blowing engine.

Fig. 18-15 Thrust increases with short liftratings.

Vertical/short take-off and landing


Fig. 18-16 Plenum chamber burning.

Fig. 18-17 Remote augmented lift system.


29. The low forward speeds of V/STOL aircraftduring take-off and transition do not permit thegeneration of adequate aerodynamic forces from thenormal flight control surfaces, it is thereforenecessary to provide one or more of the followingadditonal methods of controlling pitch, roll and yaw.

Reaction controls30. This system bleeds air from the engine andducts it through nozzles at the four extremities of theaircraft (fig. 18-18), The air supply to the nozzles isautomatically cut off when the main engine swivellingpropulsion nozzles are turned for normal flight orwhen the lift engines are shut down. The thrust of thecontrol nozzles is varied by changing their areawhich varies the amount of airflow passed.

Differential engine throttling31. This method of control is used on multi-enginedaircraft with the engines positioned in a suitable con-figuration. A rapid response rate is essential toenable the engines to be used for aircraft stabilityand control. It is usually necessary to combine differ-ential throttling with differential thrust vectoring togive aircraft control in all areas.

Automatic control systems32. Although it is possible for the pilot to control aV/STOL aircraft manually, some form of automationcan be of benefit and in particular will reduce the pilotworkload. The pilot's control column is electronicallyconnected to a computer or stabilizer that receivessignals from the control column, compares them withsignals from the sensors that measure the attitude ofthe aircraft, and automatically adjusts the reactioncontrols, differential throttling or thrust vectoringcontrols to maintain stability.

Vertical/short take-off and landing


Fig. 18-18 Reaction control system.

Rolls-Royce Turbomeca Adour MK151

Napier Gazelle

The Gazelle turbo-shaft engine first ran inDecember 1955 at 1260 shp, a figure laterincreased to 1610 shp on production engines.Gazelles were used to power BristolBelvedere and Westland Wessex helicopters.Gazelle production was taken over by Rolls-Royce in 1961.

19: Noise suppression

Contents Page

Introduction 199Engine noise 199Methods of suppressing noise202Construction and materials 205


1. Airport regulations and aircraft noise certificationrequirements, all of which govern the maximumnoise level aircraft are permitted to produce, havemade jet engine noise suppression one of the mostimportant fields of research.

2. The unit that is commonly used to express noiseannoyance is the Effective Perceived Noise deciBel(EPNdB). It takes into account the pitch as well asthe sound pressure (deciBel) and makes allowancefor the duration of an aircraft flyover. Fig. 19-1compares the noise levels of various jet enginetypes.

3. Airframe self-generated noise is a factor in anaircraft's overall noise signature, but the principalnoise source is the engine.


4. To understand the problem of engine noisesuppression, it is necessary to have a workingknowledge of the noise sources and their relativeimportance. The significant sources originate in thefan or compressor, the turbine and the exhaust jet orjets. These noise sources obey different laws andmechanisms of generation, but all increase, to avarying degree, with greater relative airflow velocity.Exhaust jet noise varies by a larger factor than thecompressor or turbine noise, therefore a reduction inexhaust jet velocity has a stronger influence than anequivalent reduction in compressor and turbine bladespeeds.


Noise suppression


Fig. 19-1 Comparative noise levels of various engine types.

Fig. 19-2 Exhaust mixing and shock structure.

5. Jet exhaust noise is caused by the violent andhence extremely turbulent mixing of the exhaustgases with the atmosphere and is influenced by theshearing action caused by the relative speedbetween the exhaust jet and the atmosphere. Thesmall eddies created near the exhaust duct causehigh frequency noise but downstream of the exhaustjet the larger eddies create low frequency noise.Additionally, when the exhaust jet velocity exceedsthe local speed of sound, a regular shock pattern isformed within the exhaust jet core. This produces adiscrete (single frequency) tone and selective ampli-fication of the mixing noise, as shown in fig. 19-2. Areduction in noise level occurs if the mixing rate isaccelerated or if the velocity of the exhaust jetrelative to the atmosphere is reduced. This can beachieved by changing the pattern of the exhaust jetas shown in fig. 19-3.

6. Compressor and turbine noise results from theinteraction of pressure fields and turbulent wakesfrom rotating blades and stationary vanes, and canbe defined as two distinct types of noise; discretetone (single frequency) and broadband (a wide rangeof frequencies). Discrete tones are produced by theregular passage of blade wakes over the stagesdownstream causing a series of tones andharmonics from each stage. The wake intensity islargely dependent upon the distance between therows of blades and vanes. If the distance is shortthen there is an intense pressure field interactionwhich results in a strong tone being generated. Withthe high bypass engine, the low pressurecompressor (fan) blade wakes passing overdownstream vanes produce such tones, but of alower intensity due to lower velocities and largerblade/vane separations. Broadband noise is

Noise suppression


Fig. 19-3 Change of exhaust jet pattern to reduce noise level.

produced by the reaction of each blade to thepassage of air over its surface, even with a smoothairstream. Turbulence in the airstream passing overthe blades increases the intensity of the broadbandnoise and can also induce tones.

7. With the pure jet engine the exhaust jet noise isof such a high level that the turbine and compressornoise is insignificant at all operating conditions,except low landing-approach thrusts. With the by-pass principle, the exhaust jet noise drops as thevelocity of the exhaust is reduced but the lowpressure compressor and turbine noise increasesdue to the greater internal power handling.

8. The introduction of a single stage low pressurecompressor (fan) significantly reduces thecompressor noise because the overall turbulenceand interaction levels are diminished. When the by-pass ratio is in excess of approximately 5 to 1, the jetexhaust noise has reduced to such a level that theincreased internal noise source is predominant. Acomparison between low and high by-pass enginenoise sources is shown in fig. 19-4.

9. Listed amongst the several other sources ofnoise within the engine is the combustion chamber. Itis a significant but not a predominant source, due inpart to the fact that it is 'buried' in the core of theengine. Nevertheless it contributes to the broadbandnoise, as a result of the violent activities which occurwithin the combustion chamber.


10. Noise suppression of internal sources isapproached in two ways; by basic design to minimizenoise originating within or propagating from theengine, and by the use of acoustically absorbentlinings. Noise can be minimized by reducing airflowdisruption which causes turbulence. This is achievedby using minimal rotational and airflow velocities andreducing the wake intensity by appropriate spacingbetween the blades and vanes. The ratio betweenthe number of rotating blades and stationary vanescan also be advantageously employed to containnoise within the engine.

Noise suppression


Fig. 19-4 Comparative noise sources of low and high by-pass engines.

11. As previously described, the major source ofnoise on the pure jet engine and low by-pass engineis the exhaust jet, and this can be reduced byinducing a rapid or shorter mixing region. Thisreduces the low frequency noise but may increasethe high frequency level. Fortunately, highfrequencies are quickly absorbed in the atmosphereand some of the noise which does propagate to thelistener is beyond the audible range, thus giving theperception of a quieter engine. This is achieved byincreasing the contact area of the atmosphere withthe exhaust gas stream by using a propelling nozzleincorporating a corrugated or lobe-type noisesuppressor (fig. 19-5).

12. In the corrugated nozzle, freestreamatmospheric air flows down the outside corrugationsand into the exhaust jet to promote rapid mixing. Inthe lobe-type nozzle, the exhaust gases are dividedto flow through the lobes and a small central nozzle.This forms a number of separate exhaust jets thatrapidly mix with the air entrained by the suppressorlobes. This principle can be extended by the use of aseries of tubes to give the same overall area as thebasic circular nozzle.

13. Deep corrugations, lobes, or multi-tubes, givethe largest noise reductions, but the performancepenalties incurred limit the depth of the corrugationsor lobes and the number of tubes. For instance, toachieve the required nozzle area, the overalldiameter of the suppressor may have to beincreased by so much that excessive drag andweight results. A compromise which gives anoticeable reduction in noise level with the leastsacrifice of engine thrust, fuel consumption oraddition of weight is therefore the designer's aim.

14. The high by-pass engine has two exhauststreams to eject to atmosphere. However, theprinciple of jet exhaust noise reduction is the sameas for the pure or low by-pass engine, i.e. minimizethe exhaust jet velocity within overall performanceobjectives. High by-pass engines inherently have alower exhaust jet velocity than any other type of gasturbine, thus leading to a quieter engine, but furthernoise reduction is often desirable. The mostsuccessful method used on by-pass engines is tomix the hot and cold exhaust streams within theconfines of the engine (fig. 19-5) and expel the lowervelocity exhaust gas flow through a single nozzle(Part 6).

15. In the high by-pass ratio engine thepredominant sources governing the overall noiselevel are the fan and turbine. Research has produced

a good understanding of the mechanisms of noisegeneration and comprehensive noise design rulesexist. As previously indicated, these are founded onthe need to minimize turbulence levels in the airflow,reduce the strength of interactions between rotatingblades and stationary vanes, and the optimum use ofacoustically absorbent linings.

Noise suppression


Fig. 19-5 Types of noise suppressor.

Noise suppression


Fig. 19-6 Noise absorbing materials and location.

16. Noise absorbing 'lining' material convertsacoustic energy into heat. The absorbent linings (fig.19-6) normally consist of a porous skin supported bya honeycomb backing, to provide the requiredseparation between the facesheet and the solidengine duct. The acoustic properties of the skin andthe liner depth are carefully matched to the characterof the noise, for optimum suppression. The disad-vantage of liners is the slight increase in weight andskin friction and hence a slight increase in fuelconsumption. They do however, provide a verypowerful suppression technique.


17. The corrugated or lobe-type noise suppressorforms the exhaust propelling nozzle and is usually aseparate assembly bolted to the jet pipe. Provision isusually made to adjust the nozzle area so that it can

be accurately calibrated. Guide vanes are fitted tothe lobe-type suppressor to prevent excessive lossesby guiding the exhaust gas smoothly through thelobes to atmosphere. The suppressor is a fabricatedwelded structure and is manufactured from heat-resistant alloys.

18. Various noise absorbing lining materials areused on jet engines. They fall mainly within twocategories, lightweight composite materials that areused in the lower temperature regions and fibrous-metallic materials that are used in the highertemperature regions. The noise absorbing materialconsists of a perforate metal or composite facingskin, supported by a honeycomb structure on a solidbacking skin which is bonded to the parent metal ofthe duct or casing. For details of manufacture ofthese materials refer to Part 22.

Noise suppression


Rolls-Royce Conway

Rolls-Royce RM60

Produced in response to an Admiralty contractfor a coastal-craft engine with good cruisingeconomy, the RM60, although based onaeroengine philosophy, was designed fromthe first as a marine gas turbine. Two RM60swent to sea in 1953 in the former steamgunboat HMS Grey Goose, the world's firstwarship to be powered solely by gas turbines.

20: Thrust distribution

Contents Page

Introduction 207Distribution of the thrust forces 207Method of calculating the thrust forces 209Calculating the thrust of the engine 209

Compressor casingDiffuser ductCombustion chambersTurbine assemblyExhaust unit and jet pipePropelling nozzleEngineInclined combustion chambers

Afterburning 212


1. Although the principles of jet propulsion (see Part1) will be familiar to the reader, the distribution of thethrust forces within the engine may appearsomewhat obscure- These forces are in effect gasloads resulting from the pressure and momentumchanges of the gas stream reacting on the enginestructure and on the rotating components. They arein some locations forward propelling forces and inothers opposing or rearward forces. The amount that

the sum of the forward forces exceeds the sum of therearward forces is normally known as the rated thrustof the engine.


2. The diagram in fig. 20-1 is of a typical single-spool axial flow turbo-jet engine and illustrates wherethe main forward and rearward forces act. The originof these forces is explained by following the engineworking cycle shown in Part 2.


3. At the start of the cycle, air is induced into theengine and is compressed. The rearward accelera-tions through the compressor stages and theresultant pressure rise produces a large reactiveforce in a forward direction. On the next stage of itsjourney the air passes through the diffuser where itexerts a small reactive force, also in a forwarddirection,

4. From the diffuser the air passes into thecombustion chambers (Part 4) where it is heated,and in the consequent expansion and acceleration ofthe gas large forward forces are exerted on thechamber walls.

5. When the expanding gases leave the combustionchambers and flow through the nozzle guide vanesthey are accelerated and deflected on to the bladesof the turbine (Part 5). Due to the acceleration anddeflection, together with the subsequent straighten-ing of the gas flow as it enters the jet pipe, consider-able 'drag' results; thus the vanes and blades aresubjected to large rearward forces, the magnitude of

which may be seen on the diagram. As the gas flowpasses through the exhaust system (Part 6), smallforward forces may act on the inner cone or bullet,but generally only rearward forces are produced andthese are due to the 'drag' of the gas flow at thepropelling nozzle.

6. It will be seen that during the passage of the airthrough the engine, changes in its velocity andpressure occur (Part 2). For instance, where aconversion from velocity (kinetic) energy to pressureenergy is required the passages are divergent inshape, similar to that used in the compressordiffuser. Conversely, where it is required to convertthe energy stored in the combustion gases tovelocity, a convergent passage or nozzle, similar tothat used in the turbine, is employed. Where theconversion is to velocity energy, 'drag' loads orrearward forces are produced; where the conversionis to pressure energy, forward forces are produced.Part 2, fig. 2-3 illustrates velocity and pressurechanges at two points on the engine.

Thrust distribution


Fig. 20-1 Thrust distribution of a typical single-spool axial flow engine.


7. The thrust forces or gas loads can be calculatedfor the engine, or for any flow section of the engine,provided that the areas, pressures, velocities andmass flow are known for both the inlet and outlet ofthe particular flow section.

8. The distribution of thrust forces shown in fig. 20-1 can be calculated by considering each componentin turn and applying some simple calculations. Thethrust produced by the engine is mainly the productof the mass of air passing through the engine and thevelocity increase imparted to it (i.e. Newtons SecondLaw of Motion), however, the pressure differencebetween the inlet to and the outlet from the particularflow section will have an effect on the overall thrustof the engine and must be included in the calculation.

9. To calculate the resultant thrust for a particularflow section it is necessary to calculate the totalthrust at both inlet and outlet, the resultant thrustbeing the difference between the two valuesobtained.

10. Calculation of the thrust is achieved using thefollowing formula:

Where A = Area of flow section in = Pressure in lb. per = Mass flow in lb. per sec. vJ = Velocity of flow in feet per sec. g = Gravitational constant 32.2 ft. per

sec. per sec.


11. When applying the above method to calculatethe individual thrust loads on the various componentsit is assumed that the engine is static. The effect ofaircraft forward speed on the engine thrust will bedealt with in Part 21. In the following calculations 'g'is taken to be 32 for convenience. To assist in thesecalculations the locations concerned are illustratedby a number of small diagrams.

Compressor casing12. To obtain the thrust on the compressor casing itis necessary to calculate the conditions at the inlet to

the compressor and the conditions at the outlet fromthe compressor. Since the pressure and the velocityat the inlet to the compressor are zero, it is onlynecessary to consider the force at the outlet from thecompressor. Therefore, given that the compressor-

OUTLET Area (A) = 182 (P) = 94 lb. per

(gauge)Velocity (vJ) = 406 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

= 19,049 lb. of thrust in a forward direction.

Diffuser duct13. The conditions at the diffuser duct inlet are thesame as the conditions at the compressor outlet, i.e.19,049 lb.Therefore, given that the diffuser--OUTLET Area (A) = 205

Pressure (P) = 95 lb. per

Velocity (vJ) = 368 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

= 21,235 - 19,049

= 2,186 lb. of thrust in a forward direction.

Thrust distribution


gM)PxA(Thrust JV+=


406x153)94x182( −+=


368x153)95x205( −+=


W)PxA( JV −+=


M)PxA( JV −+=

Combustion chambers14. The conditions at the combustion chamber inletare the same as the conditions at the diffuser outlet,i.e. 21,235 lb. Therefore, given that the combustionchamber-OUTLET Area (A) = 580

Pressure (P) = 93 lb. per

Velocity (vJ) = 309 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

= 55,417 - 21,235

= 34,182 !b. of thrust in a forward direction.

Turbine assembly15. The conditions at the turbine inlet are the sameas the conditions at the combustion chamber outlet,i.e. 55,417 lb.

Therefore given that the turbine--

OUTLET Area (A) = 480 (P) = 21 lb. per

(gauge)Velocity (vJ) = 888 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

= 14,326 - 55,417

= -41,091

This negative value means a force acting in arearward direction.

Exhaust unit and jet pipe16. The conditions at the inlet to the exhaust unitare the same as the conditions at the turbine outlet,i.e. 14,326 lb. Therefore, given that the jet pipe--OUTLET Area (A) = 651

Pressure (P) = 21 lb. per

Velocity (vJ) = 643 ft. per sec.Mass flow (W) = 153 lb. per sec.

Thrust distribution



W)PxA( JV −+=


W)PxA( JV −+=


309x153)93x580( −+=


888x153)21x480( −+=

The thrust

= 16,745 - 14,326

= 2,419 lb. of thrust in a forward direction.

Propelling nozzle17. The conditions at the inlet to the propellingnozzle are the same as the conditions at the jet pipeoutlet, i.e. 16,745 lb. Therefore, given that the propelling nozzle--OUTLET Area (A) = 332

Pressure (P) = 6 lb. per

Velocity (vJ) = 1,917 ft. per sec.Mass flow (W) = 153 lb. per sec.

The thrust

= 11,158 - 16,745

= 5,587lb. acting in a rearward direction.

It is emphasized that these are basic calculationsand such factors as the effect of air offtakes havebeen ignored.

18. Based on the individual calculations, the sum ofthe forward or positive loads is 57,836 lb. and thesum of the rearward or negative loads is 46,678 lb.Thus, the resultant (gross or total) thrust is 11,158 lb.

Engine19. It will be of interest to calculate the thrust of theengine by considering the engine as a whole, as theresultant thrust should be equal to the sum of theindividual gas loads previously calculated.

20. Although the momentum change of the gasstream produces most of the thrust developed by the

engine (momentum thrust = ), an additional

thrust is produced when the engine operates with thepropelling nozzle in a 'choked' condition (Part 6). Thisthrust results from the aerodynamic forces which arecreated by the gas stream and exert a pressure

Thrust distribution



W)PxA( JV −+= 745,16g

W)PxA( JV −+=


643x153)21x651( −+= 745,1632

917,1x153)6x332( −+=


across the exit area of the propelling nozzle(pressure thrust). Algebraically, this force isexpressed as (P-P0) A.

Where A = Area of propelling nozzle in P = Pressure in lb. per P0 = Atmospheric pressure in lb. per

Therefore, assuming values of mass flow, pressureand area to be the same as in the previous calcula-tions i.e.

Area of propelling nozzle (A) = 332 (P) = 6 lb. per

(gauge)Atmospheric Pressure (P) = 0 lb. per

(gauge)Mass flow (W) = 153 lb. per sec.Velocity (vJ) = 1,917 ft. per sec.

The thrust

= 1,992 + 9,166

= 11,158 lb., the same as previously calculated

by combining the gas loads on the individualengine locations.

21. On engines that operate with a non-chokednozzle, the (P-P0) A function does not apply and thethrust results only from the gas stream momentumchange.

Inclined combustion chambers22. In the previous example (Para. 14) the flowthrough the combustion chamber is axial, however, ifthe combustion chamber is inclined towards the axisof the engine, then the axial thrust will be less thanfor an axial flow chamber. This thrust can be obtainedby multiplying the sum of the outlet thrust by thecosine of the angle (see fig. 20-2). The

cosine = and for a given angle

is obtained by consulting a table of cosines. It shouldbe emphasized that if the inlet and outlet are atdifferent angles to the engine axis, it is necessary tomultiply the inlet and outlet thrusts separately by thecosine of their respective angles.


23. When the engine is fitted with an afterburner(Part 16), the gases passing through the exhaust

Thrust distribution


Fig. 20-2 A hypothetical combustion chamber showing values required for calculating thrust.


WA)PP( JV0 −+⋅−=


917,1X153332)06( −+⋅−=


system are reheated to provide additional thrust. Theeffect of afterburning is to increase the volume of theexhaust gases, thus producing a higher exit velocityat the propelling nozzle.

24. Assuming that an afterburner jet pipe andpropelling nozzle are fitted to the engine used in theprevious calculations, and the new conditions at thepropelling nozzle are as follows-

OUTLET Area (A) = 455 (P) = 5 lb. per

(gauge)Velocity (vJ) = 2,404 ft. per sec.Mass flow (W) = 157 lb. per sec.

The thrust

= 14,069 - 16,745

= 2,676 lb. acting in a rearward direction.

Therefore, compared with the previous calculation inpara. 17, it will be seen that the negative thrust isreduced from -5,587 lb. to -2,676 lb.; the overallpositive thrust is thus increased by 2,911 lb; which isequivalent to a thrust increase of more than 25 percent.

25. To arrive at the total thrust of the engine withafterburning the calculations in para. 20 should usethe above figures.

Thrust distribution



W)PxA( JV −+=


404,2x157)5x455( −+=

Rolls-Royce RB168 MK807

Blackburn Nimbus

The Nimbus was developed from the A129turbo-shaft which, in its turn, was a modifiedTurbomeca Artouste built under licence. TheNimbus developed 968 hp, but for helicopteruse was flat-rated at 710 hp. The engine wasused in Westland Wasp and Scout helicoptersand four 700 hp units were used to power theexperimental 5RN-2 hovercraft.

21: Performance

Contents Page

Introduction 215Engine thrust on the test bench 217

Comparison between thrust and horse-power

Engine thrust in flight 218Effect of forward speedEffect of afterburning on engine thrustEffect of altitudeEffect of temperature

Propulsive efficiency 223 Fuel consumption and power-to-weight relationship 225


1. The performance requirements of an engine areobviously dictated to a large extent by the type ofoperation for which the engine is designed. Thepower of the turbo-jet engine is measured in thrust,produced at the propelling nozzle or nozzles, andthat of the turbo-propeller engine is measured inshaft horse-power (s.h.p.) produced at the propellershaft. However, both types are in the main assessedon the amount of thrust or s.h.p. they develop for agiven weight, fuel consumption and frontal area.

2. Since the thrust or s.h.p. developed is dependenton the mass of air entering the engine and the accel-eration imparted to it during the engine cycle, it isobviously influenced, as subsequently described, bysuch variables as the forward speed of the aircraft,altitude and climatic conditions, These variablesinfluence the efficiency of the air intake, thecompressor, the turbine and the jet pipe; conse-quently, the gas energy available for the productionof thrust or s.h.p. also varies.

3. In the interest of fuel economy and aircraft range,the ratio of fuel consumption to thrust or s.h.p. shouldbe as low as possible. This ratio, known as thespecific fuel consumption (s.f.c.), is expressed inpounds of fuel per hour per pound of net thrust ors.h.p. and is determined by the thermal andpropulsive efficiency of the engine. In recent yearsconsiderable progress has been made in reducings.f.c. and weight. These factors are further explainedin para. 46.


4. Whereas the thermal efficiency is often referredto as the internal efficiency of the engine, thepropulsive efficiency is referred to as the externalefficiency. This latter efficiency, described in para. 37,explains why the pure jet engine is less efficient thanthe turbo-propeller engine at lower aircraft speedsleading to development of the by-pass principle and,more recently, the propfan designs.

5. The thermal and the propulsive efficiency alsoinfluence, to a large extent, the size of thecompressor and turbine, thus determining the weightand diameter of the engine for a given output.

6. These and other factors are presented in curvesand graphs, calculated from the basic gas laws (Part2), and are proved in practice by bench and flighttesting, or by simulating flight conditions in a highaltitude test cell. To make these calculations, specificsymbols are used to denote the pressures and tem-peratures at various locations through the engine; for

instance, using the symbols shown in fig. 21-1 the

overall compressor pressure ratio is . These

symbols vary slightly for different types of engine; forinstance, with high by-pass ratio engines, and alsowhen afterburning (Part 16) is incorporated,additional symbols are used.

7. To enable the performance of similar engines tobe compared, it is necessary to standardize in someconventional form the variations of air temperatureand pressure that occur with altitude and climaticconditions. There are in use several differentdefinitions of standard atmospheres, the one in mostcommon use being the International StandardAtmosphere (I.S.A.). This is based on a temperaturelapse rate of approximately 1.98 K. degrees per1,000ft,, resulting in a fall from 288.15 deg.K. (15deg.C) at sea level to 216.65 deg.K (-56.5 deg.C.) at36,089 ft. (the tropopause). Above this altitude the



Fig. 21-1 Temperature and pressure notation of a typical turbo-jet engine.




temperature is constant up to 65.617ft. The I.S.A.standard pressure at sea level is 14.69 pounds persquare inch falling to 3.28 pounds per square inch atthe tropopause (refer to I.S.A. table fig. 21-10).


8. The thrust of the turbo-jet engine on the testbench differs somewhat from that during flight.Modern test facilities are available to simulateatmospheric conditions at high altitudes thusproviding a means of assessing some of theperformance capability of a turbo-jet engine in flightwithout the engine ever leaving the ground. This isimportant as the changes in ambient temperatureand pressure encountered at high altitudes consider-ably influence the thrust of the engine.

9. Considering the formula derived in Part 20 forengines operating under 'choked' nozzle conditions,

it can be seen that the thrust can be further affectedby a change in the mass flow rate of air through theengine and by a change in jet velocity. An increase inmass airflow may be obtained by using waterinjection (Part 17) and increases in jet velocity byusing afterburning (Part 16).

10. As previously mentioned, changes in ambientpressure and temperature considerably influence thethrust of the engine. This is because of the way theyaffect the air density and hence the mass of airentering the engine for a given engine rotationalspeed. To enable the performance of similar enginesto be compared when operating under differentclimatic conditions, or at different altitudes, correctionfactors must be applied to the calculations to returnthe observed values to those which would be foundunder I.S.A. conditions. For example, the thrustcorrection for a turbo-jet engine is: Thrust (lb.) (corrected) =

thrust (lb.) (observed) x

where P0 = atmospheric pressure in inches of mercury (in. Hg.) (observed)

30 = I.S.A. standard sea level pressure (in.Hg.)

11. The observed performance of the turbo-propeller engine is also corrected to I.S.A.conditions, but due to the rating being in s.h.p. and

not in pounds of thrust the factors are different. Forexample, the correction for s.h.p. is:S.h.p. (corrected) =

s.h.p. (observed)

where P0 = atmospheric pressure (in.Hg.)(observed)

T0 = atmospheric temperature in deg.C.(observed)

30 = I.S.A. standard sea level pressure(in.Hg.)

273 + 15 = I.S.A. standard sea leveltemperature in deg.K.

273 + T0 = Atmospheric temperature in deg.K.

In practice there is always a certain amount of jetthrust in the total output of the turbo-propeller engineand this must be added to the s.h.p. The correctionfor jet thrust is the same as that in para. 10.

12. To distinguish between these two aspects of thepower output, it is usual to refer to them as s.h.p. andthrust horse-power (t.h.p.). The total equivalenthorse-power is denoted by t.e.h.p. (sometimese.h.p.) and is the s.h.p. plus the s.h.p. equivalent tothe net jet thrust. For estimation purposes it is takenthat, under sea- level static conditions, one s.h.p. isequivalent to approximately 2.6 lb. of jet thrust.Therefore :

13. The ratio of jet thrust to shaft power isinfluenced by many factors. For instance, the higherthe aircraft operating speed the larger may be therequired proportion of total output in the form of jetthrust. Alternatively, an extra turbine stage may berequired if more than a certain proportion of the totalpower is to be provided at the shaft. In general,turbo-propeller aircraft provide one pound of thrustfor every 3.5 h,p. to 5 h.p.

Comparison between thrust and horse-power14. Because the turbo-jet engine is rated in thrustand the turbo-propeller engine in s.h.p., no directcomparison between the two can be made without apower conversion factor. However, since the turbo-propeller engine receives its thrust mainly from thepropeller, a comparison can be made by convertingthe horse-power developed by the engine to thrust orthe thrust developed by the turbo-jet engine to t.h.p.;that is, by converting work to force or force to work.For this purpose, it is necessary to take into accountthe speed of the aircraft.



gWA)PP(Thrust JV

0 +⋅−=


00 T27315273x



6.2.lbthrustjet.p.h.s.p.h.e.t +=

15. The t.h.p. is expressed as

where F = lb. of thrustV = aircraft speed (ft. per sec.)

Since one horse-power is equal to 550 per sec.and 550 ft. per sec. is equivalent to 375 miles perhour, it can be seen from the above formula that onelb. of thrust equals one t.h.p. at 375 m.p.h. It is alsocommon to quote the speed in knots (nautical milesper hour); one knot is equal to 1.1515 m.p.h, or onepound of thrust is equal to one t.h.p. at 325 knots.

16. Thus if a turbo-jet engine produces 5,000 lb. ofnet thrust at an aircraft speed of 600 m.p.h. the t.h.p.

would be

However, if the same thrust was being produced bya turbo-propeller engine with a propeller efficiency of55 per cent at the same flight speed of 600 m.p.h.,then the t.h.p. would be

Thus at 600 m.p.h. one lb. of thrust is the equivalentof about 3 t.h.p.


17. Since reference will be made to gross thrust,momentum drag and net thrust, it will be helpful todefine these terms:from Part 20, gross or total thrust is the product of themass of air passing through the engine and the jetvelocity at the propelling nozzle, expressed as:

The momentum drag is the drag due to themomentum of the air passing into the engine relative

to the aircraft velocity, expressed as where

W = Mass flow in lb. per sec.V = Velocity of aircraft in feet per sec.g = Gravitational constant 32.2 ft. per sec. per

sec.The net thrust or resultant force acting on the aircraftin flight is the difference between the gross thrustand the momentum drag.

18. From the definitions and formulae stated inpara, 17; under flight conditions, the net thrust of the



Fig. 21-2 The balance of forces and expression for thrust and momentum drag.



600x000,5 =


100x000,8 =

gWA)PP( Jv

0 +−


engine, simplifying, can be expressed as:

Fig. 21-2 provides a diagrammatic explanation.

Effect of forward speed19. Since reference will be made to 'ram ratio' andMach number, these terms are defined as follows:

Ram ratio is the ratio of the total air pressure atthe engine compressor entry to the static airpressure at the air intake entry.

Mach number is an additional means ofmeasuring speed and is defined as the ratio ofthe speed of a body to the local speed of sound.Mach 1.0 therefore represents a speed equal tothe local speed of sound.

20. From the thrust equation in para. 18, it isapparent that if the jet velocity remains constant,independent of aircraft speed, then as the aircraftspeed increases the thrust would decrease in directproportion. However, due to the 'ram ratio' effect fromthe aircraft forward speed, extra air is taken into theengine so that the mass airflow and also the jetvelocity increase with aircraft speed. The effect ofthis tends to offset the extra intake momentum drag

due to the forward speed so that the resultant loss ofnet thrust is partially recovered as the aircraft speedincreases. A typical curve illustrating this point isshown in fig. 21-3. Obviously, the 'ram ratio' effect, orthe return obtained in terms of pressure rise at entryto the compressor in exchange for the unavoidableintake drag, is of considerable importance to theturbo-jet engine, especially at high speeds. Abovespeeds of Mach 1.0, as a result of the formation ofshock waves at the air intake, this rate of pressurerise will rapidly decrease unless a suitably designedair intake is provided (Part 23); an efficient air intakeis necessary to obtain maximum benefit from the ramratio effect.

21. As aircraft speeds increase into the supersonicregion, the ram air temperature rises rapidlyconsistent with the basic gas laws (Part 2). This



Fig. 21-3 Thrust recovery with aircraftspeed.

Fig. 21-4 The effect of aircraft speed onthrust and fuel consumption.

g)Vv(WA)PP( J


temperature rise affects the compressor delivery airtemperature proportionately and, in consequence, tomaintain the required thrust, the engine must besubjected to higher turbine entry temperatures. Sincethe maximum permissible turbine entry temperatureis determined by the temperature limitations of theturbine assembly, the choice of turbine materials andthe design of blades and stators to permit cooling arevery important.

22. With an increase in forward speed, theincreased mass airflow due to the 'ram ratio' effectmust be matched by the fuel flow (Part 10) and theresult is an increase in fuel consumption. Becausethe net thrust tends to decrease with forward speedthe end result is an increase in specific fuelconsumption (s.f.c.), as shown by the curves for atypical turbo-jet engine in fig, 21-4.

23. At high forward speeds at low altitudes the 'ramratio' effect causes very high stresses on the engineand, to prevent overstressing, the fuel flow is auto-matically reduced to limit the engine speed andairflow. The method of fuel control is described inPart 10.

24. The effect of forward speed on a typical turbo-propeller engine is shown by the trend curves in fig.21 -5. Although net jet thrust decreases, s.h.p.increases due to the 'ram ratio1 effect of increasedmass flow and matching fuel flow. Because it isstandard practice to express the s.f.c. of a turbo-propeller engine relative to s.h.p., an improved exhibited. However, this does not provide a truecomparison with the curves shown in fig. 21-4, for atypical turbo-jet engine, as s.h.p, is absorbed by thepropeller and converted into thrust and, irrespectiveof an increase in s.h.p., propeller efficiency andtherefore net thrust deteriorates at high subsonicforward speeds. In consequence, the turbo-propellerengine s.f.c, relative to net thrust would, in generalcomparison with the turbo-jet engine, show animprovement at low forward speeds but a rapid dete-rioration at high speeds.

Effect of afterburning on engine thrust25. At take-off conditions, the momentum drag ofthe airflow through the engine is negligible, so thatthe gross thrust can be considered to be equal to thenet thrust. If afterburning (Part 16) is selected, anincrease in take-off thrust in the order of 30 per centis possible with the pure jet engine and considerablymore with the by-pass engine. This augmentation ofbasic thrust is of greater advantage for certainspecific operating requirements.

26. Under flight conditions, however, this advantageis even greater, since the momentum drag is thesame with or without afterburning and, due to theram effect, better utilization is made of every pound



Fig. 21-5 The effect of aircraft speed ons.h.p. and fuel consumption.

of air flowing through the engine. The followingexample, using the static values given in Part 16,illustrates why afterburning thrust improves underflight conditions.

27. Assuming an aircraft speed of 600 m.p.h. (880ft.per sec.), then Momentum drag is:

This means that every pound of air per secondflowing through the engine and accelerated up to thespeed of the aircraft causes a drag of about 27.5 lb.

28. Suppose each pound of air passed through theengine gives a gross thrust of 77.5 lb. Then the netthrust given by the engine per lb. of air per second is77.5 - 27.5 = 50 lb.

29. When afterburning is selected, assuming the 30per cent increase in static thrust given in para. 25,the gross thrust will be 1.3 x 77.5 - 100.75 lb. Thus,under flight condition of 600 m.p.h., the net thrust perpound of air per second will be 100.75 - 27.5 = 73.25lb. Therefore, the ratio of net thrust due to

afterburning is = 1.465. In other words, a 30

per cent increase in thrust under static conditionsbecomes a 46.5 per cent increase in thrust at 600m.p.h.

30. This larger increase in thrust is invaluable forobtaining higher speeds and higher altitude perform-ances. The total and specific fuel consumptions arehigh, but not unduly so for such an increase inperformance.

31. The limit to the obtainable thrust is determinedby the afterburning temperature and the remainingusable oxygen in the exhaust gas stream. Becauseno previous combustion heating takes place in theduct of a by-pass engine, these engines with theirlarge residual oxygen surplus are particularly suitedto afterburning and static thrust increases of up to 70per cent are obtainable. At high forward speedsseveral times this amount is achieved.

Effect of altitude32. With increasing altitude the ambient airpressure and temperature are reduced. This affectsthe engine in two interrelated ways:

The fall of pressure reduces the air density andhence the mass airflow into the engine for agiven engine speed. This causes the thrust ors.h.p. to fall. The fuel control system, asdescribed in Part 10, adjusts the fuel pump

output to match the reduced mass airflow, somaintaining a constant engine speed.

The fall in air temperature increases the densityof the air, so that the mass of air entering thecompressor for a given engine speed is greater.This causes the mass airflow to reduce at alower rate and so compensates to some extentfor the loss of thrust due to the fall in atmosphericpressure. At altitudes above 36,089 feet and upto 65,617 feet, however, the temperatureremains constant, and the thrust or s.h.p. isaffected by pressure only.

Graphs showing the typical effect of altitude onthrust, s.h.p, and fuel consumption are illustrated infig. 21-6 and fig. 21-7.

Effect of temperature33. On a cold day the density of the air increases sothat the mass of air entering the compressor for agiven engine speed is greater, hence the thrust ors.h.p, is higher. The denser air does, however,increase the power required to drive the compressoror compressors; thus the engine will require morefuel to maintain the same engine speed or will run ata reduced engine speed if no increase in fuel isavailable.

34. On a hot day the density of the air decreases,thus reducing the mass of air entering thecompressor and, consequently, the thrust of theengine for a given r.p.m. Because less power will berequired to drive the compressor, the fuel controlsystem reduces the fuel flow to maintain a constantengine rotational speed or turbine entry temperature,as appropriate; however, because of the decrease inair density, the thrust will be lower. At a temperatureof 45 deg.C., depending on the type of engine, athrust loss of up to 20 per cent may be experienced.This means that some sort of thrust augmentation,such as water injection (Part 17), may be required.

35. The fuel control system (Part 10) controls thefuel flow so that the maximum fuel supply is heldpractically constant at low air temperature conditions,whereupon the engine speed falls but, because ofthe increased mass airflow as a result of the increasein air density, the thrust remains the same. Forexample, the combined acceleration and speedcontrol fuel system (Part 10) schedules fuel flow tomaintain a constant engine r.p.m., hence thrustincreases as air temperature decreases until, at apredetermined compressor delivery pressure, thefuel flow is automatically controlled to maintain aconstant compressor delivery pressure and,




880 =


therefore, thrust. Fig. 21-8 illustrates this for a twin-spool engine where the controlled engine r.p.m. ishigh pressure compressor speed and thecompressor delivery pressure is expressed as P3. Itwill also be apparent from this graph that the lowpressure compressor speed is always less than itslimiting maximum and that the difference in the twospeeds is reduced by a decrease in ambient airtemperature. To prevent the L.P. compressor over-speeding, fuel flow is also controlled by an L.P.governor which, in this case, takes a passive role.

36. The pressure ratio control fuel system (Part 10)schedules fuel flow to maintain a constant enginepressure ratio and, therefore, thrust below a prede-



Fig. 21-6 The effects of altitude on thrustand fuel consumption.

Fig. 21-7 The effect of altitude on s.h.p. andfuel consumption.

termined ambient air temperature. Above thistemperature the fuel flow is automatically controlledto prevent turbine entry temperature limitations frombeing exceeded, thus resulting in reduced thrust and,overall, similar curve characteristics to those shownin fig. 21-8. In the instance of a triple-spool enginethe pressure ratio is expressed as P4/P1. i.e. H.P.compressor delivery pressure/engine inlet pressure.


37. Performance of the jet engine is not onlyconcerned with the thrust produced, but also with theefficient conversion of the heat energy of the fuel intokinetic energy, as represented by the jet velocity, andthe best use of this velocity to propel the aircraftforward, i.e. the efficiency of the propulsive system.

38. The efficiency of conversion of fuel energy tokinetic energy is termed thermal or internal efficiencyand, like all heat engines, is controlled by the cyclepressure ratio and combustion temperature.Unfortunately, this temperature is limited by thethermal and mechanical stresses that can betolerated by the turbine. The development of newmaterials and techniques to minimize theselimitations is continually being pursued.

39. The efficiency of conversion of kinetic energy topropulsive work is termed the propulsive or externalefficiency and this is affected by the amount of kinetic

energy wasted by the propelling mechanism. Wasteenergy dissipated in the jet wake, which represents a

loss, can be expressed as where (vJ-V)

is the waste velocity. It is therefore apparent that atthe aircraft lower speed range the pure jet streamwastes considerably more energy than a propellersystem and consequently is less efficient over thisrange. However, this factor changes as aircraftspeed increases, because although the jet streamcontinues to issue at a high velocity from the engineits velocity relative to the surrounding atmosphere isreduced and, in consequence, the waste energy lossis reduced.

40. Briefly, propulsive efficiency may be expressedas:

or simply

Work done is the net thrust multiplied by the aircraftspeed. Therefore, progressing from the net thrustequation given in para. 18, the following equation isarrived at:Propulsive efficiency =



Fig. 21-8 The effect of air temperature on a typical twin-spool engine.

g2)Vv(W 2

J −

airflow engine to impartedEnergy aircraft the on done Work

exhaust in wasted work+ done Workdone Work










In the instance of an engine operating with a non-choked nozzle (Part 20), the equation becomes:

41. This latter equation can also be used for thechoked nozzle condition by using vj to represent thejet velocity when fully expanded to atmosphericpressure, thereby dispensing with the nozzlepressure term (P-P0)A.

42. Assuming an aircraft speed (V) of 375 m.p.h.and a jet velocity (vj) of 1,230 rn.p.h., the efficiencyof a turbo-jet is:

On the other hand, at an aircraft speed of 600 m.p.h.the efficiency is:

Propeller efficiency at these values of V is approxi-mately 82 and 55'per cent, respectively, and from



Fig. 21-9 Propulsive efficiencies and aircraft speed.










3752 =+



6002 =+


reference to fig. 21-9 it can be seen that for aircraftdesigned to operate at sea level speeds belowapproximately 400 m.p.h. it is more effective toabsorb the power developed in the jet engine bygearing it to a propeller instead of using it directly inthe form of a pure jet stream. The disadvantage ofthe propeller at the higher aircraft speeds is its rapidfall off in efficiency, due to shock waves createdaround the propeller as the blade tip speedapproaches Mach 1.0. Advanced propellertechnology, however, has produced a multi-bladed,swept back design capable of turning with tip speedsin excess of Mach 1.0 without loss of propellerefficiency. By using this design of propeller in acontra-rotating configuration, thereby reducing swirllosses, a 'prop-fan' engine, with very good propulsiveefficiency capable of operating efficiently at aircraftspeeds in excess of 500 m.p.h. at sea level, can beproduced.

43. To obtain good propulsive efficiencies withoutthe use of a complex propeller system, the by-passprinciple (Part 2) is used in various forms. With thisprinciple, some part of the total output is provided bya jet stream other than that which passes through theengine cycle and this is energized by a fan or avarying number of LP. compressor stages. Thisbypass air is used to lower the mean jet temperatureand velocity either by exhausting through a separatepropelling nozzle, or by mixing with the turbinestream to exhaust through a common nozzle.

44. The propulsive efficiency equation for a high by-pass ratio engine exhausting through separatenozzles is given below, where W1 and VJ1 relate tothe by-pass function and W2 and vJ2 to the enginemain function.

Propulsive efficiency =

By calculation, substituting the following values,which will be typical of a high by-pass ratio engine oftriple-spool configuration, it will be observed that apropulsive efficiency of approximately 85 per centresults.

V = 583 rn.p.h.W1 = 492 lb. per sec.W2 = 100 lb. per sec.VJ1 = 781 m.p.h.VJ2 = 812 m.p.h.

Propulsive efficiency can be further improved byusing the rear mounted contra-rotating fan configura-tion of the by-pass principle. This gives very high by-

pass ratios in the order of 15:1, and reduced 'drag'results due to the engine core being 'washed' by thelow velocity aircraft slipstream and not the relativelyhigh velocity fan efflux.

45. The improved propulsive efficiency of thebypass system bridges the efficiency gap betweenthe turbo-propeller engine and the pure turbo-jetengine. A graph illustrating the various propulsiveefficiencies with aircraft speed is shown in fig. 21-9.


46. Primary engine design considerations, particu-larly for commercial transport duty, are those of lowspecific fuel consumption and weight. Considerableimprovement has been achieved by use of the by-pass principle, and by advanced mechanical andaerodynamic features, and the use of improvedmaterials. With the trend towards higher by-passratios, in the range of 15:1, the triple-spool andcontra-rotating rear fan engines allow the pressureand by-pass ratios to be achieved with short rotors,using fewer compressor stages, resulting in a lighterand more compact engine.

47. S.f.c. is directly related to the thermal andpropulsive efficiencies; that is, the overall efficiencyof the engine. Theoretically, high thermal efficiencyrequires high pressures which in practice also meanshigh turbine entry temperatures. In a pure turbo-jetengine this high temperature would result in a highjet velocity and consequently lower the propulsiveefficiency (para. 40). However, by using the by-passprinciple, high thermal and propulsive efficienciescan be effectively combined by bypassing aproportion of the L.P. compressor or fan delivery airto lower the mean jet temperature and velocity asreferred to in para. 43. With advanced technologyengines of high by-pass and overall pressure ratios,a further pronounced improvement in s.f.c. isobtained.

48. The turbines of pure jet engines are heavybecause they deal with the total airflow, whereas theturbines of by-pass engines deal only with part of theflow; thus the H.P. compressor, combustionchambers and turbines, can be scaled down. Theincreased power per lb. of air at the turbines, to takeadvantage of their full capacity, is obtained by theincrease in pressure ratio and turbine entrytemperature. It is clear that the by-pass engine islighter, because not only has the diameter of the highpressure rotating assemblies been reduced but theengine is shorter for a given power output. With a low













Fig. 21-10 International Standard Atmosphere.

by-pass ratio engine, the weight reduction comparedwith a pure jet engine is in the order of 20 per centfor the same air mass flow.

49. With a high by-pass ratio engine of the triple-spool configuration, a further significant improvementin specific weight is obtained- This is derived mainlyfrom advanced mechanical and aerodynamic design,which in addition to permitting a significant reductionin the total number of parts, enables rotatingassemblies to be more effectively matched and towork closer to optimum conditions, thus minimizingthe number of compressor and turbine stages for a

given duty. The use of higher strength light-weightmaterials is also a contributory factor.

50. For a given mass flow less thrust is produced bythe by-pass engine due to the lower exit velocity.Thus, to obtain the same thrust, the by-pass enginemust be scaled to pass a larger total mass airflowthan the pure turbo-jet engine. The weight of theengine, however, is still less because of the reducedsize of the H.P. section of the engine. Therefore, inaddition to the reduced specific fuel consumption, animprovement in the power-to-weight ratio is obtained.



Rolls-Royce RB168 Mk202/Mk203

Rolls-Royce RB39 Clyde

Encouraged by results obtained from theTrent, Rolls-Royce decided to go ahead withan engine designed from the start as a turbo-prop. Named the Clyde it utilized the axialcompressor from the Metrovick F2 as firststage and a scaled up supercharger impellerfrom a Merlin as second stage. First running inAugust 1945 at 2000 shp, later enginesproduced up to 4200 shp.

22: ManufactureContents Page

Introduction 229Manufacturing strategy 230Forging 231Casting 233Fabrication 234Welding 235

Tungsten inert gas (T.I.G.) welding Electron beam welding (E.B.W.)

Electro-chemical machining (E.C.M.) 237

Stem drilling Capillary drilling

Electro-discharge machining (E.D.M.) 238 Composite materials andsandwich casings 240Inspection 240


1. During the design stages of the aircraft gasturbine engine, close liaison is maintained betweendesign, manufacturing, development and productsupport to ensure that the final design is a matchbetween the engineering specification and the man-ufacturing process capability.

2. The functioning of this type of engine, with itshigh power-to-weight ratio, demands the highestpossible performance from each component.Consistent with this requirement, each componentmust be manufactured at the lowest possible weightand cost and also provide mechanical integritythrough a long service life. Consequently, themethods used during manufacture are diverse andare usually determined by the duties eachcomponent has to fulfil.

3. No manufacturing technique or process that Inany way offers an advantage is ignored and mostavailable engineering methods and processes areemployed in the manufacture of these engines, Insome instances, the technique or process mayappear by some standards to be elaborate, timeconsuming and expensive, but is only adopted afterconfirmation that it does produce maximizedcomponent lives comparable with rig test achieve-ments.

4. Engine components are produced from a varietyof high tensile steel and high temperature nickel andcobalt alloy forgings. A proportion of components arecast using the investment casting process. Whilstfabrications, which form an increasing content, areproduced from materials such as stainless steel,titanium and nickel alloys using modern joining


techniques i.e., tungsten inert gas welding,resistance welding, electron beam welding and hightemperature brazing in vacuum furnaces.

5. The methods of machining engine componentsinclude grinding, turning, drilling, boring andbroaching whenever possible, with the more difficultmaterials and configurations being machined byelectro-discharge, electro-chemical, laser holedrilling and chemical size reduction.

6. Structural components i.e., cold spoiler, locationrings and by-pass ducts, benefit by considerableweight saving when using composite materials.

7. In addition to the many manufacturing methods,chemical and thermal processes are used on partfinished and finished components. These includeheat treatment, electro-plating, chromate sealing,chemical treatments, anodizing to prevent corrosion,chemical and mechanical cleaning, wet and dryabrasive blasting, polishing, plasma spraying, elec-trolytic etching and polishing to reveal metallurgicaldefects. Also a variety of barrelling techniques forremoval o! burrs and surface improvement. Mostprocesses are concerned with surface changes,

some give resistance to corrosion whilst others canbe used to release unwanted stress.

8. The main structure of an aero gas turbine engineis formed by a number of circular casings, ref. fig. 22-1, which are assembled and secured together byflanged joints and couplings located with dowels andtenons. These engines use curvic and hurthcouplings to enable accurate concentricity of matingassemblies which in turn assist an airline operatorwhen maintenance is required.


9. Manufacturing is changing and will continue tochange to meet the increasing demands ofaeroengine components for fuel efficiency, cost andweight reductions and being able to process thematerials required to meet these demands.

10. With the advent of micro-processors andextending the use of the computer, full automation ofcomponents considered for in house manufactureare implemented in line with supply groups manufac-turing strategy, all other components beingresourced within the world-wide supplier network.



Fig. 22-1 Arrangements of a triple-spool turbo-jet engine.

11. This automation is already applied in themanufacture of cast turbine blades with the sevencell and computer numerical controlled (C.N.C.)grinding centres, laser hardfacing and film coolinghole drilling by electro-discharge machining (E.D.M.).Families of turbine and compressor discs areproduced in flexible manufacturing cells, employingautomated guided vehicles delivering palletizedcomponents from computerized storage to C.N.C.machining cells that all use batch of one techniques.The smaller blades, with very thin airfoil sections, areproduced by integrated broaching and 360 degreeelectrochemical machining (E.C.M.) while inspectionand processing are being automated using thecomputer.

12. Tolerances between design and manufacturingare much closer when the design specification ismatched by the manufacturing proven capability.

13. Computer Aided Design (C.A.D.) and ComputerAided Manufacture (CAM.) provides an equivalentlink when engine components designed by C.A.D.can be used for the preparation of manufacturingdrawings, programmes for numerically controlledmachines, tool layouts, tool designs, operation

sequence, estimating and scheduling. Computersimulation allows potential cell and flow linemanufacture to be proven before physical machinepurchase and operation, thus preventing equipmentnot fulfilling their intended purpose.

14. Each casing is manufactured from the lightestmaterial commensurate with the stress and tempera-tures to which it is subjected in service. For example,magnesium alloy, composites and materials ofsandwich construction are used for air intakecasings, fan casings and low pressure compressorcasings, since these are the coolest parts of theengine. Alloy si eels are used for the turbine andnozzle casings where the temperatures are high andbecause these casings usually incorporate theengine rear mounting features. For casingssubjected to intermediate temperatures i.e. by-passduct and combustion outer casings, aluminium alloysand titanium alloys are used.


15. The engine drive shafts, compressor discs,turbine discs and gear trains are forged to as nearoptimum shape as is practicable commensurate withnon-destructive testing i.e., ultrasonic, magneticparticle and penetrant inspection. With turbine andcompressor blades, the accurately produced thinairfoil sections with varying degrees of camber andtwist, in a variety of alloys, entails a high standard ofprecision forging, ret. fig. 22-2. Neverthelessprecision forging of these blades is a recognisedpractice and enables one to be produced from ashaped die with the minimum of further work.



Fig. 22-2 Precision forging.



Fig. 22-3 Method of producing an engine component by sand casting.

16. The high operating temperatures at which theturbine discs must operate necessitates the use ofnickel base alloys. The compressor discs at the rearend of the compressor are produced from creep-resisting steels, or even nickel base alloys, becauseof the high temperatures to which they are subjected.The compressor discs at the front end of thecompressor are produced from titanium. The higherstrength of titanium at the moderate operating tem-peratures at the front end of the compressor,together with its lower weight provides a consider-able advantage over steel.

17. Forging calls for a very close control of thetemperature during the various operations. Anexceptionally high standard of furnace controlequipment, careful maintenance and cleanliness ofthe forging hammers, presses and dies, is essential.

18. Annular combustion rings can be cold forged toexacting tolerances and surfaces which alleviatesthe need for further machining before weldingtogether to produce the combustion casing.

19. H.P. compressor casings of the gas turbineengine are forged as rings or half rings which, whenassembled together, form the rigid structure of theengine. They are produced in various materials, i.e.,stainless steel, titanium and nickel alloys.


20. An increasing percentage of the gas turbineengine is produced from cast components using



Fig. 22-4 Automatic investment casting.

sand casting, ref. fig. 22-3, die casting andinvestment casting techniques; the latter becomingthe foremost in use because of its capability toproduce components with surfaces that require nofurther machining. It is essential that all castings aredefect free by the disciplines of cleanliness duringthe casting process otherwise they could causecomponent failure.

21. All casting techniques depend upon care withmethods of inspection such as correct chemicalcomposition, test of mechanical properties, radiolog-ical and microscopic examination, tensile strengthand creep tests.

22. The complexity of configurations together withaccurate tolerances in size and surface finish istotally dependent upon close liaison with design,manufacturing, metallurgist, chemist, die maker,furnace operator and final casting.

23. In the pursuit of ever increasing performance,turbine blades are produced from high temperature

nickel alloys that are cast by the investment castingor lost wax' technique. Directionally solidified andsingle crystal turbine blades are cast using thistechnique in order to extend their cyclic lives.

24. Figure 22-4 illustrates automatic casting used inthe production of equi-axed, directional solidified andsingle crystal turbine blades. The lost wax process isunparalleled in its ability to provide the higheststandards of surface finish, repeatable accuracy andsurface detail in a cast component. The increasingdemands of the engine has manifested itself in theneed to limit grain boundaries and provide complexinternal passages. The moulds used for directionalsolidified and single crystal castings differ from con-ventional moulds in that they are open at both ends,the base of a mould forms a socketed bayonet fittinginto which a chill plate is located during casting.Metal is introduced from the central sprue into themould cavities via a ceramic filter. These andorientated seed crystals, if required, are assembledwith the patterns prior to investment. Extensiveautomation is possible to ensure the wax patternsare coated with the shell material consistently byusing robots. The final casting can also have theirrises removed using elastic cut-off wheels drivenfrom robot arms, ref. fig. 22-5.


25. Major components of the gas turbine engine i.e.bearing housings, combustion and turbine casings,exhaust units, jet pipes, by-pass mixer units and lowpressure compressor casings can be produced asfabricated assemblies using sheet materials such asstainless steel titanium and varying types of nickelalloys.



Fig. 22-5 Robot cut-off

26. Other fabrication techniques for themanufacture of the low pressure compressor widechord fan blade comprise rolled titanium side panelsassembled in dies, hot twisted in a furnace and finallyhot creep formed to achieve the necessary configu-ration. Chemical milling is used to recess the centreof each panel which sandwiches a honeycomb core,both panels and the honeycomb are finally joinedtogether using automated furnaces where anactivated diffusion bonding process takes place, ref.fig. 22-6.


27. Welding processes are used extensively in thefabrication of gas turbine engine components i.e.,resistance welding by spot and seam, tungsten inertgas and electron beam are amongst the most widelyused today. Care has to be taken to limit thedistortion and shrinkage associated with thesetechniques.

Tungsten inert gas (T.I.G.) welding28. The most common form of tungsten inert gaswelding, fig, 22-7, in use is the direct current straightpolarity i.e., electrode negative pole. This is widelyused and the most economical method of producinghigh quality welds for the range of high strength/hightemperature materials used in gas turbine engines.For this class of work, high purity argon shielding gasis fed to both sides of the weld and the welding torchnozzle is fitted with a gas lens to ensure maximumefficiency for shielding gas coverage. A consumable



Fig. 22-6 Wide chord fan bladeconstruction.

Fig. 22-7 Typical tungsten inert gas welding details.

four per cent thoriated tungsten electrode, togetherwith a suitable non-contact method o! arc starting isused and the weld current is reduced in a controlledmanner at the end of each weld to prevent theformation of finishing cracks. All welds are visuallyand penetrant inspected and in addition, weldsassociated with rotating parts i.e., compressor and/orturbine are radiologically examined to QualityAcceptance Standards. During welding operationsand to aid in the control of distortion and shrinkagethe use of an expanding fixture is recommendedand, whenever possible, mechanised weldingemployed together with the pulsed arc technique ispreferred. A typical T.I.G. welding operation isillustrated in fig. 22-8.

Electron beam welding (E.B.W.)29. This system, which can use either low or highvoltage, uses a high power density beam ofelectrons to join a wide range of different materialsand of varying thickness. The welding machine ref.fig. 22-9, comprises an electron gun, optical viewingsystem, work chamber and handling equipment,vacuum pumping system, high or low voltage powersupply and operating controls. Many major rotatingassemblies for gas turbine engines are manufac-tured as single items in steel, titanium and nickelalloys and joined together i.e., intermediate and highpressure compressor drums. This technique allows



Fig. 22-8 Tungsten inert gas welding.

Fig. 22-9 Electron beam welding.

design flexibility in that distortion and shrinkage arereduced and dissimilar materials, to serve quitedifferent functions, can be homogeneously joinedtogether. For example, the H.P. turbine stub shaftsrequiring a stable bearing steel welded to a materialwhich can expand with the mating turbine disc.Automation has been enhanced by the application ofcomputer numerical control (C.N.C.) to the workhandling and manipulation. Seam tracking to ensurethat the joint is accurately followed and close loopunder bead control to guarantee the full depth ofmaterial thickness is welded. Focus of the beam iscontrolled by digital voltmeters. See fig. 22-10 forweld examples.


30. This type of machining employs both electricaland chemical effects in the removal of metal.Chemical forming, electro-chemical drilling and elec-trolytic grinding are techniques of electro-chemicalmachining employed in the production of gas turbineengine components.

31. The principle of the process is that when acurrent flows between the electrodes immersed in asolution of salts, chemical reactions occur in whichmetallic ions are transported from one electrode to

another (fig. 22-11). Faraday's law of electrolysisexplains that the amount of chemical reactionproduced by a current is proportional to the quantityof electricity passed.

32. In chemical forming, (fig. 22-11), the toolelectrode (the cathode) and the workpiece (theanode) are connected into a direct current circuit.Electrolytic solution passes, under pressure, throughthe tool electrode and metal is removed from thework gap by electrolytic action. A hydraulic ramadvances the tool electrodes into the workpiece toform the desired passage.

33. Electrolytic grinding employs a conductivewheel impregnated with abrasive particles. Thewheel is rotated close to the surface of theworkpiece, in such a way that the actual metalremoval is achieved by electro-chemical means. Theby-products, which would inhibit the process, areremoved by the sharp particles embodied in thewheel.

34. Stem drilling and capillary drilling techniquesare used principally in the drilling of small holes,usually cooling holes, such as required whenproducing turbine blades.



Fig. 22-10 Examples of T.I.G. and E.B. welds.

Stem drilling35. This process consists of tubes (cathode)produced from titanium and suitably insulated toensure a reaction at the tip. A twenty per centsolution of nitric acid is fed under pressure onto theblade producing holes generally in the region of0.026 in. diameter. The process is more speedy inoperation than electro-discharge machining and iscapable of drilling holes up to a depth two hundredtimes the diameter of the tube in use.

Capillary drilling36. Similar in process to stem drilling but usingtubes produced from glass incorporating a core ofplatinum wire (cathode). A twenty per cent nitric acidsolution is passed through the tube onto theworkpiece and is capable of producing holes assmall as 0.009 in. diameter. Depth of the hole is upto forty times greater than the tube in use andtherefore determined by tube diameter.

37. Automation has also been added to the processof electro-chemical machining (E.C.M.) with the intro-duction of 360 degree E.G. machining of smallcompressor blades, ref. fig. 22-12. For some bladesof shorter length airfoil, this technique is more costeffective than the finished shaped airfoil when usingprecision forging techniques. Blades produced byE.C.M. employ integrated vertical broachingmachines which take pre-cut lengths of bar material,produce the blade root feature, such as a fir-tree, andthen by using this as the location, fully E.C.M. fromboth sides to produce the thin airfoil section in oneoperation.


38. This type of machining removes metal from theworkpiece by converting the kinetic energy of electricsparks into heat as the sparks strike the workpiece.

39. An electric spark results when an electricpotential between two conducting surfaces reachesthe point at which the accumulation of electrons hasacquired sufficient energy to bridge the gap betweenthe two surfaces and complete the circuit. At thispoint, electrons break through the dielectric mediumbetween the conducting surfaces and, moving fromnegative (the tool electrode) to positive (theworkpiece), strike the latter surface with greatenergy; fig, 22-13 illustrates a typical spark erosioncircuit.

40. When the sparks strike the workpiece, the heatis so intense that the metal to be removed is instan-taneously vaporized with explosive results. Away



Fig. 22-11 Electro-chemical machining.

from the actual centre of the explosion, the metal istorn into fragments which may themselves be meltedby the intense heat. The dielectric medium, usuallyparaffin oil. pumped into the gap between the toolelectrode and the workpiece, has the tendency toquench the explosion and to sweep away metallicvapour and molten particles.

41. The amount of work that can be effected in thesystem is a function of the energy of the individualsparks and the frequency at which they occur.

42. The shape of the tool electrode is a mirror imageof the passage to be machined in the workpiece and,to maintain a constant work gap, the electrode is fedinto the workpiece as erosion is effected.



Fig. 22-12 Typical automated manufacture of compressor blades.

Fig. 22-13 Electro-discharge machiningcircuit.


43. High power to weight ratio and low componentcosts are very important considerations in the designof any aircraft gas turbine engine, but when thefunction of such an engine is to support a verticaltake-off aircraft during transition, or as an auxiliarypower unit, then the power to weight ratio becomesextremely critical.

44. In such engines, the advantage of compositematerials allows the designer to produce structuresin which directional strengths can be varied bydirectional lay-up of fibres according to the appliedloads.

45. Composite materials have and will continue toreplace casings which, in previous engines, wouldhave been produced in steels or titanium. By-passduct assemblies comprising of three casings arecurrently being produced up to 4ft-7in. in diameterand 2ft-0in in length using pre-cured compositematerials for the casing fabric. Flanges and mountingbosses are added during the manufacturing process,which are then drilled for both location and machined

for peripheral feature attachment on C.N.C.machining centres, which at one component load,completely machine all required features. Examplesof composite material applications are illustrated infig. 22-14.

46. Conventional cast and fabricated casings andcowlings are also being replaced by casings ofsandwich construction which provide strength alliedwith lightness and also act as a noise suppressionmedium. Sandwich construction casings comprise ahoneycomb structure of aluminium or stainless steelinterposed between layers of dissimilar material. Thematerials employed depend upon the environment inwhich they are used.


47. During the process of manufacture, componentparts need to be inspected to ensure defect freeengines are produced. Using automated machineryand automated inspection, dimensional accuracy ismaintained by using multi-directional applied probesthat record sizes and position of features. The C.N.C.inspection machine can inspect families ofcomponents at pre-determined allotted intervals



Fig. 22-14 Some composite material applications.

without further operator intervention. In the chipmachining (i.e., turning, boring, milling etc.) andmetal forming processes C.N.C. machine toolsenable consistency of manufacture which can be sta-tistically inspected i.e., one in ten. Component

integrity is achieved by use of ultrasonic, radiologi-cal, magnetic particle and penetrant inspectiontechniques, as well as electrolytic and acid etching toensure all material properties are maintained to bothlaboratory and quality acceptance standards.



Fig. 22-15 Advanced integrated manufacturing system (A.I.M.S.).

Rolls-Royce Turbomeca RTM 332 Turboshaft

Rolls-Royce RB 93 Soar

Developed as a lightweight expendableengine for winged missiles, the Soar first ranin 1952. At that time it had the best thrust toweight ratio of any gas turbine in the world,producing 1810 lb thrust from only 275 lb totalweight. The Soar was flight tested in a GlosterMeteor with one engine on each wingtip. Itwas also built under licence in the USA as theJ81 for the XQ4 supersonic drone.

23: Power plant installation

Contents Page

Introduction 243Power plant location 243Air intakes 245Engine and jet pipe mountings 248Accessories 249Cowlings 249


1. When a gas turbine engine is installed in anaircraft it usually requires a number of accessoriesfitting to it and connections made to various aircraftsystems. The engine, jet pipe and accessories, andin some installations a thrust reverser, must besuitably cowled and an air intake must be providedfor the compressor, the complete installation formingthe aircraft power plant.


2. The power plant location and aircraft configura-tion are of an integrated design and this dependsupon the duties that the aircraft has to perform.Turbo-jet engine power plants may be in the form ofpod installations that are attached to the wings bypylons (fig. 23-1), or attached to the sides of the rearfuselage by short stub wings (fig. 23-2), or they maybe buried in the fuselage or wings. Some aircrafthave a combination of rear fuselage and tail-mounted power plants, others, as shown in fig. 23-3,have wing-mounted pod installations with a thirdengine buried in the tail structure. Turbo-propellerengines, however, are normally limited to installationin the wings or nose of an aircraft.


Power plant installation


Fig. 23-1 Wing-mounted pod installation.

Fig. 23-2 Fuselage - mounted pod installation.

Fig. 23-3 Tail and wing-mounted pod installation.

3. The position of the power plant must not affectthe efficiency of the air intake, and the exhaust gasesmust be discharged clear of the aircraft and itscontrol surfaces. Any installation must also be suchthat it produces the minimum drag effect.

4. Power plant installations are numbered from leftto right when viewed from the rear of the aircraft.

5. Supersonic aircraft usually have the power plantsburied in the aircraft for aerodynamic reasons.Vertical lift aircraft can use either the buried installa-tion or the podded power plant, or in some instancesboth types may be combined in one aircraft (Part 18).


6. The main requirement of an air intake is that,under all operating conditions, delivery of-the air tothe engine is achieved with the minimum loss ofenergy occurring through the duct. To enable thecompressor to operate satisfactorily, the air mustreach the compressor at a uniform pressuredistributed evenly across the whole inlet area.

7. The ideal air intake for a turbo-jet engine fitted toan aircraft flying at subsonic or low supersonicspeeds, is a short, pitot-type circular intake (fig. 23-4). This type of intake makes the fullest use of theram effect on the air due to forward speed, andsuffers the minimum loss of ram pressure withchanges of aircraft attitude. However, as sonic speed

is approached, the efficiency of this type of air intakebegins to fall because of the formation of a shockwave at the intake lip.

8. The pitot-type intake can be used for engines thatare mounted in pods or in the wings, although the lattersometimes require a departure from the circular cross-section because of the wing thickness (fig. 23-5).

Power plant installation


Fig. 23-5 Wing leading edge intakes.

Fig. 23-4 Pitot-type intake.

9. Single engined aircraft sometimes use a pilot-type intake; however, because this generally involvesthe use of a long duct ahead of the compressor, adivided type of intake on each side of the fuselage isoften used (fig. 23-6).

10. The disadvantage of the divided type of airintake is that when the aircraft yaws, a loss of rampressure occurs on one side of the intake, as shownin fig. 23-7, causing an uneven distribution of airflowinto the compressor.

Power plant installation


Fig. 23-6 Single engined aircraft with fuselage intakes.

Fig. 23-7 Loss of ram pressure in divided intakes.

11. At higher supersonic speeds, the pitot type of airintake is unsuitable due to the severity of theshockwave that forms and progressively reduces theintake efficiency as speed increases. A more suitabletype of intake for these higher speeds is known asthe external/internal compression intake (fig. 23-8).This type of intake produces a series of mild shockwaves without excessively reducing the intakeefficiency.

12. As aircraft speed increases still further, so alsodoes the intake compression ratio and, at high Machnumbers, it is necessary to have an air intake that hasa variable throat area and spill valves to accommodate

and control the changing volumes of air (fig. 23-9).The airflow velocities encountered in the higher speedrange of the aircraft are much higher than the enginecan efficiently use; therefore, the air velocity must bedecreased between the intake and the engine air inlet.The angle of the variable throat area intake automati-cally varies with aircraft speed and positions the shockwave to decrease the air velocity at the engine inletand maintain maximum pressure recovery within theinlet duct. However, continued development enablesthis to be achieved by careful design of the intake andducting. This, coupled with auxiliary air doors to permitextra air to be taken in under certain engine operatingconditions, allows the airflow to be controlled withoutthe use of variable geometry intakes. The fuselageintakes shown in fig. 23-10 are of the variable throatarea type.

Power plant installation


Fig. 23-8 External/internal compressionintake.

Fig. 23-9 Variable throat area intake.


13. The engine is mounted in the aircraft in amanner that allows the thrust forces developed bythe engine to be transmitted to the aircraft mainstructure, in addition to supporting the engine weightand carrying any flight loads. Because of the widevariations in the temperature of the engine casings,

the engine is mounted so that the casings canexpand freely in both a longitudinal and a radialdirection. Types of engine mountings, however, varyto suit the particular installation requirement. Turbo-jet engines are usually either side mounted orunderslung as illustrated in fig. 23-11. Turbo-propeller engines are mounted forward on a tubularframework as illustrated in fig. 23-13.

Power plant installation


Fig. 23-10 Fuselage intakes.

Fig. 23-11 Typical turbo-jet engine mountings.

14. The jet pipe is normally attached to the rear ofthe engine and supported by the engine mountings.In some installations, particularly where long jetpipes are employed, an additional mounting isprovided, usually in the form of small rollers attachedto each side of the jet pipe. The rollers locate inairframe-mounted channels and support the weightof the jet pipe, whilst still allowing it to freely expandin a longitudinal direction.


15. An aircraft power plant installation generallyincludes a number of accessories that are electrical-ly operated, mechanically driven or driven by highpressure air.

16. Electrically operated accessories such asengine control actuators, amplifiers, air controlvalves and solenoids, are supplied with power fromthe aircraft electrical system or an engine drivendedicated electrical generator.

17. Mechanically driven units, such as generators,constant speed drive units, hydraulic pumps, low andhigh pressure fuel pumps, and engine speedsignalling, measuring or governing units are drivenfrom the engine through internal and externalgearboxes (Part 7).

18. Air-driven accessories, such as the air starterand possibly the thrust reverser, afterburner and

water injection pumps, are driven by air tapped fromthe engine compressor. Air conditioning and cabinpressurization units may have a separate air-drivencompressor or use air direct from the enginecompressor. The amount of air that is taken for allaccessories and services must always be a verysmall percentage of the total airflow, as it representsa thrust or power loss and an increase in specific fuelconsumption.


19. Access to an engine mounted in the wing orfuselage is by hinged doors; on pod and turbo-propeller installations the main cowlings are hinged.Access for minor servicing is by small detachable orhinged panels. All fasteners are of the quick-releasetype.

20. A turbo-propeller engine, or a turbo-jet enginemounted in a pod, is usually far more accessible thana buried engine because of the larger area of hingedcowling that can be provided. The accessibility of apodded turbo-fan engine is shown in fig, 23-12 andthat of a turbo-propeller engine is shown in fig, 23-13.

Power plant installation


Fig. 23-12 Engine accessibility, turbo-fan engine.

Fig. 23-13 Engine accessibility, turbo-propeller engine.

I.A.E. International Aero Engines V2500

Rolls-RoyceRB 162

Design of the RB162 began in 1959 usingexperience, gained on the RB108, ofsimplified lightweight constructions andsystems. These measures, combined withlightweight materials, served to keep theengine weight down to 280 lb; giving a thrustto weight ratio of 16:1. First run in December1961, the RB162 was used to provide lift for avariety of VTOL research aircraft.

24: Maintenance

Contents Page

Introduction 251 On-wing maintenance 252

Scheduled maintenance Unscheduled maintenance

Condition monitoring 252Flight deck indicators In-flight recorders Ground indicators

Maintenance precautions 254Trouble shooting 254Adjustments 256Ground testing 256


1. Maintenance covers both the work that isrequired to maintain the engine and its systems in anairworthy condition while installed in an aircraft (on-wing or line maintenance) and the work required toreturn the engine to airworthy condition whenremoved from an aircraft (overhaul or shopmaintenance). On-wing maintenance is covered inthis Part; overhaul is covered in Part 25.

2. Because many aspects of maintenance aresubject to the approval of a recognized authority, itshould be fully understood that the information givenin this Part is of a general nature and is not intendedas a substitute for any official instructions.

3. The comprehensive instructions covering theactual work to be done to support scheduledmaintenance (para. 8) and unscheduled maintenance(para. 10) are contained in the aircraft maintenancemanual. Both this publication, and the aircraftmaintenance schedule mentioned in para. 8, arebased on manufacturers' recommendations and areapproved by the appropriate airworthiness authority.

4. The maximum time an engine can remaininstalled in an aircraft (engine life) is limited to a fixedperiod agreed between the engine manufacturer andairworthiness authority. On some engines this periodis referred to as the time between overhaul (T.B.O.)and on reaching it the engine is removed forcomplete overhaul.


5. Because the T.B.O. is actually determined by thelife of one or two assemblies within the engine,during overhaul, it is generally found that the otherassemblies are mechanically sound and fit tocontinue in service for a much longer period.Therefore, with the introduction of modular enginesand the improved inspection and monitoringtechniques available, the T.B.O. method on limitingthe engine's life on-wing has been replaced by the'on-condition' method.

6. Basically this means that a life is not declared for the totalengine but only for certain parts of the engine. On reachingtheir life limit, these parts are replaced and the enginecontinues in service, the remainder of the engine beingoverhauled 'on condition', Modular constructed engines areparticularly suited to this method, as the module containinga life limited part can be replaced by a similar module andthe engine returned to service with minimum delay, Themodule is then disassembled for life limited partreplacement, repair or complete overhaul as required.


7. On-wing maintenance falls into two basiccategories: scheduled maintenance andunscheduled maintenance.

Scheduled maintenance8. Scheduled maintenance embraces the periodicand recurring checks that have to be effected inaccordance with the engine section of theappropriate aircraft maintenance schedule. Thesechecks range from transit items, which do notnormally entail opening cowls, to more elaboratechecks within specified time limits, usually calculatedin aircraft flying hours and phased with the aircraftcheck cycle.

9. Continuous 'not-exceed-limit' maintenance,whereby checks are carried out progressively and asconvenient within given time limits rather than atspecific aircraft check periods, has been widelyadopted to supersede the check cycle. With theprogressive introduction of condition monitoringdevices (para. 11) of increased efficiency andreliability, a number of traditionally acceptedscheduled checks may become unnecessary.Extracts from a typical maintenance schedule areshown in fig. 24-1.

Unscheduled maintenance10. Unscheduled maintenance covers work neces-sitated by occurrences that are not normally relatedto time limits, e.g. bird ingestion, a strike by lightning,a crash or heavy landing, Unscheduled work

required may also result from malfunction, troubleshooting, scheduled maintenance, and occasionally,manufacturers' specific recommendations. This typeof maintenance usually involves rectificationadjustment or replacement.


11. Condition monitoring devices must giveindication of any engine deterioration at the earliestpossible stage and also enable the area or module inwhich deterioration is occurring to be identified. Thisfacilitates quick diagnosis, which can be followed byscheduled monitoring and subsequent programmedrectification at major bases, thereby avoiding in-flightshut-down, with resultant aircraft delay, andminimizing secondary damage. Monitoring devicesand facilities can be broadly categorized as flight deckindicators, in-flight recorders and ground indicators.

Flight deck indicators12. Flight deck indicators are used to monitorengine parameters such as thrust or power, r.p.m.,turbine gas temperature, oil pressure and vibration.Most of the indicators used are described in Part 12.Other devices, however, may be used and theseinclude:

Accelerometers for more reliable and precisevibration monitoring.Radiation pyrometers for direct measurement ofturbine blade temperature.Return oil temperature indicators.Remote indicators for oil tank content.Engine surge or stall detectors.Rub indicators to sense eccentric running ofrotating assemblies.

In-flight recorders13. Selected engine parameters are recorded, eithermanually or automatically, during flight. Therecordings are processed and analyzed for significanttrends indicative of the commencement of failure. Anin-flight recording device that may be used is thetime/temperature cycle recorder. The purpose of thisdevice is to accurately record the engine time spentoperating at critical high turbine gas temperatures,thus providing a more realistic measure of 'hot-end'life than that provided by total engine running hours.

14. Automatic systems (Part 12) known as aircraftintegrated data systems (A.I.D.S.) are able to recordparameters additional to those normally displayede.g. certain pressures, temperatures and flows.

15. Many of the electronic components used inmodern control systems have the ability to monitor



their own and associated component operation. Anyfault detected is recorded in its built-in memory forsubsequent retrieval and rectification by the groundcrew. On aircraft that feature electronic engineparameter flight deck displays (Part 12) certain faultsare also automatically brought to the flight crew'sattention.

Ground indicators16. The devices used or checked on the ground, asdistinct from those used or checked in flight, mayconveniently be referred to as ground indicators; thistitle is also taken to embrace instruments used forengine internal inspection.

17. Internal viewing instruments can be eitherflexible or rigid, designed either for end or angledviewing and, in some instances adaptable for still or

video photography which may be linked to closedcircuit television. These instruments are used forexamining and assessing the condition of thecompressor and turbine assemblies, nozzle guidevanes (fig. 24-2) and combustion system, and can beinserted through access ports located at strategicpoints in the engine main casings.

18. The engine condition indicators includemagnetic chip detectors, oil filters and certain fuelfilters. These indicators are frequently used to sub-stantiate indications of failures shown by flight deckmonitoring and in-flight recordings. For instance,inspection of the oil filters and chip detectors canreveal deposits from which experienced personnelcan recognize early signs of failure. Somemaintenance organizations progressively log oil filter



Fig. 24-1 A typical maintenance schedule (extracts).

and magnetic chip detector history and catalogue theyield of particles. Fuel filters may incorporate a silverstrip indicator that detects any abnormal concentra-tion of sulphur in the fuel.


19. During engine maintenance, it is necessary toobserve certain precautions. The ignition system ispotentially lethal and, therefore, before any work isdone on the high energy ignition units, igniter plugsor harness, the low tension supply to the units mustbe disconnected and at least one minute allowed toelapse before disconnecting the high tension lead.Similarly, before carrying out work on unitsconnected to the electrical system, the system mustbe made safe, either by switching off power or bytripping and tagging appropriate circuit breakers.With some installations, the isolation of certainassociated systems may be required.

20. When the oil system is being replenished, caremust be taken that no oil is spilt. If any oil is acciden-tally spilt, it should be cleaned off immediately as it isinjurious to paintwork and to certain rubbercompounds such as could be found in the electricalharnesses, Oil can also be toxic through absorptionif allowed to come into contact with the human skinfor prolonged periods. Care should be taken not tooverfill the oil system; this may easily occur if theaircraft is not on level ground or if the engine hasbeen stationary for a long period before the oil levelis checked.

21. Before an inspection of the air intake or exhaustsystem is made it must be ascertained that there Isno possibility of the starter system being operated orthe ignition system being energized.

22. A final inspection of the engine, air intake andexhaust system, must always be made after anyrepair, adjustment or component change, to ensurethat no loose items, no matter how small, have beenleft inside. Unless specific local instructions ruleotherwise, air intake and exhaust blanks or coversshould be fitted when engines are not running.


23. The procedure for locating a fault is commonlyreferred to as trouble shooting, and the requirementunder this procedure is for quick and accuratediagnosis with the minimum associated work and theprevention of unnecessary unit or engine removals.

24. The basic principle of effective trouble shootingis to clearly define and interpret the reportedsymptom and then proceed to a logical andsystematic method of diagnosis (fig. 24-3).



Fig. 24-2 Inspection of H.P. nozzle guidevanes.

Fig. 24-3 Trouble shooting - logicalsequence.

25. The reported symptom will frequently originatefrom flight deck instrument readings and, unless it isapparent from supporting information that thereadings are genuine, instrumentation should bechecked before proceeding further. Similarly, quick

elimination checks should normally be undertakenbefore more involved tasks. The manufacturers'maintenance manual contains trouble shootinginformation, usually in chart form and fig. 24-4 showsa typical example.



Fig. 24-4 A typical trouble shooting chart.

26. The progressive introduction of improved andmore reliable condition monitoring devices (para. 11)will have considerable influence on accepted troubleshooting practice, since to a large extent thesedevices are designed to pin-point, at an early stage,the specific system or assembly at fault. Thedevelopment of suitable test sets could eventuallyeliminate the need for engine ground testing aftertrouble shooting


27. There are usually some adjustments that can bemade to the engine controlling the fuel trimmingdevices. Typical functions for which adjustmentprovision is normally made include idling andmaximum r.p.m., acceleration and decelerationtimes, and compressor air bleed valve operation.

28. Adjustment of an engine should be made only ifit is quite certain that no other fault exists that couldbe responsible for the particular condition, Themaintenance manual instructions relative to theadjustment must be closely adhered to at all times. Inmany instances, subject to local instructions, aground adjustment can be made with the enginerunning.

29. Adjusters are usually designed with some formof friction locking (fig. 24-5) that dispenses withlocknuts, lockplates and locking wire. On someengines, provision is also made for fitting remoteadjustment equipment (fig. 24-6) that permitsadjustment to be made during ground test with thecowls closed, the adjustment usually being madefrom the flight deck.


30. The basic purpose of engine ground testing is toconfirm performance and mechanical integrity and tocheck a fault or prove a rectification during troubleshooting. Ground testing is essential after engineinstallation, but scheduled ground testing may notnormally be called for where satisfactory operationon the last flight is considered to be the authority oracceptance for the subsequent flight. In someinstances, this is backed up by specific checks madein cruise or on approach and, of course, by evidencefrom flight deck indicators and recordings.

31. For economic reasons and because of thenoise problem, ground testing is kept to a minimum

and is usually only carried out after engine installa-tions, during trouble shooting, or to test an aircraftsystem. With the improved maintenance methodsand introduction of system test sets which simulaterunning conditions during the checking of a staticengine, the need for ground testing, particularly athigh power, is becoming virtually unnecessary.

32. Before a ground test is made, certainprecautions and procedures must be observed toprevent damage to the engine or aircraft and injury topersonnel.

33. Because of the mass of air that will be drawninto the intake and the resultant high velocity andtemperature of the exhaust gases during a groundtest, danger zones exist at the front and rear of theaircraft. These zones will extend for a considerabledistance, and atypical example is shown in fig. 24-7.The jet efflux must be clear o! buildings and otheraircraft. Personnel engaged in ground testing mustensure that any easily detachable clothing issecurely fastened and should wear acoustic earmuffs.



Fig. 24-5 Typical friction locked adjusters.



Fig. 24-6 Remote adjustment equipment fitted to a turbo-propeller engine.

34. The aircraft should be headed into wind andpositioned so that the air intake and exhaust are overfirm concrete, or a prepared area that is free fromloose material and loose objects, and clear ofequipment. Where noise suppression installationsare used, the aircraft should be positioned in

accordance with local instructions. When verticaltake-off aircraft are being tested, protective steelplates and deflectors may be used to prevent grounderosion and engine ingestion of exhaust gases anddebris. Aircraft wheels should be securely chockedand braked; with vertical take-off aircraft, anchoring



Fig. 24-7 Ground running danger zones.

or restraining devices are also used. Adequate firefighting equipment must be readily available andlocal fire regulations must be strictly enforced.

35. Before an engine is started, the air intake andjet pipe must be inspected to ensure that they arefree from any debris or obstruction. Each operatorwill detail his individual pre-start inspection require-ments; a typical example of this for a multi-enginedaircraft is shown in fig. 24-8.

36. The starting drill varies between differentaircraft types and a starting check procedure isnormally used. Generally, all non-essential systems

are switched or selected off; warning and emergencysystems are checked when applicable. Finally, afterensuring that the low pressure fuel supply is selectedon, the starting cycle is initiated.

37. At a predetermined point during the startingcycle, the high pressure fuel shut-off valve (cock) isopened to allow fuel to pass to the fuel spraynozzles, this point varying with aircraft and enginetype; on some installations the shut-off valve may beopened before the starting cycle is initiated. Duringthe engine light-up period and subsequent accelera-tion to idling speed, the engine exhaust gastemperature must be carefully monitored to ensure



Fig. 24-8 A pre-start inspection sequence.

that the maximum temperature limitation is notexceeded. If the temperature limitation appears likelyto be exceeded, the shut-off valve must be closedand the starting cycle cancelled; the cause andpossible effect of the high temperature must then beinvestigated before the engine is again started.

38. When a turbo-propeller engine is being started,the propeller must be set to the correct starting pitch,as recommended by the engine manufacturer. Toprovide the minimum resistance to turning and thusprevent an excessive exhaust gas temperatureoccurring during the starting cycle, some propellershave a special fine pitch setting.

39. Throttle movements should be kept to aminimum and be smooth and progressive to avoidthermal stresses associated with rapid changes intemperature. Rapid throttle movements to check theacceleration and deceleration capabilities of theengine should be made only after all other majorchecks have proved satisfactory and after someslower accelerations and decelerations have provedsuccessful.

40. Before an engine is stopped, it should normallybe allowed to run for a short period at idling speed toensure gradual cooling of the turbine assembly. Theonly action required to stop the engine is the closing



Fig. 24-9 Overheated turbine blades.

of the shut-off valve. The shut-off valve must not bere-opened during engine rundown, as the resultingsupply of fuel can spontaneously ignite withconsequent severe overheating of the turbineassembly. An example of turbine blades which havebeen subjected to overheating is shown in fig. 24-9.

41. The time taken for the engine to come to restafter the shut-off valve is closed is known as the'rundown time' and this can give an indication of anyrubbing inside the engine. However it should beborne in mind that variations in wind velocity anddirection may affect the run-down time of an engine.



Rolls-Royce Tyne

Bristol BE 25 Orion

The Orion was a two-spool turbo-propdesigned to operate at its full rated power to20,000ft, achieved by throttling its sea levelpower to a maximum of 5150 ehp. Flighttesting commenced in August 1956 with theengine installed in the port outer nacelle of theBristol Britannia. Development was disconti-nued owing to lack of demand for turbo-propsat this time.

25: Overhaul


1. It is most important that the cost of maintainingan engine in service is considered at the designstage. All aspects of engine repairability are alsoconsidered, both to reduce the requirement foroverhaul or repair and to avoid, where possible,designs which make repairs difficult to effect. Engineconstruction must allow the operator to complete theoverhaul or repair work as quickly and cheaply aspossible.

2. In service, the engine is inspected at routineperiods based on manufacturers' recommendationsand agreed between the operator and the relevantairworthiness authority. The engine is removed fromthe aircraft when it fails, during these inspections, tomeet the specified standards. This concept is a formof 'on-condition' monitoring, reference para. 9,however, regardless of condition, some engines areremoved when a stipulated number of engine flying

hours have been achieved, this concept is known astime between overhauls (T.B.O.). Operators will oftenremove engines in order to acquire 'fleet stagger'thus preventing a situation when an unacceptablenumber of engines require removal at the sameperiod of time.

3. The length of time between overhauls varies con-siderably between different engine types, beingestablished as a result of discussions between theoperator, the airworthiness authority and the manu-facturer, when such considerations as the totalexperience gained with the particular engine series,the type of operation, the utilization, and sometimesclimatic conditions, are taken into account. Inimproving the overhaul period the airworthinessauthority may take into consideration the backgroundof the operator, his servicing facilities and theexperience of his maintenance personnel.

4. When a new type of engine enters service,sampling (i.e. engine removal, dismantling andinspection) may be called for at a modest life. Thesampling will be continued until the life at which theengine should be overhauled is indicated by thecondition of the sample engines or by its reliabilityrecord in service. In some instances, the ultimate life

Contents Page

Introduction 263 Overhaul/Repair 265

DisassemblyCleaningInspectionRepairBalancingMoment weighing of bladesAssemblingTestingPreparing for storage/despatch


obtained may be two, three, or even four times theoriginal period permitted. The development of theT.B.O, from the introduction of an engine into service,through several years of operation, is shown as anexample in fig, 25-1.

5. Among the main factors affecting the overhaulperiod for an engine is the use to which it is put inservice. For example, a military engine will generallyhave a much lower T.B.O. than its civil counterpart,as performance capability is the operating criterionrather than economics. Due to the effect of rapidtemperature changes in the hot parts of the engine,the most arduous treatment is the frequent changingof power output to which short-haul transports andfighter aircraft are subjected.

6. When aircraft are based in areas with exception-ally high humidity or salt content in the atmosphere,there exists the added danger of corrosion, resultingin the need for more frequent overhauls. In peacetime, some military aircraft have a very low utilization,this introduces the additional problem of certainmaterials used in its construction deteriorating beforethe engine has otherwise reached a condition whichwould normally require an overhaul. Elapsed time, aswell as flying hours, would then influence theoverhaul period.

7. In addition to scheduled overhauls, there areproblems that arise from damage and defects. Aproportion of these, which are uneconomic orimpractical to rectify in the aircraft, necessitateunscheduled removals and require the engine to bereturned to an engineering base or an overhaul shopfor partial or complete overhaul.

8. The purpose of overhaul is to restore an engineenabling it to complete a further life by complyingwith new engine performance acceptance limitationsand maintaining the same reliability. This is achievedby dismantling the engine in order that parts can beinspected for condition and to determine the need forrenewal or repair of those parts whose deteriorationwould reduce the performance, or would not remainin a serviceable condition until the next overhaul.

9. The design of the modular constructed engine(Part 22, fig. 22-1) makes it particularly suited to adifferent technique of overhaul/repair. This techniqueis based on 'on-condition' monitoring (Part 24). Thismeans that a life is not declared for the total enginebut only certain parts of the engine. On reaching theirlife limit, these parts are replaced and the enginereturned to service, the remainder of the enginebeing overhauled 'on-condition'.



Fig. 25-1 Example of growth of time between overhaul (T.B.O.).

10. Modular construction, together with associatedtooling, enables the engine to be disassembled intoa number of major assemblies (modules). Moduleswhich contain life limited parts can be replaced bysimilar assemblies and the engine returned toservice with minimum delay. The removed modulesare disassembled into mini-modules for life limitedpart replacement, repair or complete overhaul asrequired.


11. The high cost of new engines has a consider-able influence on the overhaul/repair arrangements,as the number of spare engines normally bought bythe operator is kept to an absolute minimum. Thismeans that an unserviceable engine must be quicklyrestored to serviceability by changing a module, or apart if the modular construction will permit it, or bycareful scheduling of planned removals for overhaulsat time expiry. This scheduling, through theworkshop, of engines or modules that require the useof specialized equipment for repair is important, bothto keep the flow of work even and to staggerremovals to avoid aircraft being grounded byshortage of serviceable engines or modules.

12. Because the work that is to be implementedmust be planned and subsequently recorded, theengine or module is received in the workshop withdocuments to show its modification standard and itsreason for rejection from service. The planning willinclude a list of the modifications that can or must beincorporated to improve engine reliability orperformance or to reduce operating costs.

13. The layout of the overhaul/repair workshop isdesigned to facilitate work movement through thecomplete range of operations, to achieve maximumutilization of floor space and to allow specialequipment to be sited in positions that will suit thegeneral flow pattern. All these considerations areaimed at achieving a quick turnround of engines. Asan example of how shop layouts may be planned, atypical arrangement is shown in fig, 25-2.

Disassembly14. The engine can be disassembled in the verticalor horizontal position. When it is disassembled in thevertical position, the engine is mounted, usually frontend downwards, on a floor-fixture stool or a ram-topfixture. To enable it to be disassembled horizontally,the engine is mounted in a special turnover stand.

15. When the floor-fixture stool is used, thepersonnel use a mobile work platform to raisethemselves to a reasonable working position aroundthe engine. When the ram-top fixture is used, the ramand engine are retracted into a pit, so enabling theworkmen to remain at floor level.

16. The engine is disassembled into main sub-assemblies or modules, which are fitted in trans-portation stands and despatched to the separateareas where they are further disassembled toindividual parts. The individual parts are conveyed insuitable containers to a cleaning area in preparationfor inspection.

Cleaning17. The cleaning agents used during overhaulrange from organic solvents to acid and otherchemical cleaners, and extend to electrolyticcleaning solutions.

18. Organic solvents include kerosine for washing,trichloroethane for degreasing and paint strippingsolutions which can generally be used on themajority of components for carbon and paintremoval. The more restricted and sometimes rigidlycontrolled acid and other chemical cleaners are usedfor corrosion, heat scale and carbon removal fromcertain components. To give the highest degree ofcleanliness to achieve the integrity of inspection thatis considered necessary on certain major rotatingparts, such as turbine discs, electrolytic cleaningsolutions are often used.

19. Aircraft which operate at high altitudes canbecome contaminated with radio-active particlesheld in the atmosphere, this radio-activity is retainedin the dirt and carbon deposits in the engine.

20. If contamination is suspected the radio-activitylevel of the engine must be determined to ensure thelimitations agreed by the health authorities are notexceeded, Evidence of contamination will entailadditional cleaning in a designated region, separatefrom the overhaul area, to safeguard the health ofpersonnel in the workshop. Arrangements have tomade with the health authorities for disposal of thewaste radio-active cleaning material.

Inspection21. After cleaning, and prior to inspection, thesurfaces of some parts, e.g. turbine discs, areetched. This process removes a small amount ofmaterial from the surface of the part, leaving an even





Fig. 25-2 Typical overhaul workshop layout.

matt finish which reveals surface defects that cannotnormally be seen with the naked eye. The metalremoval is normally achieved either by an electrolyt-ic process in which the part forms the anode, or byimmersing the part for a short time in a special acidbath. Both methods must be carefully controlled toavoid the removal of too much material.

22. After the components have been cleaned theyare visually and, when necessary, dimensionallyinspected to establish general condition and thensubjected to crack inspection. This may includebinocular and magnetic or penetrant inspectiontechniques, used either alone or consecutively,depending on the components being inspected andthe degree of inspection considered necessary.

23. The non-dimensional inspections can bedivided into visual examination for general conditionand inspection for cracks. The visual examinationdepends on the inspector's judgement, based onexperience and backed by guidance from the manu-facturer. Although the visual examination of manyparts of the engine conform to normal engineeringpractice, for some parts the acceptance standardsare specialized, for example, the combustionchambers, which are subjected to very high temper-atures and high speed airflows in service.

24. Dimensional inspection consists of measuringspecific components to ensure that they are withinthe limits and tolerances laid down and known as'Fits and Clearances'. Some of the components aremeasured at each overhaul, because only a smallamount of wear or distortion is permissible or toenable the working clearances with matingcomponents to be calculated. Other components aremeasured only when the condition found duringvisual inspection requires dimensional verification.The tolerances laid down for overhaul, supported byservice experience, are often wider than those usedduring original manufacture.

25. The detection of cracks that are not normallyvisible to the naked eye is most important, particular-ly on major rotating parts such as turbine discs, sincefailure to detect them could result in crackpropagation during further service and eventuallylead to component failure. Various methods ofaccentuating these are used for inspection, the twoprincipal techniques being penetrant inspection fornonmagnetic materials and electro-magneticinspection for those parts that can be magnetized.

26. Two forms of penetrant inspection in commonuse are known as the dye penetrant and the

fluorescent test. With the dye test, a penetratingcoloured dye is induced to enter any cracks or poresin the surface of the part. The surface is then washedand a developer fluid containing white absorbents isapplied. Dye remaining in cracks or other surfacedefects is drawn to the surface of the developer bycapillary action and the resultant stains indicate theirlocations.

27. Fluorescent testing is based on the principlethat when ultra-violet radiation falls on a chemicalcompound, known as fluorescent ink, it is absorbedand its energy re-emitted as visible light. If a suitableink is allowed to penetrate surface cavities, theplaces where it is trapped will be revealed under therays of an ultra-violet lamp by brilliant lightemissions.

28. Magnetic crack testing (fig. 25-3) can only beapplied to components which can be magnetized.The part is first magnetized and then sprayed with, orimmersed in, a low viscosity fluid which containsferrous particles and is known as 'ink'. The two wallsof a crack in the magnetized part form magneticpoles and the magnetic field between these polesattracts the particles in the ink, so indicating the crack(fig. 25-4). In some instances, the ink may containfluorescent particles which enable their build-up to beviewed under an ultra-violet lamp, A part that hasbeen magnetically crack tested must be de-magnetized after inspection.

29. Chromic acid anodizing may be used as ameans of crack detection on aluminium parts, e.g.compressor blades. This process, in addition toproviding an oxide film that protects againstcorrosion, gives a surface that reveals even thesmallest flaws.



Fig. 25-3 Magnetic crack testing.

30. When the requirement for a detailed inspectionon a component such as a turbine disc is necessary,etching of the disc surfaces would be followed bybinocular inspection of the blade retention areas. Thewhole disc would then be subjected to magneticcrack test, followed by re-inspection of the discincluding a further binocular inspection of the bladeretention areas.

Repair31. To ensure that costs are maintained at thelowest possible level, a wide variety of techniquesare used to repair engine parts to make them suitablefor further service. Welding, the fitting of interferencesleeves or liners, machining and electro-plating aresome of the techniques employed during repair.

32. The welding techniques detailed in Part 22 areextensively used and range from welding of cracksby inert gas welding to the renewing of sections offlame tubes and jet pipes by electric resistancewelding.

33. On some materials now being used for gasturbine engine parts, different techniques may haveto be employed. An example of this is the highstrength titanium alloys which suffer from brittlewelds if they are allowed to become contaminated byoxygen during the cooling period. Parts made in

these alloys, which have to withstand high stressloadings in service, are often welded in a bag orplastic dome that is purged by an inert gas beforewelding commences.

34. More advanced materials and constructionsmay have to be welded by electron-beam welding.This method not only enables dissimilar metals to bewelded, but also complete sections of the moreadvanced fabricated constructions, e.g. a section ofa fabricated rotor drum, to be replaced at a lowpercentage cost of a new drum.

35. Some repair methods, such as welding, mayaffect the properties of the materials and, to restorethe materials to a satisfactory condition, it may benecessary to heat treat the parts to remove thestresses, reduce the hardness of the weld area orrestore the strength of the material in the heataffected area, Heat treatment techniques are alsoused for removing distortions after welding. The partsare heated to a temperature sufficient to remove thestresses and, during the heat treatment process,fixtures are often used to ensure the parts maintaintheir correct configuration.

36. Electro-plating methods are also widely used forrepair purposes and these range from chromiumplating, which can be used to provide a very hardsurface, to thin coatings of copper or silver plating,which can be applied to such areas as bearinglocations on a shaft to restore a fitting diameter thatis only slightly worn.

37. Many repairs are effected by machiningdiameters and/or faces to undersize dimensions orbores to oversize dimensions and then fitting shims,liners or metal spraying coatings of wear resistantmaterial. The effected surfaces are then restored totheir original dimensions by machining or grinding.

38. The inspection of parts after they have beenrepaired consists mainly of a penetrant or magneticinspection. However, further inspection may berequired on parts that have been extensivelyrepaired and this may involve pressure testing or X-ray inspection of welded areas.

39. Re-balancing of the main rotating assembly willbe necessary during overhaul, even though all theoriginal parts may be refitted, and this is done asdescribed in para 40.



Fig. 25-4 Cracks revealed by magneticcrack detection.

Balancing40. Because of the high rotational speeds, anyunbalance in the main rotating assembly of a gasturbine engine is capable of producing vibration andstresses which increase as the square of therotational speed. Therefore very accurate balancingof the rotating assembly is necessary.

41. The two main methods of measuring andcorrecting unbalance are single plane (static)balancing and two plane (dynamic) balancing. Withsingle plane, the unbalance is only in one plane i.e.,centrally through the component at 90 degrees to theaxis. This is appropriate for components such asindividual compressor or turbine discs.

42. For compressor and/or turbine rotor assembliespossessing appreciable axial length, unbalance maybe present at many positions along the axis. Ingeneral it is not possible to correct this combinationof distributed unbalance in a single plane. However,if two correction planes are chosen, usually at axiallyopposed ends of the assembly, it is always possibleto find a combination of two unbalance weights whichare equivalent for the unbalances present in theassembled rotor, hence two plane balancing.

43. To illustrate this point refer to fig. 25-5, the dis-tribution of unbalance in the rotor has been reducedto an equivalent system of two unbalances 'A' and'B'. The rotor is already in static balance because inthis example 'A' and 'B' are equal and opposed.However, when the part is rotating, each weightproduces its own centrifugal force in opposition to theother causing unbalance couples, with the tendencyto turn the part end-over-end. This action is restrictedby the bearings, with resultant stresses and vibration.It will be seen, therefore, that to bring the part to astate of dynamic balance, an equal amount of weightmust be removed at 'A' and 'B' or added at 'P' and 'O'.When the couples set up by the centrifugal forces areequal, it is said that a part is dynamically balanced.Unbalance is expressed in units of ounce-inches,thus one ounce of excess weight displaced twoinches from the axis of a rotor is two ounce inches ofunbalance.

44. When balancing assemblies such as L.P.compressor rotors, the readings obtained are incon-sistent due to blade scatter. Blade scatter is causedby the platform and root or retaining pin clearancesallowing the blades to interlock at the platforms andassume a different radial position during each

balancing run. This only occurs at the relatively lowr.p.m. used for balancing, because, during enginerunning, the blades will assume a consistent radialposition as they are centrifuged outwards.

45. To obtain authentic balance results when bladescatter is present, it is necessary to record readingsfrom several balance runs, e.g. 8 runs, thereafterdetermining a vector mean.

46. A typical dynamic balancing machine forindicating the magnitude and angular position ofunbalance in each plane is shown in fig. 25-6.Correction of unbalance may be achieved by one ora combination of the following basic methods; redis-tribution of weight, addition of weight and removal ofweight.

47. Redistribution of weight is possible for suchassemblies as turbine and compressor discs, whenblades of different weight can be interchanged and,on some engines, clamped weights are provided forpositioning around the disc.

48. The addition of weight is probably the mostcommon method used, certain parts of the assemblyhaving provision for the fitting of screwed or rivetedplugs, heavy wire, balancing plates or nuts.



Fig. 25-5 Unbalance couples due tocentrifugal force.

49. Removing weight by machining metal frombalancing lands is the third basic method, butnormally it is only employed on initial manufacturewhen balancing a component, e.g. a turbine shaft ora compressor shaft, that is part of a larger assembly.

50. Modular assembled engines demand differentbalancing methods which include the use ofsimulated engine rotors. The dummy rotors mustreproduce the bearing span, weight, centre of gravityand dynamic characteristics of the sub-assembly it



Fig. 25-6 Dynamic balancing machine.

Fig. 25-7 Simulated engine rotor assemblies.

replaces and must be produced and maintained sothat their own contribution to the measuredunbalance is minimal. In order to obtain the correctdynamic reactions when balancing a compressorand/or turbine rotor assembly on its own, with theintention of making it an independent module, asimulated engine rotor must be used to replace themating assembly, ref. fig. 25-7. The compressorand/or turbine rotor assembly having then been inde-pendently balanced with the appropriate dummyrotor is thus corrected both for its own unbalance andinfluence due to geometric errors on any othermating assembly.

Moment weighing of blades51. With the introduction of the large fan blade,moment weighing of blades has assumed a greatersignificance, ref. fig. 25-8. This operation takes intoaccount the mass of each blade and also the positionof its centre of gravity relative to the centre line of thedisc into which the blade is assembled. Themechanical system of blade moment weighing maybe integrated with a computer, ref. fig, 25-9, whichwill automatically optimise the blade distribution. Themoment weight of a blade in units i.e. or,is identical to the unbalance effect of the blade wheninstalled into a disc. The recorded measurement ofblade moment weights enables each blade to be

distributed around the disc in order that theseunbalances are cancelled.

Assembling52. The engine can be built in the vertical orhorizontal position, using the ram or stand illustratedin fig. 25-TO and 25-11 respectively. Assembling ofthe engine sub-assemblies or modules is done inseparate areas, thus minimizing the build time on thebuild rams or stands.



Fig. 25-8 Principle of blade moment weighing.

Fig. 25-9 Integrated blade momentweighing.



Fig. 25-10 Engine assembly --- vertical.

Fig. 25-11 Engine assembly--- horizontal.

53. During assembling, inspection checks aremade. These checks can establish dimensions toenable axial and radial clearances to be calculatedand adjustments to be made, or they can ascertainthat vital fitting operations have been correctlyeffected. Dimensional checks are effected duringdisassembly to establish datums which must berepeated on subsequent re-assernbly.

54. To ensure that the nuts, bolts and setscrewsthroughout the engine and its accessories areuniformly tight, controlled torque tightening isapplied, fig. 25-12, the torque loading figure isdetermined by the thread diameter and the differingcoefficients of friction allied with thread finish i.e.,silver or cadmium plating and the lubricant used.

Testing55. On completion of assembly, every productionand/or overhauled engine must be tested in a 'sea-level' test cell (fig. 25-14), i.e. a test cell in which theengine is run at ambient temperature and pressureconditions, the resultant performance figures beingcorrected to International Standard Atmosphere(I.S.A.) sea-level conditions (Part 21).

56. To ensure that the engine performance meetsthat guaranteed to the customer and the require-ments of the Government licensing and purchasingauthorities, each engine is tested to an acceptancetest schedule.

57. In addition to the 'sea-level' tests, sampleengines are tested to a flight evaluation test



Fig. 25-12 Torque tightening.



Fig. 25-13 A high altitude test cell.

Fig. 25-14 A sea-level test cell.

schedule. These tests cover such characteristics asanti-icing, combustion and reheat efficiencies,performance, mechanical reliability, and oil and fuelconsumptions at the variety of conditions to whichthe engine may be subjected during its operationallife. Flight evaluation testing can be effected byinstalling the engine in an aircraft or in an altitude testcell (fig. 25-13) to test the variations of air humidity,pressure and temperature on the engine, itsaccessories and the oil and fuel systems. When in anaircraft, the engine is operated at the actualatmospheric conditions specified in the schedule, butin an altitude test cell, the engine is installed in anenclosed cell and tested to the schedule inconditions that are mechanically simulated.

58. Mechanical simulation comprises supplying theengine inlet with an accurately controlled massairflow at the required temperature and humidity, andadjusting the atmospheric pressure within theexhaust cell to coincide with pressure at varyingaltitudes.

59. The data which is accumulated from either 'sea-level' or altitude testing is retained for futuredevelopment, engine life assessment, material capa-bilities or any aspect of engine history.

60. During the testing of turbo-jet engines there is aneed to reduce the exhaust noise to withinacceptable limits. This may be achieved by severaldifferent means, each involving costly equipment.However, a typical silencer would do this byexpansion in the first section, damping by acoustictubes and final diffusion by a large exit through which

the hot gas would be directed upwards at a lowvelocity.

Preparing for storage/despatch61. The preparation of the engine/module forstorage and/or despatch is of major importance,since storage and transportation calls for specialtreatment to preserve the engine. To resist corrosionduring storage, the fuel system is inhibited by specialoil and all apertures are sealed off. The external andinternal surfaces of the engine are also protected byspecial inhibiting powders or by paper impregnatedwith inhibiting powder and the engine is enclosed ina re-usable bag (fig. 25-15) or plastic sheeting intowhich a specific amount of desiccant is inserted. Iftransportation by rail or sea is involved, the inhibitedand bagged engine may be packed in a woodencrate or metal case.



Fig. 25-15 Transportation stand and storagebag.


Appendix 1

Conversion factorsUNIT ABBREVIATIONS

in = inchft = footyd = yardoz = ouncelb = poundcwt = hundredweightBtu = British thermal unithp = horsepowerHg = mercury

s = secondmin = minuteh = hourf = forceW = wattkW = kilowatt (Wx1000)mm = millimetre (mx0.001)m = metre

km = kilometre (mx1000)g = gramkg = kilogramN = newtonPa = pascalkPa = kilopascalJ = JoulekJ = kilojoule (Jx1000)MJ = megajoule (Jx1 000 000)

CONVERSION FACTORS - Exact values are printed in bold type.

LENGTH 1 in = 25.4 mm1 ft = 0.3048 m1 mile = 1 .60934 km1 International nautical mile = 1.852 km

AREA 1 in2 = 645.16 mm2

1 ft2 = 92903.04 mm2

VOLUME 1 UK fluid ounce = 28413.1 mm3

1 US fluid ounce = 29573.5 mm3

1 Imperial pint = 568261.0 mm3

1 US liquid pint = 473176.0 mm3

1 UK gallon = 4546090.0 mm3

1 US gallon = 3785410.0 mm3

1 in3 = 16387.1 mm3

1 ft3 = 0.0283168 m3

MASS 1 oz (avoir.) = 28. 3495 g1 lb = 0.45359237 kg1 UK ton = 1.01605 tonne 1 short ton (2000 lb) = 0.907 tonne


DENSITY 1 lb/in3 = 27679.9 kg/m3

1 lb/ft3 = 16.0185 kg/m3

VELOCITY 1 in/min = 0.42333 mm/s 1 ft/min = 0.00508 m/s1 ft/s = 0.3048 m/s1 mile/h = 1.60934 km/h 1 International knot = 1.852 km/h

ACCELERATION 1 ft/s2 = 0.3048 m/s2

MASS FLOW RATE 1 lb/h = 1.25998x10-4 kg/s

FORCE 1 Ibf = 4.44822 N 1 kgf = 9.80665 N1 tonf = 9964.02 N

PRESSURE 1 in Hg (0.0338639 bar) = 3386.39 Pa1 Ibf/in2 (0.0689476 bar) = 6894.76 Pa1 bar = 100.0 kPa1 standard atmosphere = 101.325 kPa

MOMENT (torque) 1 Ibf in = 0.112985 Nm1 Ibf ft = 1.35582 Nm

ENERGY/ HEAT/ WORK 1 hph = 2.68452 MJ1 therm = 105.506 MJ1 Btu = 1.05506 kJ1 kWh = 3.6 MJ

HEAT FLOW RATE 1 Btu/h = 0.293071 W

POWER 1 hp (550 ft Ibf/s) = 0.745700 kW

KINEMATIC VISCOSITY 1 ft2/s = 929.03 stokes = 0.092903 m2/s

SPECIFIC ENTHALPY 1 Btu/ft3 = 37.2589 kJ/rn2

1 Btu/lb = 2.326 kJ/kg

PLANE ANGLE 1 radian (rad) = 57.2958 degrees1 degree = 0.0174533 rad = 1.1111 grade1 second = 4.84814x10-6 rad = 0.0003 grade1 minute = 2.90888x10-4 rad = 0.0185 grade

VELOCITY OF ROTATION 1 revolution/min = 0.104720 rad/s