_aviation__The_Jet_Engine_Gas_turbine__turbojet__turbofan_Rolls-Royce

Vertical/short take-off and landingpipe, the joint faces of which are so angled that, 10. The switch-in deflector consists of one or a pairwhen the sections are counter-rotated, the nozzle of heavily reinforced doors which form part of the jetmoves from the horizontal to the vertical position (fig. pipe wall when the engine is operating in the forward18-5). To avoid either a side component o! thrust or a thrust condition. To select lift thrust, the doors arethrust line offset from the engine axis during the moved to blank off the conventional propelling nozzlemovement of the nozzle it is necessary that the first and direct the exhaust flow into a lift nozzle (fig. 18-joint face is perpendicular to the axis of the jet pipe. 8). The lift nozzles may be designed so that they canIf it is desired that the nozzle does not rotate, as may be mechanically rotated to vary the angle of thebe the case if it is a variable area nozzle, a third joint thrust and permit intermediate lift/thrust positions toface which is perpendicular to the axis of the nozzle be selected.is required.9. The two and four nozzle deflector systems use 11. A second type of switch-in deflector system isside mounted nozzles (fig. 18-6) which can rotate on used on the tandem fan or hybrid fan vectored thrustsimple bearings through an angle of well over 90 engine (fig. 18-9). In this case the deflector system isdegrees so that reverse thrust can be provided if situated between the stages of the fan of a mixedrequired. A simple drive system, for example, a flow turbo-fan engine. In normal flight the valve issprocket and chain, can be used and by mechanical positioned so that the engine operates in the sameconnections all the nozzles can be made to deflect manner as a mixed flow turbo-fan and for lift thrustsimultaneously. For forward flight, to avoid a high the valve is switched so that the exhaust flow fromperformance loss and consequent increase in fuel the front part of the fan exhausts through downwardconsumption, careful design of the exhaust unit and facing lift nozzles and a secondary inlet is opened tonozzle aerodynamic passages are essential to provide the required airflow to the rear part of the fanminimize the pressure losses due to turning the and the main engine. On a purely subsonic V/STOLexhaust flow through two close coupled bends (fig. aircraft where fuel consumption is important the18-7). valve may be dispensed with and the engine operated permanently in the latter high by-passFig. 18-7 Nozzle duct configuration. mode described above. 12. Thrust deflecting nozzles will create an upstream pressure distortion which may excite vibration of the fan or low pressure turbine blades if the nozzle system is close to these components. Snubbers (Part 3) may be used on the fan blades to resist vibration. On the low pressure turbine, shrouds at the blade tips (Part 5) or wire lacing may be used to achieve the same result. Lift engines 13. The lift engine is designed to produce vertical thrust during the take-off and landing phases of V/STOL aircraft. Because the engine is not used in normal flight it must be light and have a small volume to avoid causing a large penalty on the aircraft. The lift engine may be a turbo-jet which for a given thrust gives the lowest weight and volume. Should a low jet velocity be necessary a lift fan may be employed. 14. Pure lift-jet engines have been developed with thrust/weight ratios of about 20:1 and still higher values are projected for the future. Weight is reduced by keeping the engine design simple and also by extensive use of composite materials (fig. 18-10). Because the engine is operated for only limited periods during specific flight conditions i.e. during take-off and landing, the fuel system can be simplified and a total loss oil system (Part 8), in which 191

Vertical/short take-off and landingFig. 18-8 Switch-In deflector system. exhaust nozzle may be replaced by a multi-lobe nozzle to increase the rate of mixing with thethe used lubricating oil is ejected overboard, can be surrounding air.used. 16. The lift-fan engine is designed to reduce the jet15. Lift engines can be designed to operate in the exhaust velocity, to reduce ground erosion and allowvertical or horizontal position and a thrust deflecting operation from unprepared ground surfaces. It alsonozzle fitted to provide some of the advantages of reduces the jet noise significantly. A range of designthrust vectoring. Alternatively, the engine may be options have been considered for this type of enginemounted so that it can swivel through a large angle and some are shown on fig. 18-11.to provide thrust vectoring. The lift-jet engine willhave an extremely hot, high velocity jet exhaust andto reduce ground erosion by the jet the normalFig. 18-9 Vectored thrust engine.192

Vertical/short take-off and landingFig. 18-10 A lift-jet engine.Remote lift systems 18. The remote lift-fan (fig. 18-12) is mounted in the17. Direct lift remote systems duct the by-pass air or aircraft wing or fuselage, and is driven mechanicallyengine exhaust air to downward facing lift nozzles or by air or gas ducted into a tip turbine, The driveremote from the engine. These nozzles may be in the system is provided by the main propulsion powerfront fuselage of the aircraft or in the wings. The plant or by a separate engine.engine duct is blocked by means of a diverter similarto that described in para. 10. 19. The advantage of the remote lift system is that it gives some freedom to the aircraft to position theFig. 18-11 Lift-fan engine configurations. 193

Vertical/short take-off and landingFig. 18-12 Remote lift fan. through at least 90 degrees to provide thrustpropulsion system to the best advantage whilst still vectoring (fig. 18-13). In addition to these propulsionmaintaining the resultant thrust near the aircraft engines, one or more lift engines may be installed tocentre of gravity in the jet lift mode. This freedom is provide supplementary lift during the take-off andachieved at a cost of increased volume, particularly landing phase of flight.with the gas driven systems, due to the size of theducts to feed the gas to the remote lift system. 21. The swivelling engine system can only be usedAlthough the mechanically driven remote lift-fan with two or more engines. This then introduces theeliminates the need for these large gas ducts, it is problem of safety in the event of an engine failure.done at the expense of long shafts and high power So, although there is only a small weight penalty andgearboxes and clutch systems. no increase in fuel consumption, safety considera-Swivelling engines tions tend to offset these advantages compared to20. This method consists of having propulsion some of the other powered lift systems. The normalengines which can be mechanically swiveled closed method of providing aircraft control at low speeds is by differential throttling and vectoring of the enginesFig. 18-13 Jet lift with swivelling nozzles. which simplifies the basic engine design but makes194 the control system more complex. Bleed air for STOL 22. Fig. 18-14 shows one method how STOL can be achieved with a form of 'flap blowing'. The turbo- fan engine has a geared variable pitch fan and an oversized low pressure (L. P.) compressor from the exit of which air is bled and ducted to the flap system in the wing trailing edge. The variable pitch fan enables high L.P. compressor speed and thus high bleed pressure to be maintained over a wide range of thrusts. This gives excellent control at greatly different aircraft flight conditions. LIFT THRUST AUGMENTATION 23. In many cases on V/STOL aircraft augmentation of the lift thrust is necessary to avoid an engine which is oversized for normal flight with the consequent effects of higher engine weight and fuel consumption than would be the case for a conventional aircraft- This lift thrust augmentation can be achieved in a number 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 ratings 24. Experience has shown that an engine rating structure can be devised which provides high thrust levels for short periods of time without reducing engine life. Operation in ground effect and the take- off and landing manoeuvres require maximum thrust for less than 15 seconds so that use of a short lift rating for that time is feasible. Fig. 18-15 shows an example of thrust permissible with a 15 second short lift rating compared to that with a 2.5 minute normal lift rating.

Vertical/short take-off and landingFig. 18-14 Flap blowing engine. Fig. 18-15 Thrust increases with short lift ratings.25. At high ambient temperatures, the engine mayrun into a turbine temperature limit before reaching Ejectorsits maximum r.p.m. and suffer a thrust loss as a 28. The principle of the ejector is that a small, highresult. Restoration of the thrust can be achieved by energy jet entrains large quantities of ambient air bymeans of water injection into the combustion viscous mixing and an increase in thrust over that ofchamber (Part 17) which allows operation at a higher the high energy jet results. A number of projectedturbine gas temperature for a given turbine blade V/STOL aircraft have incorporated this concept usingtemperature. If desired, water injection can also be either all the engine exhaust air or just the bypassused to increase the thrust at low ambient tempera- flow.tures. 195Lift 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.

Vertical/short take-off and landingFig. 18-16 Plenum chamber burning.Fig. 18-17 Remote augmented lift system.196

Vertical/short take-off and landingFig. 18-18 Reaction control system. Differential engine throttling 31. This method of control is used on multi-enginedAIRCRAFT CONTROL aircraft with the engines positioned in a suitable con- figuration. A rapid response rate is essential to29. The low forward speeds of V/STOL aircraft enable the engines to be used for aircraft stabilityduring take-off and transition do not permit the and control. It is usually necessary to combine differ-generation of adequate aerodynamic forces from the ential throttling with differential thrust vectoring tonormal flight control surfaces, it is therefore give aircraft control in all areas.necessary to provide one or more of the followingadditonal methods of controlling pitch, roll and yaw. Automatic control systems 32. Although it is possible for the pilot to control aReaction controls V/STOL aircraft manually, some form of automation30. This system bleeds air from the engine and can be of benefit and in particular will reduce the pilotducts it through nozzles at the four extremities of the workload. The pilot's control column is electronicallyaircraft (fig. 18-18), The air supply to the nozzles is connected to a computer or stabilizer that receivesautomatically cut off when the main engine swivelling signals from the control column, compares them withpropulsion nozzles are turned for normal flight or signals from the sensors that measure the attitude ofwhen the lift engines are shut down. The thrust of the the aircraft, and automatically adjusts the reactioncontrol nozzles is varied by changing their area controls, differential throttling or thrust vectoringwhich varies the amount of airflow passed. controls to maintain stability. 197

Rolls-Royce Turbomeca Adour MK151Napier Gazelle The Gazelle turbo-shaft engine first ran in December 1955 at 1260 shp, a figure later increased to 1610 shp on production engines. Gazelles were used to power Bristol Belvedere and Westland Wessex helicopters. Gazelle production was taken over by Rolls- Royce in 1961.

19: Noise suppressionContents PageIntroduction 199Engine noise 199Methods of suppressing noise202Construction and materials 205INTRODUCTION 3. Airframe self-generated noise is a factor in an aircraft's overall noise signature, but the principal1. Airport regulations and aircraft noise certification noise source is the engine.requirements, all of which govern the maximumnoise level aircraft are permitted to produce, have ENGINE NOISEmade jet engine noise suppression one of the mostimportant fields of research. 4. To understand the problem of engine noise suppression, it is necessary to have a working2. The unit that is commonly used to express noise knowledge of the noise sources and their relativeannoyance is the Effective Perceived Noise deciBel importance. The significant sources originate in the(EPNdB). It takes into account the pitch as well as fan or compressor, the turbine and the exhaust jet orthe sound pressure (deciBel) and makes allowance jets. These noise sources obey different laws andfor the duration of an aircraft flyover. Fig. 19-1 mechanisms of generation, but all increase, to acompares the noise levels of various jet engine varying degree, with greater relative airflow velocity.types. Exhaust jet noise varies by a larger factor than the compressor or turbine noise, therefore a reduction in exhaust jet velocity has a stronger influence than an equivalent reduction in compressor and turbine blade speeds. 199

Noise suppressionFig. 19-1 Comparative noise levels of various engine types.Fig. 19-2 Exhaust mixing and shock structure.200

Noise suppression5. Jet exhaust noise is caused by the violent and 6. Compressor and turbine noise results from thehence extremely turbulent mixing of the exhaust interaction of pressure fields and turbulent wakesgases with the atmosphere and is influenced by the from rotating blades and stationary vanes, and canshearing action caused by the relative speed be defined as two distinct types of noise; discretebetween the exhaust jet and the atmosphere. The tone (single frequency) and broadband (a wide rangesmall eddies created near the exhaust duct cause of frequencies). Discrete tones are produced by thehigh frequency noise but downstream of the exhaust regular passage of blade wakes over the stagesjet the larger eddies create low frequency noise. downstream causing a series of tones andAdditionally, when the exhaust jet velocity exceeds harmonics from each stage. The wake intensity isthe local speed of sound, a regular shock pattern is largely dependent upon the distance between theformed within the exhaust jet core. This produces a rows of blades and vanes. If the distance is shortdiscrete (single frequency) tone and selective ampli- then there is an intense pressure field interactionfication of the mixing noise, as shown in fig. 19-2. A which results in a strong tone being generated. Withreduction in noise level occurs if the mixing rate is the high bypass engine, the low pressureaccelerated or if the velocity of the exhaust jet compressor (fan) blade wakes passing overrelative to the atmosphere is reduced. This can be downstream vanes produce such tones, but of aachieved by changing the pattern of the exhaust jet lower intensity due to lower velocities and largeras shown in fig. 19-3. blade/vane separations. Broadband noise isFig. 19-3 Change of exhaust jet pattern to reduce noise level. 201

Noise suppressionproduced by the reaction of each blade to the 9. Listed amongst the several other sources ofpassage of air over its surface, even with a smooth noise within the engine is the combustion chamber. Itairstream. Turbulence in the airstream passing over is a significant but not a predominant source, due inthe blades increases the intensity of the broadband part to the fact that it is 'buried' in the core of thenoise and can also induce tones. engine. Nevertheless it contributes to the broadband noise, as a result of the violent activities which occur7. With the pure jet engine the exhaust jet noise is within the combustion chamber.of such a high level that the turbine and compressornoise is insignificant at all operating conditions, METHODS OF SUPPRESSING NOISEexcept low landing-approach thrusts. With the by-pass principle, the exhaust jet noise drops as the 10. Noise suppression of internal sources isvelocity of the exhaust is reduced but the low approached in two ways; by basic design to minimizepressure compressor and turbine noise increases noise originating within or propagating from thedue to the greater internal power handling. engine, and by the use of acoustically absorbent linings. Noise can be minimized by reducing airflow8. The introduction of a single stage low pressure disruption which causes turbulence. This is achievedcompressor (fan) significantly reduces the by using minimal rotational and airflow velocities andcompressor noise because the overall turbulence reducing the wake intensity by appropriate spacingand interaction levels are diminished. When the by- between the blades and vanes. The ratio betweenpass ratio is in excess of approximately 5 to 1, the jet the number of rotating blades and stationary vanesexhaust noise has reduced to such a level that the can also be advantageously employed to containincreased internal noise source is predominant. A noise within the engine.comparison between low and high by-pass enginenoise sources is shown in fig. 19-4.Fig. 19-4 Comparative noise sources of low and high by-pass engines.202

Noise suppression11. As previously described, the major source of Fig. 19-5 Types of noise suppressor.noise on the pure jet engine and low by-pass engineis the exhaust jet, and this can be reduced by a good understanding of the mechanisms of noiseinducing a rapid or shorter mixing region. This generation and comprehensive noise design rulesreduces the low frequency noise but may increase exist. As previously indicated, these are founded onthe high frequency level. Fortunately, high the need to minimize turbulence levels in the airflow,frequencies are quickly absorbed in the atmosphere reduce the strength of interactions between rotatingand some of the noise which does propagate to the blades and stationary vanes, and the optimum use oflistener is beyond the audible range, thus giving the acoustically absorbent linings.perception of a quieter engine. This is achieved byincreasing the contact area of the atmosphere with 203the 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

Noise suppressionFig. 19-6 Noise absorbing materials and location.204

Noise suppression16. Noise absorbing 'lining' material converts be accurately calibrated. Guide vanes are fitted toacoustic energy into heat. The absorbent linings (fig. the lobe-type suppressor to prevent excessive losses19-6) normally consist of a porous skin supported by by guiding the exhaust gas smoothly through thea honeycomb backing, to provide the required lobes to atmosphere. The suppressor is a fabricatedseparation between the facesheet and the solid welded structure and is manufactured from heat-engine duct. The acoustic properties of the skin and resistant alloys.the liner depth are carefully matched to the characterof the noise, for optimum suppression. The disad- 18. Various noise absorbing lining materials arevantage of liners is the slight increase in weight and used on jet engines. They fall mainly within twoskin friction and hence a slight increase in fuel categories, lightweight composite materials that areconsumption. They do however, provide a very used in the lower temperature regions and fibrous-powerful suppression technique. metallic materials that are used in the higher temperature regions. The noise absorbing materialCONSTRUCTION AND MATERIALS consists of a perforate metal or composite facing skin, supported by a honeycomb structure on a solid17. The corrugated or lobe-type noise suppressor backing skin which is bonded to the parent metal offorms the exhaust propelling nozzle and is usually a the duct or casing. For details of manufacture ofseparate assembly bolted to the jet pipe. Provision is these materials refer to Part 22.usually made to adjust the nozzle area so that it can 205

Rolls-Royce ConwayRolls-Royce RM60 Produced in response to an Admiralty contract for a coastal-craft engine with good cruising economy, the RM60, although based on aeroengine philosophy, was designed from the first as a marine gas turbine. Two RM60s went to sea in 1953 in the former steam gunboat HMS Grey Goose, the world's first warship to be powered solely by gas turbines.

20: Thrust distributionContents PageIntroduction 207Distribution of the thrust 207forces 209Method of calculating the 209thrust forcesCalculating the thrust of 212the engine Compressor casing Diffuser duct Combustion chambers Turbine assembly Exhaust unit and jet pipe Propelling nozzle Engine Inclined combustion chambersAfterburningINTRODUCTION the sum of the forward forces exceeds the sum of the rearward forces is normally known as the rated thrust1. Although the principles of jet propulsion (see Part of the engine.1) will be familiar to the reader, the distribution of thethrust forces within the engine may appear DISTRIBUTION OF THE THRUST FORCESsomewhat obscure- These forces are in effect gasloads resulting from the pressure and momentum 2. The diagram in fig. 20-1 is of a typical single-changes of the gas stream reacting on the engine spool axial flow turbo-jet engine and illustrates wherestructure and on the rotating components. They are the main forward and rearward forces act. The originin some locations forward propelling forces and in of these forces is explained by following the engineothers opposing or rearward forces. The amount that working cycle shown in Part 2. 207

Thrust distributionFig. 20-1 Thrust distribution of a typical single-spool axial flow engine.3. At the start of the cycle, air is induced into the which may be seen on the diagram. As the gas flowengine and is compressed. The rearward accelera- passes through the exhaust system (Part 6), smalltions through the compressor stages and the forward forces may act on the inner cone or bullet,resultant pressure rise produces a large reactive but generally only rearward forces are produced andforce in a forward direction. On the next stage of its these are due to the 'drag' of the gas flow at thejourney the air passes through the diffuser where it propelling nozzle.exerts a small reactive force, also in a forwarddirection, 6. It will be seen that during the passage of the air through the engine, changes in its velocity and4. From the diffuser the air passes into the pressure occur (Part 2). For instance, where acombustion chambers (Part 4) where it is heated, conversion from velocity (kinetic) energy to pressureand in the consequent expansion and acceleration of energy is required the passages are divergent inthe gas large forward forces are exerted on the shape, similar to that used in the compressorchamber walls. diffuser. Conversely, where it is required to convert the energy stored in the combustion gases to5. When the expanding gases leave the combustion velocity, a convergent passage or nozzle, similar tochambers and flow through the nozzle guide vanes that used in the turbine, is employed. Where thethey are accelerated and deflected on to the blades conversion is to velocity energy, 'drag' loads orof the turbine (Part 5). Due to the acceleration and rearward forces are produced; where the conversiondeflection, together with the subsequent straighten- is to pressure energy, forward forces are produced.ing of the gas flow as it enters the jet pipe, consider- Part 2, fig. 2-3 illustrates velocity and pressureable 'drag' results; thus the vanes and blades are changes at two points on the engine.subjected to large rearward forces, the magnitude of208

Thrust distributionMETHOD OF CALCULATING THE THRUST the compressor and the conditions at the outlet fromFORCES the compressor. Since the pressure and the velocity at the inlet to the compressor are zero, it is only7. The thrust forces or gas loads can be calculated necessary to consider the force at the outlet from thefor the engine, or for any flow section of the engine, compressor. Therefore, given that the compressor-provided that the areas, pressures, velocities andmass flow are known for both the inlet and outlet of OUTLET Area (A) = 182 sq.in.the particular flow section. Pressure (P) = 94 lb. per sq.in.8. The distribution of thrust forces shown in fig. 20-1 can be calculated by considering each component (gauge)in turn and applying some simple calculations. Thethrust produced by the engine is mainly the product Velocity (vJ) = 406 ft. per sec.of the mass of air passing through the engine and the Mass flow (W) = 153 lb. per sec.velocity increase imparted to it (i.e. Newtons SecondLaw of Motion), however, the pressure difference The thrustbetween the inlet to and the outlet from the particularflow section will have an effect on the overall thrust = (A x P) + MVJ − 0of the engine and must be included in the calculation. g9. To calculate the resultant thrust for a particular = (182 x 94) + 153 x 406 − 0flow section it is necessary to calculate the total 32thrust at both inlet and outlet, the resultant thrustbeing the difference between the two values = 19,049 lb. of thrust in a forward direction.obtained.10. Calculation of the thrust is achieved using thefollowing formula: Thrust = (A x P) + MVJ gWhere A = Area of flow section in sq.in. P = Pressure in lb. per sq.in. W = Mass flow in lb. per sec. Diffuser duct vJ = Velocity of flow in feet per sec. g = Gravitational constant 32.2 ft. per 13. The conditions at the diffuser duct inlet are the sec. per sec. same as the conditions at the compressor outlet, i.e. 19,049 lb. Therefore, given that the diffuser-- OUTLET Area (A) = 205 sq.in. Pressure (P) = 95 lb. per sq.in.CALCULATING THE THRUST OF THE ENGINE (gauge)11. When applying the above method to calculate Velocity (vJ) = 368 ft. per sec.the individual thrust loads on the various components Mass flow (W) = 153 lb. per sec.it is assumed that the engine is static. The effect ofaircraft forward speed on the engine thrust will be The thrustdealt with in Part 21. In the following calculations 'g'is taken to be 32 for convenience. To assist in these = (A x P) + W VJ − 19,049calculations the locations concerned are illustrated gby a number of small diagrams. = (205 x 95) + 153 x 368 − 19,049Compressor casing 3212. To obtain the thrust on the compressor casing itis necessary to calculate the conditions at the inlet to = 21,235 - 19,049 = 2,186 lb. of thrust in a forward direction. 209

Thrust distribution Therefore given that the turbine-- OUTLET Area (A) = 480 sq.in. Pressure (P) = 21 lb. per sq.in. (gauge) Velocity (vJ) = 888 ft. per sec. Mass flow (W) = 153 lb. per sec. The thrust = (A x P) + WVJ − 55,417 g = (480 x 21) + 153 x 888 − 55,417 32Combustion chambers = 14,326 - 55,41714. The conditions at the combustion chamber inlet = -41,091are the same as the conditions at the diffuser outlet, This negative value means a force acting in a rearward direction.i.e. 21,235 lb. Therefore, given that the combustionchamber-OUTLET Area (A) = 580 sq.in. Pressure (P) = 93 lb. per sq.in. (gauge) Velocity (vJ) = 309 ft. per sec. Mass flow (W) = 153 lb. per sec.The thrust= (A x P) + W VJ − 21,235 g= (580 x 93) + 153 x 309 − 21,235 32= 55,417 - 21,235= 34,182 !b. of thrust in a forward direction. Exhaust unit and jet pipe 16. The conditions at the inlet to the exhaust unit are the same as the conditions at the turbine outlet, i.e. 14,326 lb. Therefore, given that the jet pipe--Turbine assembly OUTLET Area (A) = 651 sq.in.15. The conditions at the turbine inlet are the sameas the conditions at the combustion chamber outlet, Pressure (P) = 21 lb. per sq.in.i.e. 55,417 lb. (gauge)210 Velocity (vJ) = 643 ft. per sec. Mass flow (W) = 153 lb. per sec.

Thrust distributionThe thrust The thrust= (A x P) + W VJ − 14,326 = (A x P) + W VJ − 16,745 g g= (651 x 21) + 153 x 643 − 14,326 = (332 x 6) + 153 x 1,917 − 16,745 32 32= 16,745 - 14,326 = 11,158 - 16,745= 2,419 lb. of thrust in a forward direction. = 5,587lb. acting in a rearward direction. It is emphasized that these are basic calculations and such factors as the effect of air offtakes have been ignored. 18. Based on the individual calculations, the sum of the forward or positive loads is 57,836 lb. and the sum of the rearward or negative loads is 46,678 lb. Thus, the resultant (gross or total) thrust is 11,158 lb.Propelling nozzle Engine 19. It will be of interest to calculate the thrust of the engine by considering the engine as a whole, as the resultant thrust should be equal to the sum of the individual gas loads previously calculated.17. The conditions at the inlet to the propelling 20. Although the momentum change of the gasnozzle are the same as the conditions at the jet pipe stream produces most of the thrust developed by theoutlet, i.e. 16,745 lb. engine (momentum thrust =WVJ gTherefore, given that the propelling nozzle-- ), an additionalOUTLET Area (A) = 332 sq.in. thrust is produced when the engine operates with the propelling nozzle in a 'choked' condition (Part 6). This Pressure (P) = 6 lb. per sq.in. (gauge) thrust results from the aerodynamic forces which are Velocity (vJ) = 1,917 ft. per sec. created by the gas stream and exert a pressure Mass flow (W) = 153 lb. per sec. 211

Thrust distributionacross the exit area of the propelling nozzle by combining the gas loads on the individual(pressure thrust). Algebraically, this force is engine locations.expressed as (P-P0) A. 21. On engines that operate with a non-chokedWhere A = Area of propelling nozzle in sq.in. nozzle, the (P-P0) A function does not apply and the P = Pressure in lb. per sq.in. thrust results only from the gas stream momentum P0 = Atmospheric pressure in lb. per sq.in. change.Therefore, assuming values of mass flow, pressure Inclined combustion chambersand area to be the same as in the previous calcula- 22. In the previous example (Para. 14) the flowtions i.e. through the combustion chamber is axial, however, if the combustion chamber is inclined towards the axisArea of propelling nozzle (A) = 332 sq.in. of the engine, then the axial thrust will be less than for an axial flow chamber. This thrust can be obtainedPressure (P) = 6 lb. per sq.in. by multiplying the sum of the outlet thrust by the cosine of the angle (see fig. 20-2). The (gauge) BaseAtmospheric Pressure (P) = 0 lb. per sq.in. cosine = Hypotenuse and for a given angle (gauge) is obtained by consulting a table of cosines. It shouldMass flow (W) = 153 lb. per sec. be emphasized that if the inlet and outlet are at different angles to the engine axis, it is necessary toVelocity (vJ) = 1,917 ft. per sec. multiply the inlet and outlet thrusts separately by the cosine of their respective angles.The thrust AFTERBURNING= (P − P0) ⋅ A + W VJ − 0 g 23. When the engine is fitted with an afterburner (Part 16), the gases passing through the exhaust= (6 − 0) ⋅ 332 + 153 X 1,917 − 0 32= 1,992 + 9,166= 11,158 lb., the same as previously calculatedFig. 20-2 A hypothetical combustion chamber showing values required for calculating thrust.212

Thrust distribution 24. Assuming that an afterburner jet pipe and propelling nozzle are fitted to the engine used in the previous calculations, and the new conditions at the propelling nozzle are as follows- OUTLET Area (A) = 455 sq.in. Pressure (P) = 5 lb. per sq.in. (gauge) Velocity (vJ) = 2,404 ft. per sec. Mass flow (W) = 157 lb. per sec. The thrust = (A x P) + W VJ − 16,745 g = (455 x 5) + 157 x 2,404 − 16,745 32 = 14,069 - 16,745 = 2,676 lb. acting in a rearward direction.system are reheated to provide additional thrust. The Therefore, compared with the previous calculation ineffect of afterburning is to increase the volume of the para. 17, it will be seen that the negative thrust isexhaust gases, thus producing a higher exit velocity reduced from -5,587 lb. to -2,676 lb.; the overallat the propelling nozzle. positive thrust is thus increased by 2,911 lb; which is equivalent to a thrust increase of more than 25 per cent. 25. To arrive at the total thrust of the engine with afterburning the calculations in para. 20 should use the above figures. 213

Rolls-Royce RB168 MK807Blackburn Nimbus The Nimbus was developed from the A129 turbo-shaft which, in its turn, was a modified Turbomeca Artouste built under licence. The Nimbus developed 968 hp, but for helicopter use was flat-rated at 710 hp. The engine was used in Westland Wasp and Scout helicopters and four 700 hp units were used to power the experimental 5RN-2 hovercraft.

21: PerformanceContents PageIntroduction 215Engine thrust on the test 217bench 218 Comparison between thrust and 223 horse-power 225Engine thrust in flight Effect of forward speed Effect of afterburning on engine thrust Effect of altitude Effect of temperaturePropulsive efficiencyFuel consumption andpower-to-weight relationshipINTRODUCTION 2. Since the thrust or s.h.p. developed is dependent on the mass of air entering the engine and the accel-1. The performance requirements of an engine are eration imparted to it during the engine cycle, it isobviously dictated to a large extent by the type of obviously influenced, as subsequently described, byoperation for which the engine is designed. The such variables as the forward speed of the aircraft,power of the turbo-jet engine is measured in thrust, altitude and climatic conditions, These variablesproduced at the propelling nozzle or nozzles, and influence the efficiency of the air intake, thethat of the turbo-propeller engine is measured in compressor, the turbine and the jet pipe; conse-shaft horse-power (s.h.p.) produced at the propeller quently, the gas energy available for the productionshaft. However, both types are in the main assessed of thrust or s.h.p. also varies.on the amount of thrust or s.h.p. they develop for agiven weight, fuel consumption and frontal area. 3. In the interest of fuel economy and aircraft range, the ratio of fuel consumption to thrust or s.h.p. should be as low as possible. This ratio, known as the specific fuel consumption (s.f.c.), is expressed in pounds of fuel per hour per pound of net thrust or s.h.p. and is determined by the thermal and propulsive efficiency of the engine. In recent years considerable progress has been made in reducing s.f.c. and weight. These factors are further explained in para. 46. 215

Performance4. Whereas the thermal efficiency is often referred instance, using the symbols shown in fig. 21-1 theto as the internal efficiency of the engine, thepropulsive efficiency is referred to as the external overall compressor pressure ratio is P3 . Theseefficiency. This latter efficiency, described in para. 37, P1explains why the pure jet engine is less efficient thanthe turbo-propeller engine at lower aircraft speeds symbols vary slightly for different types of engine; forleading to development of the by-pass principle and,more recently, the propfan designs. instance, with high by-pass ratio engines, and also5. The thermal and the propulsive efficiency also when afterburning (Part 16) is incorporated,influence, to a large extent, the size of thecompressor and turbine, thus determining the weight additional symbols are used.and diameter of the engine for a given output. 7. To enable the performance of similar engines to6. These and other factors are presented in curves be compared, it is necessary to standardize in someand graphs, calculated from the basic gas laws (Part conventional form the variations of air temperature2), and are proved in practice by bench and flight and pressure that occur with altitude and climatictesting, or by simulating flight conditions in a high conditions. There are in use several differentaltitude test cell. To make these calculations, specific definitions of standard atmospheres, the one in mostsymbols are used to denote the pressures and tem- common use being the International Standardperatures at various locations through the engine; for Atmosphere (I.S.A.). This is based on a temperature lapse rate of approximately 1.98 K. degrees per 1,000ft,, resulting in a fall from 288.15 deg.K. (15 deg.C) at sea level to 216.65 deg.K (-56.5 deg.C.) at 36,089 ft. (the tropopause). Above this altitude theFig. 21-1 Temperature and pressure notation of a typical turbo-jet engine.216

Performancetemperature is constant up to 65.617ft. The I.S.A. not in pounds of thrust the factors are different. Forstandard pressure at sea level is 14.69 pounds per example, the correction for s.h.p. is:square inch falling to 3.28 pounds per square inch at S.h.p. (corrected) =the tropopause (refer to I.S.A. table fig. 21-10). s.h.p. (observed) x 30 x 273 + 15ENGINE THRUST ON THE TEST BENCH P0 273 + T08. The thrust of the turbo-jet engine on the test where P0 = atmospheric pressure (in.Hg.)bench differs somewhat from that during flight. (observed)Modern test facilities are available to simulateatmospheric conditions at high altitudes thus T0 = atmospheric temperature in deg.C.providing a means of assessing some of the (observed)performance capability of a turbo-jet engine in flightwithout the engine ever leaving the ground. This is 30 = I.S.A. standard sea level pressureimportant as the changes in ambient temperature (in.Hg.)and pressure encountered at high altitudes consider-ably influence the thrust of the engine. 273 + 15 = I.S.A. standard sea level temperature in deg.K.9. Considering the formula derived in Part 20 forengines operating under 'choked' nozzle conditions, 273 + T0 = Atmospheric temperature in deg.K. In practice there is always a certain amount of jet thrust in the total output of the turbo-propeller engine and this must be added to the s.h.p. The correction for jet thrust is the same as that in para. 10.Thrust = (P − P0) ⋅ A + W VJ 12. To distinguish between these two aspects of the g power output, it is usual to refer to them as s.h.p. and thrust horse-power (t.h.p.). The total equivalentit can be seen that the thrust can be further affected horse-power is denoted by t.e.h.p. (sometimesby a change in the mass flow rate of air through the e.h.p.) and is the s.h.p. plus the s.h.p. equivalent toengine and by a change in jet velocity. An increase in the net jet thrust. For estimation purposes it is takenmass airflow may be obtained by using water that, under sea- level static conditions, one s.h.p. isinjection (Part 17) and increases in jet velocity by equivalent to approximately 2.6 lb. of jet thrust.using afterburning (Part 16). Therefore :10. As previously mentioned, changes in ambient t.e.h.p. = s.h.p. + jet thrust lb. 2.6pressure and temperature considerably influence the 13. The ratio of jet thrust to shaft power is influenced by many factors. For instance, the higherthrust of the engine. This is because of the way they the aircraft operating speed the larger may be the required proportion of total output in the form of jetaffect the air density and hence the mass of air thrust. Alternatively, an extra turbine stage may be required if more than a certain proportion of the totalentering the engine for a given engine rotational power is to be provided at the shaft. In general, turbo-propeller aircraft provide one pound of thrustspeed. To enable the performance of similar engines for every 3.5 h,p. to 5 h.p.to 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: Comparison between thrust and horse-power 14. Because the turbo-jet engine is rated in thrustThrust (lb.) (corrected) = and the turbo-propeller engine in s.h.p., no direct comparison between the two can be made without a thrust (lb.) (observed) x 30 power conversion factor. However, since the turbo- P0 propeller engine receives its thrust mainly from the propeller, a comparison can be made by convertingwhere P0 = atmospheric pressure in inches of the horse-power developed by the engine to thrust or mercury (in. Hg.) (observed) the thrust developed by the turbo-jet engine to t.h.p.; that is, by converting work to force or force to work.30 = I.S.A. standard sea level pressure For this purpose, it is necessary to take into account the speed of the aircraft.(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 217

Performance15. The t.h.p. is expressed as FV ENGINE THRUST IN FLIGHT 550 ft. per sec . 17. Since reference will be made to gross thrust,where F = lb. of thrust momentum drag and net thrust, it will be helpful to define these terms: V = aircraft speed (ft. per sec.) from Part 20, gross or total thrust is the product of the mass of air passing through the engine and the jetSince one horse-power is equal to 550 ft.lb. per sec. velocity at the propelling nozzle, expressed as: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 also WvJ gcommon to quote the speed in knots (nautical miles (P − P0)A +per 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. The momentum drag is the drag due to the16. Thus if a turbo-jet engine produces 5,000 lb. of momentum of the air passing into the engine relative WVnet thrust at an aircraft speed of 600 m.p.h. the t.h.p. to the aircraft velocity, expressed as g where 5,000 x 600would be 375 = 8,000 W = Mass flow in lb. per sec.However, if the same thrust was being produced by V = Velocity of aircraft in feet per sec.a turbo-propeller engine with a propeller efficiency of g = Gravitational constant 32.2 ft. per sec. per55 per cent at the same flight speed of 600 m.p.h., sec.then the t.h.p. would be The net thrust or resultant force acting on the aircraft 100 in flight is the difference between the gross thrust 55 8,000 x = 14,545 and the momentum drag.Thus at 600 m.p.h. one lb. of thrust is the equivalent 18. From the definitions and formulae stated inof about 3 t.h.p. para, 17; under flight conditions, the net thrust of theFig. 21-2 The balance of forces and expression for thrust and momentum drag.218

Performanceengine, simplifying, can be expressed as:(P − P0)A + W(vJ − V) gFig. 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 dragFig. 21-3 Thrust recovery with aircraft Fig. 21-4 The effect of aircraft speed on speed. thrust and fuel consumption. due to the forward speed so that the resultant loss of net thrust is partially recovered as the aircraft speed increases. A typical curve illustrating this point is shown in fig. 21-3. Obviously, the 'ram ratio' effect, or the return obtained in terms of pressure rise at entry to the compressor in exchange for the unavoidable intake drag, is of considerable importance to the turbo-jet engine, especially at high speeds. Above speeds of Mach 1.0, as a result of the formation of shock waves at the air intake, this rate of pressure rise will rapidly decrease unless a suitably designed air intake is provided (Part 23); an efficient air intake is necessary to obtain maximum benefit from the ram ratio effect. 21. As aircraft speeds increase into the supersonic region, the ram air temperature rises rapidly consistent with the basic gas laws (Part 2). This 219

Performancetemperature rise affects the compressor delivery air Fig. 21-5 The effect of aircraft speed ontemperature proportionately and, in consequence, to s.h.p. and fuel consumption.maintain the required thrust, the engine must besubjected to higher turbine entry temperatures. Since 26. Under flight conditions, however, this advantagethe maximum permissible turbine entry temperature is even greater, since the momentum drag is theis determined by the temperature limitations of the same with or without afterburning and, due to theturbine assembly, the choice of turbine materials and ram effect, better utilization is made of every poundthe 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 s.f.c.is 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.220

Performanceof air flowing through the engine. The following output to match the reduced mass airflow, soexample, using the static values given in Part 16, maintaining a constant engine speed.illustrates why afterburning thrust improves underflight conditions. The fall in air temperature increases the density of the air, so that the mass of air entering the27. Assuming an aircraft speed of 600 m.p.h. (880ft. compressor for a given engine speed is greater.per sec.), then Momentum drag is: This causes the mass airflow to reduce at a lower rate and so compensates to some extent 880 = 27.5 (approximat ely ) for the loss of thrust due to the fall in atmospheric 32 pressure. At altitudes above 36,089 feet and up to 65,617 feet, however, the temperatureThis means that every pound of air per second remains constant, and the thrust or s.h.p. is affected by pressure only.flowing through the engine and accelerated up to the Graphs showing the typical effect of altitude onspeed of the aircraft causes a drag of about 27.5 lb. thrust, s.h.p, and fuel consumption are illustrated in fig. 21-6 and fig. 21-7.28. Suppose each pound of air passed through theengine gives a gross thrust of 77.5 lb. Then the net Effect of temperaturethrust given by the engine per lb. of air per second is 33. On a cold day the density of the air increases so77.5 - 27.5 = 50 lb. that the mass of air entering the compressor for a given engine speed is greater, hence the thrust or29. When afterburning is selected, assuming the 30 s.h.p, is higher. The denser air does, however, increase the power required to drive the compressorper cent increase in static thrust given in para. 25, or compressors; thus the engine will require more fuel to maintain the same engine speed or will run atthe gross thrust will be 1.3 x 77.5 - 100.75 lb. Thus, a reduced engine speed if no increase in fuel is available.under flight condition of 600 m.p.h., the net thrust per 34. On a hot day the density of the air decreases,pound of air per second will be 100.75 - 27.5 = 73.25 thus reducing the mass of air entering the compressor and, consequently, the thrust of thelb. Therefore, the ratio of net thrust due to engine for a given r.p.m. Because less power will be 73.25 required to drive the compressor, the fuel controlafterburning is 50 = 1.465. In other words, a 30 system reduces the fuel flow to maintain a constant engine rotational speed or turbine entry temperature,per cent increase in thrust under static conditions as appropriate; however, because of the decrease in air density, the thrust will be lower. At a temperaturebecomes a 46.5 per cent increase in thrust at 600 of 45 deg.C., depending on the type of engine, a thrust loss of up to 20 per cent may be experienced.m.p.h. This means that some sort of thrust augmentation, such as water injection (Part 17), may be required.30. This larger increase in thrust is invaluable forobtaining higher speeds and higher altitude perform- 35. The fuel control system (Part 10) controls theances. The total and specific fuel consumptions are fuel flow so that the maximum fuel supply is heldhigh, but not unduly so for such an increase in practically constant at low air temperature conditions,performance. whereupon the engine speed falls but, because of the increased mass airflow as a result of the increase31. The limit to the obtainable thrust is determined in air density, the thrust remains the same. Forby the afterburning temperature and the remaining example, the combined acceleration and speedusable oxygen in the exhaust gas stream. Because control fuel system (Part 10) schedules fuel flow tono previous combustion heating takes place in the maintain a constant engine r.p.m., hence thrustduct of a by-pass engine, these engines with their increases as air temperature decreases until, at alarge residual oxygen surplus are particularly suited predetermined compressor delivery pressure, theto afterburning and static thrust increases of up to 70 fuel flow is automatically controlled to maintain aper cent are obtainable. At high forward speeds constant compressor delivery pressure and,several times this amount is achieved. 221Effect 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

PerformanceFig. 21-6 The effects of altitude on thrust Fig. 21-7 The effect of altitude on s.h.p. and and fuel consumption. fuel consumption.therefore, thrust. Fig. 21-8 illustrates this for a twin- 36. The pressure ratio control fuel system (Part 10)spool engine where the controlled engine r.p.m. is schedules fuel flow to maintain a constant enginehigh pressure compressor speed and the pressure ratio and, therefore, thrust below a prede-compressor 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.222

Performancetermined ambient air temperature. Above this energy wasted by the propelling mechanism. Wastetemperature the fuel flow is automatically controlledto prevent turbine entry temperature limitations from energy dissipated in the jet wake, which represents abeing exceeded, thus resulting in reduced thrust and,overall, similar curve characteristics to those shown loss, can be expressed as W(vJ − V)2 where (vJ-V)in fig. 21-8. In the instance of a triple-spool engine 2gthe pressure ratio is expressed as P4/P1. i.e. H.P. is the waste velocity. It is therefore apparent that atcompressor delivery pressure/engine inlet pressure. the aircraft lower speed range the pure jet streamPROPULSIVE EFFICIENCY wastes considerably more energy than a propeller37. Performance of the jet engine is not onlyconcerned with the thrust produced, but also with the system and consequently is less efficient over thisefficient conversion of the heat energy of the fuel intokinetic energy, as represented by the jet velocity, and range. However, this factor changes as aircraftthe best use of this velocity to propel the aircraftforward, i.e. the efficiency of the propulsive system. speed increases, because although the jet stream38. The efficiency of conversion of fuel energy to continues to issue at a high velocity from the enginekinetic energy is termed thermal or internal efficiencyand, like all heat engines, is controlled by the cycle its velocity relative to the surrounding atmosphere ispressure ratio and combustion temperature.Unfortunately, this temperature is limited by the reduced and, in consequence, the waste energy lossthermal and mechanical stresses that can betolerated by the turbine. The development of new is reduced.materials and techniques to minimize theselimitations is continually being pursued. 40. Briefly, propulsive efficiency may be expressed39. The efficiency of conversion of kinetic energy to as: Work done on the aircraftpropulsive work is termed the propulsive or external Energy imparted to engine airflowefficiency and this is affected by the amount of kinetic or simply Work done Work done + work wasted in exhaust Work done is the net thrust multiplied by the aircraft speed. Therefore, progressing from the net thrust equation given in para. 18, the following equation is arrived at: Propulsive efficiency = V (P - P0 )A + W(v J − V) g  V (P - P0 )A + W(v J − V) + W(v J − V)2  g  2g Fig. 21-8 The effect of air temperature on a typical twin-spool engine. 223

PerformanceIn the instance of an engine operating with a non- 42. Assuming an aircraft speed (V) of 375 m.p.h.choked nozzle (Part 20), the equation becomes: and a jet velocity (vj) of 1,230 rn.p.h., the efficiency of a turbo-jet is: WV(vJ − V) WV(v J − V) + 12 W(v J − V)2 2 × 375 = approx. 47 per cent Simplified to : 2V 375 + 1,230 V + vJ On the other hand, at an aircraft speed of 600 m.p.h. the efficiency is:41. This latter equation can also be used for thechoked nozzle condition by using vj to represent the 2 × 600 = approx. 66 per centjet velocity when fully expanded to atmospheric 600 + 1,230pressure, thereby dispensing with the nozzlepressure term (P-P0)A. Propeller efficiency at these values of V is approxi- mately 82 and 55'per cent, respectively, and fromFig. 21-9 Propulsive efficiencies and aircraft speed.224

Performancereference to fig. 21-9 it can be seen that for aircraft pass ratios in the order of 15:1, and reduced 'drag'designed to operate at sea level speeds below results due to the engine core being 'washed' by theapproximately 400 m.p.h. it is more effective to low velocity aircraft slipstream and not the relativelyabsorb the power developed in the jet engine by high velocity fan efflux.gearing it to a propeller instead of using it directly inthe form of a pure jet stream. The disadvantage of 45. The improved propulsive efficiency of thethe propeller at the higher aircraft speeds is its rapid bypass system bridges the efficiency gap betweenfall off in efficiency, due to shock waves created the turbo-propeller engine and the pure turbo-jetaround the propeller as the blade tip speed engine. A graph illustrating the various propulsiveapproaches Mach 1.0. Advanced propeller efficiencies with aircraft speed is shown in fig. 21-9.technology, however, has produced a multi-bladed,swept back design capable of turning with tip speeds FUEL CONSUMPTION AND POWER-TO-WEIGHTin excess of Mach 1.0 without loss of propeller RELATIONSHIPefficiency. By using this design of propeller in acontra-rotating configuration, thereby reducing swirl 46. Primary engine design considerations, particu-losses, a 'prop-fan' engine, with very good propulsive larly for commercial transport duty, are those of lowefficiency capable of operating efficiently at aircraft specific fuel consumption and weight. Considerablespeeds in excess of 500 m.p.h. at sea level, can be improvement has been achieved by use of the by-produced. pass principle, and by advanced mechanical and aerodynamic features, and the use of improved43. To obtain good propulsive efficiencies without materials. With the trend towards higher by-passthe use of a complex propeller system, the by-pass ratios, in the range of 15:1, the triple-spool andprinciple (Part 2) is used in various forms. With this contra-rotating rear fan engines allow the pressureprinciple, some part of the total output is provided by and by-pass ratios to be achieved with short rotors,a jet stream other than that which passes through the using fewer compressor stages, resulting in a lighterengine cycle and this is energized by a fan or a and more compact engine.varying number of LP. compressor stages. Thisbypass air is used to lower the mean jet temperature 47. S.f.c. is directly related to the thermal andand velocity either by exhausting through a separate propulsive efficiencies; that is, the overall efficiencypropelling nozzle, or by mixing with the turbine of the engine. Theoretically, high thermal efficiencystream to exhaust through a common nozzle. requires high pressures which in practice also means high turbine entry temperatures. In a pure turbo-jet44. The propulsive efficiency equation for a high by- engine this high temperature would result in a highpass ratio engine exhausting through separate jet velocity and consequently lower the propulsivenozzles is given below, where W1 and VJ1 relate to efficiency (para. 40). However, by using the by-passthe by-pass function and W2 and vJ2 to the engine principle, high thermal and propulsive efficienciesmain function. can be effectively combined by bypassing a proportion of the L.P. compressor or fan delivery airPropulsive efficiency = to lower the mean jet temperature and velocity as referred to in para. 43. With advanced technology W1V(v J1 −V)+W2V(v J2 −V) engines of high by-pass and overall pressure ratios,W1V(v J1 −V)+W2V(v J2 −V)+12W1V(v J1 −V)2 +12W2V(v J2 −V)2 a further pronounced improvement in s.f.c. is obtained.By calculation, substituting the following values,which will be typical of a high by-pass ratio engine of 48. The turbines of pure jet engines are heavytriple-spool configuration, it will be observed that a because they deal with the total airflow, whereas thepropulsive efficiency of approximately 85 per cent turbines of by-pass engines deal only with part of theresults. flow; thus the H.P. compressor, combustion chambers and turbines, can be scaled down. The V = 583 rn.p.h. increased power per lb. of air at the turbines, to take W1 = 492 lb. per sec. advantage of their full capacity, is obtained by the W2 = 100 lb. per sec. increase in pressure ratio and turbine entry VJ1 = 781 m.p.h. temperature. It is clear that the by-pass engine is VJ2 = 812 m.p.h. lighter, because not only has the diameter of the high pressure rotating assemblies been reduced but thePropulsive efficiency can be further improved by engine is shorter for a given power output. With a lowusing the rear mounted contra-rotating fan configura-tion of the by-pass principle. This gives very high by- 225

PerformanceFig. 21-10 International Standard Atmosphere.226

Performanceby-pass ratio engine, the weight reduction compared given duty. The use of higher strength light-weightwith a pure jet engine is in the order of 20 per cent materials is also a contributory factor.for the same air mass flow. 50. For a given mass flow less thrust is produced by49. With a high by-pass ratio engine of the triple- the by-pass engine due to the lower exit velocity.spool configuration, a further significant improvement Thus, to obtain the same thrust, the by-pass enginein specific weight is obtained- This is derived mainly must be scaled to pass a larger total mass airflowfrom advanced mechanical and aerodynamic design, than the pure turbo-jet engine. The weight of thewhich in addition to permitting a significant reduction engine, however, is still less because of the reducedin the total number of parts, enables rotating size of the H.P. section of the engine. Therefore, inassemblies to be more effectively matched and to addition to the reduced specific fuel consumption, anwork closer to optimum conditions, thus minimizing improvement in the power-to-weight ratio is obtained.the number of compressor and turbine stages for a 227

Rolls-Royce RB168 Mk202/Mk203Rolls-Royce RB39 Clyde Encouraged by results obtained from the Trent, Rolls-Royce decided to go ahead with an engine designed from the start as a turbo- prop. Named the Clyde it utilized the axial compressor from the Metrovick F2 as first stage and a scaled up supercharger impeller from a Merlin as second stage. First running in August 1945 at 2000 shp, later engines produced up to 4200 shp.

22: ManufactureContents PageIntroduction 229Manufacturing strategy 230Forging 231Casting 233Fabrication 234Welding 235 Tungsten inert gas (T.I.G.) welding 237 Electron beam welding (E.B.W.) 238Electro-chemicalmachining (E.C.M.) 240 240 Stem drilling Capillary drillingElectro-dischargemachining (E.D.M.)Composite materials andsandwich casingsInspectionINTRODUCTION 3. No manufacturing technique or process that In any way offers an advantage is ignored and most1. During the design stages of the aircraft gas available engineering methods and processes areturbine engine, close liaison is maintained between employed in the manufacture of these engines, Indesign, manufacturing, development and product some instances, the technique or process maysupport to ensure that the final design is a match appear by some standards to be elaborate, timebetween the engineering specification and the man- consuming and expensive, but is only adopted afterufacturing process capability. confirmation that it does produce maximized component lives comparable with rig test achieve-2. The functioning of this type of engine, with its ments.high power-to-weight ratio, demands the highestpossible performance from each component. 4. Engine components are produced from a varietyConsistent with this requirement, each component of high tensile steel and high temperature nickel andmust be manufactured at the lowest possible weight cobalt alloy forgings. A proportion of components areand cost and also provide mechanical integrity cast using the investment casting process. Whilstthrough a long service life. Consequently, the fabrications, which form an increasing content, aremethods used during manufacture are diverse and produced from materials such as stainless steel,are usually determined by the duties each titanium and nickel alloys using modern joiningcomponent has to fulfil. 229

ManufactureFig. 22-1 Arrangements of a triple-spool turbo-jet engine.techniques i.e., tungsten inert gas welding, some give resistance to corrosion whilst others canresistance welding, electron beam welding and high be used to release unwanted stress.temperature brazing in vacuum furnaces. 8. The main structure of an aero gas turbine engine5. The methods of machining engine components is formed by a number of circular casings, ref. fig. 22-include grinding, turning, drilling, boring and 1, which are assembled and secured together bybroaching whenever possible, with the more difficult flanged joints and couplings located with dowels andmaterials and configurations being machined by tenons. These engines use curvic and hurthelectro-discharge, electro-chemical, laser hole couplings to enable accurate concentricity of matingdrilling and chemical size reduction. assemblies which in turn assist an airline operator when maintenance is required.6. Structural components i.e., cold spoiler, locationrings and by-pass ducts, benefit by considerable MANUFACTURING STRATEGYweight saving when using composite materials. 9. Manufacturing is changing and will continue to7. In addition to the many manufacturing methods, change to meet the increasing demands ofchemical and thermal processes are used on part aeroengine components for fuel efficiency, cost andfinished and finished components. These include weight reductions and being able to process theheat treatment, electro-plating, chromate sealing, materials required to meet these demands.chemical treatments, anodizing to prevent corrosion,chemical and mechanical cleaning, wet and dry 10. With the advent of micro-processors andabrasive blasting, polishing, plasma spraying, elec- extending the use of the computer, full automation oftrolytic etching and polishing to reveal metallurgical components considered for in house manufacturedefects. Also a variety of barrelling techniques for are implemented in line with supply groups manufac-removal o! burrs and surface improvement. Most turing strategy, all other components beingprocesses are concerned with surface changes, resourced within the world-wide supplier network.230

Manufacture11. This automation is already applied in the sequence, estimating and scheduling. Computermanufacture of cast turbine blades with the seven simulation allows potential cell and flow linecell and computer numerical controlled (C.N.C.) manufacture to be proven before physical machinegrinding centres, laser hardfacing and film cooling purchase and operation, thus preventing equipmenthole drilling by electro-discharge machining (E.D.M.). not fulfilling their intended purpose.Families of turbine and compressor discs areproduced in flexible manufacturing cells, employing 14. Each casing is manufactured from the lightestautomated guided vehicles delivering palletized material commensurate with the stress and tempera-components from computerized storage to C.N.C. tures to which it is subjected in service. For example,machining cells that all use batch of one techniques. magnesium alloy, composites and materials ofThe smaller blades, with very thin airfoil sections, are sandwich construction are used for air intakeproduced by integrated broaching and 360 degree casings, fan casings and low pressure compressorelectrochemical machining (E.C.M.) while inspection casings, since these are the coolest parts of theand processing are being automated using the engine. Alloy si eels are used for the turbine andcomputer. nozzle casings where the temperatures are high and because these casings usually incorporate the12. Tolerances between design and manufacturing engine rear mounting features. For casingsare much closer when the design specification is subjected to intermediate temperatures i.e. by-passmatched by the manufacturing proven capability. duct and combustion outer casings, aluminium alloys and titanium alloys are used.13. Computer Aided Design (C.A.D.) and ComputerAided Manufacture (CAM.) provides an equivalent FORGINGlink when engine components designed by C.A.D.can be used for the preparation of manufacturing 15. The engine drive shafts, compressor discs,drawings, programmes for numerically controlled turbine discs and gear trains are forged to as nearmachines, tool layouts, tool designs, operation optimum shape as is practicable commensurate with non-destructive testing i.e., ultrasonic, magnetic particle and penetrant inspection. With turbine and compressor blades, the accurately produced thin airfoil sections with varying degrees of camber and twist, in a variety of alloys, entails a high standard of precision forging, ret. fig. 22-2. Nevertheless precision forging of these blades is a recognised practice and enables one to be produced from a shaped die with the minimum of further work. Fig. 22-2 Precision forging. 231

ManufactureFig. 22-3 Method of producing an engine component by sand casting.232

ManufactureFig. 22-4 Automatic investment casting. 17. Forging calls for a very close control of the temperature during the various operations. An16. The high operating temperatures at which the exceptionally high standard of furnace controlturbine discs must operate necessitates the use of equipment, careful maintenance and cleanliness ofnickel base alloys. The compressor discs at the rear the forging hammers, presses and dies, is essential.end of the compressor are produced from creep-resisting steels, or even nickel base alloys, because 18. Annular combustion rings can be cold forged toof the high temperatures to which they are subjected. exacting tolerances and surfaces which alleviatesThe compressor discs at the front end of the the need for further machining before weldingcompressor are produced from titanium. The higher together to produce the combustion casing.strength of titanium at the moderate operating tem-peratures at the front end of the compressor, 19. H.P. compressor casings of the gas turbinetogether with its lower weight provides a consider- engine are forged as rings or half rings which, whenable advantage over steel. assembled together, form the rigid structure of the engine. They are produced in various materials, i.e., stainless steel, titanium and nickel alloys. CASTING 20. An increasing percentage of the gas turbine engine is produced from cast components using 233

Manufacturesand casting, ref. fig. 22-3, die casting and nickel alloys that are cast by the investment castinginvestment casting techniques; the latter becoming or lost wax' technique. Directionally solidified andthe foremost in use because of its capability to single crystal turbine blades are cast using thisproduce components with surfaces that require no technique in order to extend their cyclic lives.further machining. It is essential that all castings aredefect free by the disciplines of cleanliness during 24. Figure 22-4 illustrates automatic casting used inthe casting process otherwise they could cause the production of equi-axed, directional solidified andcomponent failure. single crystal turbine blades. The lost wax process is unparalleled in its ability to provide the highest21. All casting techniques depend upon care with standards of surface finish, repeatable accuracy andmethods of inspection such as correct chemical surface detail in a cast component. The increasingcomposition, test of mechanical properties, radiolog- demands of the engine has manifested itself in theical and microscopic examination, tensile strength need to limit grain boundaries and provide complexand creep tests. internal passages. The moulds used for directional solidified and single crystal castings differ from con-22. The complexity of configurations together with ventional moulds in that they are open at both ends,accurate tolerances in size and surface finish is the base of a mould forms a socketed bayonet fittingtotally dependent upon close liaison with design, into which a chill plate is located during casting.manufacturing, metallurgist, chemist, die maker, Metal is introduced from the central sprue into thefurnace operator and final casting. mould cavities via a ceramic filter. These and orientated seed crystals, if required, are assembled23. In the pursuit of ever increasing performance, with the patterns prior to investment. Extensiveturbine blades are produced from high temperature automation is possible to ensure the wax patterns are coated with the shell material consistently by using robots. The final casting can also have their rises removed using elastic cut-off wheels driven from robot arms, ref. fig. 22-5. FABRICATION 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 low pressure compressor casings can be produced as fabricated assemblies using sheet materials such as stainless steel titanium and varying types of nickel alloys.Fig. 22-5 Robot cut-off234

Manufacture 26. Other fabrication techniques for the manufacture of the low pressure compressor wide chord fan blade comprise rolled titanium side panels assembled in dies, hot twisted in a furnace and finally hot creep formed to achieve the necessary configu- ration. Chemical milling is used to recess the centre of each panel which sandwiches a honeycomb core, both panels and the honeycomb are finally joined together using automated furnaces where an activated diffusion bonding process takes place, ref. fig. 22-6. WELDING 27. Welding processes are used extensively in the fabrication of gas turbine engine components i.e., resistance welding by spot and seam, tungsten inert gas and electron beam are amongst the most widely used today. Care has to be taken to limit the distortion and shrinkage associated with these techniques. Tungsten inert gas (T.I.G.) welding 28. The most common form of tungsten inert gas welding, fig, 22-7, in use is the direct current straight polarity i.e., electrode negative pole. This is widely used and the most economical method of producing high quality welds for the range of high strength/high temperature materials used in gas turbine engines. For this class of work, high purity argon shielding gas is fed to both sides of the weld and the welding torch nozzle is fitted with a gas lens to ensure maximum efficiency for shielding gas coverage. A consumableFig. 22-6 Wide chord fan blade Fig. 22-7 Typical tungsten inert gas construction. welding details. 235

Fig. 22-8 Tungsten inert gas welding. Manufacture four per cent thoriated tungsten electrode, together with a suitable non-contact method o! arc starting is used and the weld current is reduced in a controlled manner at the end of each weld to prevent the formation of finishing cracks. All welds are visually and penetrant inspected and in addition, welds associated with rotating parts i.e., compressor and/or turbine are radiologically examined to Quality Acceptance Standards. During welding operations and to aid in the control of distortion and shrinkage the use of an expanding fixture is recommended and, whenever possible, mechanised welding employed together with the pulsed arc technique is preferred. A typical T.I.G. welding operation is illustrated in fig. 22-8. Electron beam welding (E.B.W.) 29. This system, which can use either low or high voltage, uses a high power density beam of electrons to join a wide range of different materials and of varying thickness. The welding machine ref. fig. 22-9, comprises an electron gun, optical viewing system, work chamber and handling equipment, vacuum pumping system, high or low voltage power supply and operating controls. Many major rotating assemblies for gas turbine engines are manufac- tured as single items in steel, titanium and nickel alloys and joined together i.e., intermediate and high pressure compressor drums. This technique allowsFig. 22-9 Electron beam welding.236

ManufactureFig. 22-10 Examples of T.I.G. and E.B. welds. another (fig. 22-11). Faraday's law of electrolysis explains that the amount of chemical reactiondesign flexibility in that distortion and shrinkage are produced by a current is proportional to the quantityreduced and dissimilar materials, to serve quite of electricity passed.different functions, can be homogeneously joinedtogether. For example, the H.P. turbine stub shafts 32. In chemical forming, (fig. 22-11), the toolrequiring a stable bearing steel welded to a material electrode (the cathode) and the workpiece (thewhich can expand with the mating turbine disc. anode) are connected into a direct current circuit.Automation has been enhanced by the application of Electrolytic solution passes, under pressure, throughcomputer numerical control (C.N.C.) to the work the tool electrode and metal is removed from thehandling and manipulation. Seam tracking to ensure work gap by electrolytic action. A hydraulic ramthat the joint is accurately followed and close loop advances the tool electrodes into the workpiece tounder bead control to guarantee the full depth of form the desired passage.material thickness is welded. Focus of the beam iscontrolled by digital voltmeters. See fig. 22-10 for 33. Electrolytic grinding employs a conductiveweld examples. wheel impregnated with abrasive particles. The wheel is rotated close to the surface of theELECTRO-CHEMICAL MACHINING (E.C.M.) workpiece, in such a way that the actual metal removal is achieved by electro-chemical means. The30. This type of machining employs both electrical by-products, which would inhibit the process, areand chemical effects in the removal of metal. removed by the sharp particles embodied in theChemical forming, electro-chemical drilling and elec- wheel.trolytic grinding are techniques of electro-chemicalmachining employed in the production of gas turbine 34. Stem drilling and capillary drilling techniquesengine components. are used principally in the drilling of small holes, usually cooling holes, such as required when31. The principle of the process is that when a producing turbine blades.current flows between the electrodes immersed in asolution of salts, chemical reactions occur in which 237metallic ions are transported from one electrode to

Fig. 22-11 Electro-chemical machining. Manufacture238 Stem drilling 35. This process consists of tubes (cathode) produced from titanium and suitably insulated to ensure a reaction at the tip. A twenty per cent solution of nitric acid is fed under pressure onto the blade producing holes generally in the region of 0.026 in. diameter. The process is more speedy in operation than electro-discharge machining and is capable of drilling holes up to a depth two hundred times the diameter of the tube in use. Capillary drilling 36. Similar in process to stem drilling but using tubes produced from glass incorporating a core of platinum wire (cathode). A twenty per cent nitric acid solution is passed through the tube onto the workpiece and is capable of producing holes as small as 0.009 in. diameter. Depth of the hole is up to forty times greater than the tube in use and therefore determined by tube diameter. 37. Automation has also been added to the process of electro-chemical machining (E.C.M.) with the intro- duction of 360 degree E.G. machining of small compressor blades, ref. fig. 22-12. For some blades of shorter length airfoil, this technique is more cost effective than the finished shaped airfoil when using precision forging techniques. Blades produced by E.C.M. employ integrated vertical broaching machines which take pre-cut lengths of bar material, produce the blade root feature, such as a fir-tree, and then by using this as the location, fully E.C.M. from both sides to produce the thin airfoil section in one operation. ELECTRO-DISCHARGE MACHINING (E.D.M.) 38. This type of machining removes metal from the workpiece by converting the kinetic energy of electric sparks into heat as the sparks strike the workpiece. 39. An electric spark results when an electric potential between two conducting surfaces reaches the point at which the accumulation of electrons has acquired sufficient energy to bridge the gap between the two surfaces and complete the circuit. At this point, electrons break through the dielectric medium between the conducting surfaces and, moving from negative (the tool electrode) to positive (the workpiece), strike the latter surface with great energy; fig, 22-13 illustrates a typical spark erosion circuit. 40. When the sparks strike the workpiece, the heat is so intense that the metal to be removed is instan- taneously vaporized with explosive results. Away

ManufactureFig. 22-12 Typical automated manufacture of compressor blades.from the actual centre of the explosion, the metal is Fig. 22-13 Electro-discharge machiningtorn into fragments which may themselves be melted circuit.by 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. 239

ManufactureCOMPOSITE MATERIALS AND SANDWICH for peripheral feature attachment on C.N.C.CASINGS machining centres, which at one component load, completely machine all required features. Examples43. High power to weight ratio and low component of composite material applications are illustrated incosts are very important considerations in the design fig. 22-14.of any aircraft gas turbine engine, but when thefunction of such an engine is to support a vertical 46. Conventional cast and fabricated casings andtake-off aircraft during transition, or as an auxiliary cowlings are also being replaced by casings ofpower unit, then the power to weight ratio becomes sandwich construction which provide strength alliedextremely critical. with lightness and also act as a noise suppression medium. Sandwich construction casings comprise a44. In such engines, the advantage of composite honeycomb structure of aluminium or stainless steelmaterials allows the designer to produce structures interposed between layers of dissimilar material. Thein which directional strengths can be varied by materials employed depend upon the environment indirectional lay-up of fibres according to the applied which they are used.loads. INSPECTION45. Composite materials have and will continue toreplace casings which, in previous engines, would 47. During the process of manufacture, componenthave been produced in steels or titanium. By-pass parts need to be inspected to ensure defect freeduct assemblies comprising of three casings are engines are produced. Using automated machinerycurrently being produced up to 4ft-7in. in diameter and automated inspection, dimensional accuracy isand 2ft-0in in length using pre-cured composite maintained by using multi-directional applied probesmaterials for the casing fabric. Flanges and mounting that record sizes and position of features. The C.N.C.bosses are added during the manufacturing process, inspection machine can inspect families ofwhich are then drilled for both location and machined components at pre-determined allotted intervalsFig. 22-14 Some composite material applications.240