上海交通大学：《飞机设计 Aircraft Design》课程教学资源_参考资料_36_engine options for supersonic cruise aircraft AIAA-1978-1054-507

Sponsored by- American Institute of Aeronautics and Astronautics(AIAA) Society of Automotive Engineers(SAE) 78-1054 Engine Options for Supersonic Cruise Aircraft P.H.Calder and P.C.Gupta,Bristol Engine Division,Bristol,UK AIAA/SAE 14TH JOINT PROPULSION CONFERENCE Las Vegas,Nev./July 25-27,1978 For permission to copy or republish,contact the American Instltute of Aeronautics and Astronautics, 1290 Avenue of the Americas,New York,N.Y.10019

Sponsored by￾American Institute of Aeronautics and Astronautics (AIAA) Society of Automotive Engineers (SAE) 78-1054 Engine Options for Supersonic Cruise Aircraft c P. H. Calder and P. C. Gupta, Bristol Engine Division, Bristol, UK AIANSAE 14TH JOINT PROPULSION CONFERENCE Las Vegas, Nev.lJuly 25-27, 1978 For permission to copy or republish. contact the Amerlcan lnstltute of Aeronautics and Astronautics, 1290 Avenue of the Americas, New Vork. N.V. 10019

ENGINE OPTIONS FOR SUPERSONIC CRUISE AIRCRAFT Dr P H Calder,OBE Project Director Olympus 593 and P C Gupta Assistant Chief Engineer-Performance Rolls-Royce Limited,Aero Division,Bristol Abstract integrated with the development of the variable throat,variable divergence, The paper summarises the Supersonic thrust reversing exhaust system,which uses Cruise engine studies that have been intake throat bleed for ventilation. These carried out as part of the UK contribution have been four-company activities as has to the ICAO Noise Working Group activities, been the flight clearance of the powerplant and presents the results of their inte- to the stringent airworthiness criteria laid gration in the Aircraft Parametric Study by down jointly by the British and French British Aerospace.Duct burning and Authorities, variable cycle engines are also discussed. The recent Rolls-Royce experimental work on If the cowlings of one of the rectangu- co-annular nozzles and mechanical ejector lar paired nacelles are removed,two square suppressors is reviewed and an international engines are visiblel This is the result of collaborative programme on mechanical tailoring the mass of engine dressings, ejector suppressor testing is described. It is shown how the choice of engine cycle cables and pipework to the nacelles.See Figure1· depends on the characteristics of the air- craft,the amount of silencing available and the level of technology available at go-ahead. NOT TO BE USED. THIS FL.AP ERATW Introduction Concorde supersonic passenger services from Paris and London to South America and the Middle East commenced nearly two and a half years ago and were extended to North America two years ago,initially to Washington DC and then to New York. At the end of April 1978 some 54 000 hours of engine service experience had been accumulated in the course of 4500 flights. Concorde and its powerplant have met their payload range performance objectives. The aircraft had made convincing demonstra- tions,even before the commencement of ser- vice into New York,that it had met its design aim of being no noisier than narrow bodied aircraft (707,DC8,VC10).In fact under certain circumstances,and due to its Fig 1.Concorde Engine Bay ability to adapt to particular airport geography,Concorde can produce community The last of the batch of engines ordered noise levels which are appreciably lower for the first sixteen Concorde aircraft was than those of narrow bodied aircraft. passed off the test bed in December 1977. Early flight development engines continue The theme of this session is SST to be converted to production standard for Propulsion System Airframe Integration. spares. The Concorde is a flying example of such integration,From the very earliest design It is appropriate at this time to examine stages the integration of the engine and the performance of the production run of powerplant with the airframe has been a new engines.Figure 2 shows that the prime consideration,with all relevant guaranteed thrust and mission equivalent decisions being mutually agreed by the four specific.fuel consumption have been met Companies responsible,The development of with comfortable margins and that the mean the variable intake to match the engines, performance has achieved the brochure target and control of the engine to match the which was used for aircraft mission per- intake under certain flight conditions,was formance predictions. Copyrightmerican Institute of Aeronautics and stronautics,Inc.,1978.All rights reserved

ENGINE OPTIONS FOR SUPERSONIC CRUISE AIRCRAFT Dr P H Calder, OBE Project Director Olympus 593 and P C Gupta Assistant Chief Engineer - Performance Rolls-Royce Limited, Aero Division, Bristol Abstract The paper summarises the Supersonic Cruise engine studies that have been carried out as part of the UK contribution to the ICAO Noise Working Group activities, and presents the results of their inte￾gration in the Aircraft Parametric Study by British Aerospace. Duct burning and variable cycle engines are also discussed. The recent Rolls-Royce experimental work on co-annular nozzles and mechanical ejector suppressors is reviewed and an international collaborative programme on mechanical ejector suppressor testing is described. It is shown how the choice of engine cycle depends on the characteristics of the air￾craft, the amount of silencing available and the level of technology available at go-ahead. Introduction Concorde supersonic passenger services v from Paris and London to South America and the Middle East commenced nearly two and a half years ago and were extended to North America two years ago, initially to Washington Dc and then to New York. At the end of April 1978 some 54 000 hours of engine service experience had been accumulated in the course of 4500 flights. Concorde and its powerplant have met their payload range performance objectives. The aircraft had made convincing demonstra￾tions, even before the commencement of ser￾vice into New York, that it had met its design aim of being no noisier than narrow bodied aircraft (707, Dc8, VClO). In fact, under certain circumstances, and due to its ability to adapt to particular airport geography, Concorde can produce community noise levels which are appreciably lower than those of narrow bodied aircraft. The theme of this session is SST Propulsion System - Airframe Integration. The Concorde is a flying example of such integration. From the very earliest design stages the integration of the engine and powerplant with the airframe has been a prime consideration, with all relevant decisions being mutually agreed by the four Companies responsible. The development of the variable intake to match the engines, intake under certain flight conditions, was / and control of the engine to match the 1 Copyrl@lO American in~lilutc of AeroniUlks *ad Astronm~ica. In<.. 1978. All rights resewed. integrated with the development of the variable throat, variable divergence, thrust reversing exhaust .syrtAm-! -which uses intake throat bleed for ventilation. These have been four-company activities as has been the flight clearance of the powerplant to the stringent airworthiness criteria laid down jointly by the British and French Authorities. If the cowlings of one of the rectangu￾lar paired nacelles are removed, two square engines are visible! This is the result of tailoring the mass of engine dressings, cables and pipework to the nacelles. See Figure 1. Fig 1. Concorde Engine Bay The last of the batch of engines ordered for the first sixteen Concorde aircraft was passed off the test bed in December 1977. Early flight development engines continue to be converted to production standard for spares. It is appropriate at this time to examine the performance of the production run of new engines. Figure 2 shows that the guaranteed thrust and mission equivalent specific fuel consumption have been met with comfortable margins and that the mean performance has achieved the brochure target which was used for aircraft mission per￾formance predictions

》4CH2SA+5吃GR%EnRs7 MEAN-1032 号 A∽ L96 答9n7和m+院 N.26r 注6%》 BIO MIIMLIM ACCEPONCE OMARIMM ACEPTANE BROCHARE NOMINAL MN,--/3% 05 ENEINE SERAL N9 845se47 Fig 2. Olympus 593 Mk 610 Performance of Fig 3. Effect of Cycle on Payload at Fixed New Production Engines (Assessed at Range Nominal Control Schedule Settings) Only two of the engines were rejected on datum value,results in a loss in payload. the first acceptance run,due in both cases The turbojet had optimum performance,a to low TET,and hence thrust,at limiting conclusion that had been reached indepen- HP spool rpm or measured EGT. In both dently by Government establishments. cases a change in component enabled the Concorde was designed to have no higher engine to be cleared. noise levels than the subsonic narrow bodied jets (707,VC10 etc)which were expected to The successor to Concorde must not only form the major part of airport movements at be economically viable,but must be the envisaged entry-into-service date. The environmentally acceptable. In their 1977 results of monitoring of Concorde operations paper to the AIAA/SAE Propulsion Conference into Washington and New York have confirmed the authors reviewed the application of new that this target has been achieved. technology to meet these aims,and discussed the constraints involved with particular Considerations of optimisation of mission reference to pollution and noise. Prior to fuel,including the quite considerable sub- that date ICAO had addressed the matter of sonic element in the reserves,led to the noise regulations for present and future choice of overall pressure ratio adopted. supersonic transport aircraft,and Apart from the engine cycle,other engineer- initiated a number of studies which have ing considerations were of major importance. been proceeding over the last year. This The Olympus family of engines had a large paper gives the results of some of the UK background of experience,and a supersonic contributions to these studies which are version was already under development for relevant to the choice of engine for a a military tactical strike aircraft (TSR2) supersonic cruise aircraft,and goes on to which was later cancelled,but not before discuss further options such as duct the engines had flown, burning and cycle variability in the light of these studies.The paper also summarises Furthermore,the Bristol part of the recent and proposed experimental work on Rolls-Royce organisation (Bristol Siddeley silencing, covering the results of Engines Limited as it was then known) co-annular nozzle noise measurements, and already had collaborative arrangements with including some details of a proposed the French engine company SNECMA. international collaboration programme to advance the technology of ejector These considerations weighed heavily in suppressors. favour of the choice of an olympus derivative,once the choice of engine cycle had been made,During the design and Choice of Engine for Concorde development of the engine and powerplant the Figure 3,which is reproduced from maximum possible amount of variability was Reference 1,shows the relationship trends incorporated to obtain the best performance and lowest noise within the constraints of between payload,TET,bypass ratio and jet velocity. It takes into account both the available technology. specific fuel consumption and powerplant weight and drag effects using Concorde as Figure 3 also shows how,as technology a datum at fixed range. advances,the use of increased TET would permit the bypass ratio to be increased and jet velocity reduced without sacrifice in At the cruise turbine entry temperatures (ueowethatn payload,and that an improved payload is obtained with higher TET if the jet velocity addition of bypass,to reduce jet velocity is held constant,or allowed to rise. at take-off and flyover (cut-back)from the 2

Fig 2. Olympus 593 Mk 610 Performance of New Production Engines (Assessed at Nominal Control Schedule Settings) Only two of the engines were rejected on the first acceptance run, due in both cases to low TET, and hence thrust, at limiting HP spool rpm or measured EGT. In both cases a change in component enabled the engine to be cleared. ' The successor to Concorde must not only be economically viable, but must be environmentally acceptable. In their 1977 I paper to the AIAA/SAE Propulsion Conference the authors reviewed the application of new technology to meet these aims, and discussed the constraints involved with particular reference to pollution and noise. Prior to that date ICAO had addressed the matter of noise regulations for present and future supersonic transport aircraft, and initiated a number of studies which have been proceeding over the last year. This paper gives the results of some of the UK contributions to these studies which are relevant to the choice of engine for a supersonic cruise aircraft, and goes on to discuss further options such as duct burning and cycle variability in the light of these studies. The paper also summarises recent and proposed experimental work on silencing, covering the results of co-annular nozzle noise measurements, and including some details of a proposed international collaboration programme to advance the technology of ejector suppressors. Choice of Encline for Concorde Figure 3, which is reproduced from Reference 1, shows the relationship trends between payload, TET, bypass ratio and jet velocity. It takes into account both specific fuel consumption and powerplant weight and drag effects using Concorde as a datum at fixed range. At the cruise turbine entry temperatures envisaged at Cgncorde's early design stage (135OOK - 1400 K) Figure 3 shows that any addition of bypass, to reduce jet velocity at take-off and flyover (cut-back) from the 2 Fig 3. Effect of Cycle on Payload at Fixed Range datum value, results in a loss in payload. The turbojet had optimum performance, a conclusion that had been reached indepen￾dently by Government establishments. Concorde was designed to have no higher noise levels than the subsonic narrow bodied jets (707, VClO etc) which were expected to form the major part of airport movements at the envisaged entry-into-service date. The results of monitoring of Concorde operations into Washington and New York have confirmed that this target has been achieved. Considerations of optimisation of mission fuel, including the quite considerable sub￾sonic element in the reserves, led to the choice of overall pressure ratio adopted. Apart from the engine cycle, other engineer￾ing considerations were of major importance, The Olympus family of engines had a large background of experience, and a supersonic version was already under development for a military tactical strike aircraft (TSRZ) which was later cancelled, but not before the engines had flown. Furthermore, the Bristol part of the Rolls-Royce organisation (Bristol Siddeley Engines Limited as it was then known) already had collqborative arrangements with the French engine company SNECMA. These considerations weighed heavily in favour of the choice of an Olympus derivative, once the choice of engine cycle had been made. During the design and development of the engine and powerplant the maximum possible amount of variability was incorporated to obtain the best performance and lowest noise within the constraints of the available technology. Figure 3 also shows how, as technology advances, the use of increased TET would permit the bypass ratio to be increased and jet velocity reduced without sacrifice in payload, and that an improved payload is obtained with higher TET if the jet velocity is held constant, or allowed to rise

Advancements in Technology The United Kingdom study is being under- taken by British Aerospace (Filton and In Reference 1 the authors have reviewed Weybridge)with engine data supplied by the application of advancements in tech Rolls-Royce.In this study British nology to the performance improvement and Aerospace are evaluating operating costs noise reduction of supersonic transport and noise levels of a family of aircraft. aircraft.It is shown how performance The study covers variations in: requirements for subsonic aircraft have led to the adoption of low specific thrust, Range high bypass ratio engines,with a consequent Number of Passengers reduction in jet velocities and hence in Wing Area airport noise.It is also shown that this Aspect Ratio is not the case for supersonic cruise air- Engine Bypass Ratio craft where increasing bypass ratio beyond a certain value will degrade the mission Class II technology is being assumed,with performance (Figure 3). For such aircraft a cruise Mach No of 2.0. the requirements of performance and noise are conflicting. Engine Studies The various means of improving perfor- To provide the required engine data a mance and reducing noise were discussed in preliminary design study and part load per- some detail,in the Reference,and the con- formance estimate was made of four mixed straints to their application described. flow turbofan engines.These were two spool engines with the fan on the LP spool, In the next few sections the studies and the remainder of the core compression that have recently been made of engines for being carried by the HP spool,which was the next generation of supersonic trans- envisaged as an Olympus 593 HP spool with ports are described.These studies have some extra stages of compression. As is assumed a defined advancement in the level the case for the olympus 593,sufficient of technology available. nozzle area variation is provided to enable the appropriate LP spool rpm and TET ICAO Studies to be obtained simultaneously over a large range of flight conditions.This is an In November 1976 the ICAO Committee on essential feature of a supersonic cruise Aircraft Noise,CAN 5,set up Working engine. The ground rules for the study are Group E with the task of reporting to the given in Table 1,and some leading par- Committee (CAN 6)in late 1978 on recom- ticulars of the resulting engines are given mendations for SST noise regulations and in Table 2. guide lines.France,UK,US and USSR (as the countries who have a direct interest in SST development)have representation rights Cruise Design Point on the Working Group. Same core flow and overall pressure ratio as olympus 593 As part of their programme Working 16000K TET Group E set up a Parametric Study sub-group with the somewhat formidable task of evalu- Chosen bypass ratio ating the relationship between the noise Equal Mach Number and static pressure at mixing plane level and the operating cost of a supersonic (leads transport.Each country was to perform its to low operating point on fan chic) own studies,and report to the sub-group who would correlate the studies and report to Take-off Operation the Working Group,To assist in the cor- Cruise to take-off fan rpm relation simi- relation further sub-groups were set up, lar to olympus 593 LP compressor one on noise prediction,with the aim of 17000K TET achieving a common prediction method,and Highest safe operating point on fan chic another on economics having the task of Bypass ratio and overall pressure ratio agreeing a common method for cost follow estimation. NB Improved cruise performance is obtained Working Group E,when initiating the by allowing high operating point on fan Parametric Studies,made a forecast of the chic with unequal mixing Mach numbers timescale in which it expected the various technologies would be available for com- mencing a serious design.The technologies Table 1.-Mixed Turbofan Family m Engines were classified as: A,B,C.D.Ground Rules set at the Beginning of the Study Class I 1977-1980 Class II 1980-1985 Evident from Table 2 is the significant Class III Beyond 1985 reduction in bypass ratio from cruise to take-off which has resulted from the use The detailed classification of the various of variable bypass and optimising the cycle technologies is given by Foster in Table I matching during the course of the study. of Reference 2. In fact engine A is virtually a pure jet at 3

Advancements in Technoloqy In Reference 1 the authors have reviewed the application of advancements in tech￾nology to the performance improvement and noise reduction of supersonic transport requirements for subsonic aircraft have led to the adoption of low specific thrust, high bypass ratio engines, with a consequent reduction in jet velocities and hence in airport noise. It is also shown that this is not the case for supersonic cruise air￾craft where increasing bypass ratio beyond a certain value will degrade the mission performance (Figure 3). For such aircraft the requirements of performance and noise are conflicting. - aircraft. It is shown how performance The various means of improving perfor￾mance and reducing noise were discussed in some detail, in the Reference, and the con￾straints to their application described. In the next few sections the studies that have recently been made of engines for the next generation of supersonic trans￾ports are described. These studies have assumed a defined advancement in the level of technology available. ICAO Studies In November 1976 the ICAO Committee on Aircraft Noise, CAN 5, set up Working Group E with the task of reporting to the Committee (CAN 6) in late 1978 on recom- mendations for SST noise regulations and the countries who have a direct interest in SST development) have representation rights on the Working Group. Y guide lines. France, UK, US and USSR (as As part of their Droaramme Workina GrouD E set UD a Parametric Studv Sub-Orom .,I with' the SomEwhat formidable task of evalu￾ating the relationship between the noise level and the operating cost of a supersonic transport. Each country was to perform its own studies, and report to the sub-group who would correlate the studies and report to the Working Group. To assist in the cor￾relation further sub-groups were set up, one on noise prediction, with the aim of achieving a common prediction method, and another on economics havino the task of ~ agreeing a common method for cost estimation. Working Group E, when initiating the Parametric Studies. made a forecast of the The United Kingdom study is being under￾taken by British Aerospace (Filton and Weybridge) with engine data supplied by Rolls-Royce. In this study British Aerospace are evaluating operating costs and noise levels of a family of aircraft. The study covers variations in: Range Number of Passengers Wing Area Aspect Ratio Engine Bypass Ratio Class I1 technology is being assumed, with a cruise Mach No of 2.0. Enaine Studies To provide the required engine data a preliminary design study and part load per￾formance estimate was made of four mixed flow turbofan engines. These were two spool engines with the fan on the LP spool, and the remainder of the core compression being carried by the HP spool, which was envisaged as an Olympus 593 HP spool with some extra stages of compression. As is the case for the Olympus 593, sufficient nozzle area variation is provided to enable the appropriate LP spool rpm and TET to be obtained simultaneously over a large range of flight conditions. This is an essential feature of a supersonic cruise engine. The ground rules for the study are qiven in Table 1. and some leadina -. Dar￾ticulars of the Gesulting engines are given in Table 2. Cruise Design Point Same core flow and overall pressure ratio as Olympus 593 16000K TET Chosen bypass ratio Equal Mach Number and static pressure at mixing plane (leads to low operating point on fan chic) Take-off ODeration Cruise to take-off fan rpm relation simi￾lar to Olympus 593 LP compressor 17000K TET Highest safe operating point on fan chic Bypass ratio and overall pressure ratio follow E Improved cruise performance is obtained by allowing high operating point on fan chic with unequal mixing Mach numbers timescale in which'it expected the various technologies would be available for com- mencing a serious design. The technologies were classified as: A. B, C, D Ground Rules set at the Table 1 - Mixed Turbofan Family - Ensines Beqinninq of the Study Class I 1977-1980 Class I1 1980-1985 Evident from Table 2 is the significant Class 111 Beyond 1985 reduction in bvDass ratio from cruise to take-off which-kas resulted from the use The detailed classification of the various of variable bypass and optimising the cycle technologies is given by Foster in Table I matchinq durinq the course of the study. of Reference 2. In fact-engine-A is vihually a pure jet at 3

Engine A B C D Total airflow relative to datum 1.5 2.0 2.5 3.0 Design A,at beginning of study,= 0.5 1.0 1.5 2.0 Cruise Operating A,resulting from 0.291 0.762 1.240 1.716 optimum matching Cruise Thrust (1bf) 14500 14520 14410 14260 Cruise SFC relative to datum 0,990 0.972 0.967 0.970 Take-off operating入，resulting from 0.049 0.459 0.885 1,314 optimum matching Dry take-off thrust (1bf) 48280 50150 51370 53850 Dressed engine weight relative to datum 1.458 1.678 1.817 1,986 A=Bypass ratio A*=Design bypass ratio Quoted Cruise Mach 2.0,ISA 50C,53 000 ft Quoted Take-off 250 kt,ISA 100C,900 ft Quoted weight includes bypass duct,excludes mixer and primary nozzle Reheat for take-off if required Datum is Olympus 593 Mk 610 Table 2-Performance and Weight of Turbofan Family take-off.This variation of bypass ratio In order to provide British Aerospace is a natural feature of the layout described with a purejet tend point!for their where the fan and the remainder of the core studies,a further engine was considered. compression are on different spools and The leading particulars of this engine are nozzle variation is provided. It enables given in Table 3,and are based on a pro- the fan to be matched at high pressure ratio posal for a 15%flow increase over the for take-off.In contrast a turbojet,with- current engine.This is obtained by a mod- out the facility for allowing a sizeable ification to the first three stages of the quantity of LP compressor delivery air to current LP compressor.This modification bypass the HP system at cruise,needs to has been rig tested to show the required be matched at a low LP compressor pressure flow increase,and an earlier version has ratio at take-off to avoid LP compressor surge at cruise. Since the fan engines are matched at a Object: high fan pressure ratio at take-off rela- To provide BAe with "end point"for calcu- tive to a turbojet,the core has a rela- lations. tively higher degree of supercharge,and hence although the design core flow at Chosen Engine: cruise of the family of fan engines is the same as the olympus 593,the take-off core High temperature version of 1976 Olympus flow is some 30%to 40%higher.As a con- 593 Mk 622 proposal.Take-off airflow 15% sequence the dry take-off thrust is con- above Mk 610. siderably higher than would be given by an Olympus 593 turbojet operating at the same Leading Particulars: temperature.The necessity for reheat Cruise Mach 2.0 1SA+50C,53000ft boost at take-off is therefore minimised or TET 16000K eliminated,with consequent advantages in simplification of control equipment and Thrust 129801bf improved pollution characteristics. This SFC 1.034 x Datum also permits a shorter and hence lighter Take-ofE-250Kt,ISA+100℃，900ft exhaust system design.Eliminating reheat TET 17000K at take-off facilitates the use of ejector- type noise suppressors,as these would not Dry Thrust 377801bf easily withstand the peaky temperatures Reheated Thrust 49 300 1bf (18000K Reheat) involved in afterburner operation. Bare Engine Weight Plus Dressings 1.046 ×Datum(Datum is O1 ympus593Mk610） A more detailed discussion of the match- ing of the olympus 593 and various turbo- fan layouts has'been given by the authors 3. Table 3-Turbojet (=O)Engine

Enqine A B C D I Quoted weight includes bypass duct, excludes mixer and primary nozzle ~ 1.0 1.5 2.0 ~ ____ Design A , at beginning of study, = A * 0.5 Cruise Operating h , resulting from 0.291 0.762 1.240 1.716 optimum matching . Cruise Thrust. (lbf) 14 500 14 520 14 410 14 260 0.990 0.972 0.967 0.970 __ ~~ ... ~ . .. Cruise SFC relative to datum optimum matching Dry take-off thrust (lbf) 48 280 50 150 51 370 53 850 Dressed engine weight relative to datum 1.458 1.678 1.817 1.986 ~ _________ Take-off operating A , resulting from 0.049 0.459 0.885 i.314 - ~- .~ ~ Reheat for take-off if required Datum is Olympus 593 Mlc 610 Table 2 - Performance and Weight of Turbofan take-off. This variation of bypass ratio is a natural feature of the layout described where the fan and the remainder of the core compression are on different spools and nozzle variation is provided. It enables the fan to be matched at high pressure ratio for take-off. In contrast a turbojet, with￾out the facility for allowinq a sizeable quantity of LP compressor deiivery air to bypass the HP system at cruise, needs to be matched at a low LP compressor pressure ratio at take-off to avoid LP compressor surge at cruise. Since the fan engines are matched at a high fan pressure ratio at take-off rela￾tive to a turbojet, the core has a rela￾tively higher degree of supercharge, and hence although the design core flow at cruise of the family of fan engines is the same as the Olympus 593, the take-off core flow is some 30% to 40% higher. As a con- sequence the dry take-off thrust is con￾siderably higher than would be given by an Olympus 593 turbojet operating at the same temperature. The necessity for reheat boost at take-off is therefore minimised or eliminated, with consequent advantages in simplification of control equipment and improved pollution characteristics. This also permits a shorter and hence lighter exhaust system design. Eliminating reheat at take-off facilitates the use of ejector￾type noise suppressors, as these would not easily withstand the peaky temperatures involved in afterburner operation. A more detailed discussion of the match￾ing of the Olympus 593 and various turbo￾fan layouts has been given by the authors3 Family In order to provide British Aerospace with a purejet 'end point' for their studies, a further engine was considered. The leading particulars of this engine are given in Table 3,and are based on a pro￾posal for a 15% flow increase over the current engine. This is obtained by a mod￾ification to the first three stages of the current LP compressor. This modification has been rig tested to show the required flow increase, and an earlier version has w: To provide BAe with "end point" for calcu￾lations. 3hosen Enqine: Hiah temDerature version of 1976 01vmDuS 593 Mk 6'22 proposal. above Mk 610. Take-off airfiow 154 -eadincr Particulars: Cruise - Mach 2.0 ISA + 5OC, 53 000 ft TET 1600°K Thrust 12 980 lbf SFC 1.034 x Datum rake-off - 250 Kt, ISA + 10°C, 900 ft TET 170O0K Dry Thrust 37 780 lbf Reheated Thrust 49 300 lbf (18000K Reheat) aare Enqine Weioht Plus Dressinqs 1.046 x Datum (Datum is Olympus 593 Mk 610) Table 3 - Turbojet (A= 0) Encline 4

been run on an engine to demonstrate a 10% flow increase.The choice for this study Condition 1 2 3 of a high temperature version of this engine rather than the basic olympus 593 Mach number 2.0 1.2 0,93 0.50 was dictated by a desire to meet a take-off thrust of around 50 000 1b within a reheat Altitude（ft) 53000 40000 25000 15000 temperature of 1800K. Assumed 0.9371 0.9858 0,99 0.99 The relative weights of these engines are intake PRF shown on Figure 4 as a function of flow increase over the Olympus 593.Also shown Net thrust 14610 19430 11980 9940 is the effect of scaling up the turbojets. (1bf) It will be seen that to increase flow by increasing bypass ratio is 'a considerably Relative lighter route than scaling up an engine. importance in 73,8% 10.3% 6.6% 9.2% mission fuel Condition 1:supersonic acceleration and 20n cruise Condition 2: transonic acceleration RESSED Condition 3: ENSINE diversion cruise WEICHT Condition 4:hold 乳A7ME 心Y 5 Table 4 Flight Conditions for Mission Fuel Calculation (ISA 50C) 卜/25 OWER L4WS4kE 4材-2口 1550 SOT LINE CONSISTENT WITH ⊙7九RET产 FI0 32 OF SAE 75/056 a格3 ▲my9pEs9is 10 5 20 25 30 COWEPALANT L 4s% 550K57 7龙LA/RFLOW RELA7e7DMMP然B93 15s0 元B2F4ND SOT Fig 4. Estimated Weight of Engines for ICAO Studies 2 It is of interest to compare the results of powerplant plus fuel weight calculations MEW DTZOM for the family of engines studied with the - -- results of earlier parametric studies such as Figure 3 which are based on simple cycle calculations and weight formulae. 2 OPERATINS CRUISE PYAASS RATO 05 格 The engines have been scaled to the common cruise thrust requirement of the Fig 5. Comparison of ICAO Study Engines 700 000 1b TOGW;4200 nm range,Mach 2.0 with Original Parametric Studies cruise aircraft model described in Reference 3.The internal engine perfor- Take-off and flyover noise estimates have mance was evaluated at the flight condi- been made for the family of engines by a tions listed in Table 4 to determine the Rolls-Royce method,The results are shown fuel weight for a typical mission. The for typical thrust levels on Figure 6 and relative importance of each flight condi- Figure 7 as a function of mass flow relative tion in determining mission fuel weight is to Olympus 593.Noteworthy is the diminish- also given in Table 4. ing return of increasing mass flow at the flyover condition.The actual noise scales The full line of Figure 5 shows the have been omitted,as this and other methods results of a parametric study similar to are being reviewed by the ICAO Working Group Figure 3 and in fact using identical noise sub-group,and it would be undesirable assumptions to Figure 32 of Reference 4, to pre-empt the conclusion of the sub-group. except that a 1600 K TET (1550K stator outlet temperature)has been used.The By combining Figures 6 and 7 with the individually plotted points show the mission calculation it is possible to results for the family of engines here relate powerplant plus fuel weight described.The datum is the current with total air flow and noise level,as Olympus 593 scaled to suit the new aircraft shown on Figure 8,Figure 8 also shows the model.It will be seen that the trends of benefit to be obtained from increasing the earlier parametric study are closely cruise TET beyond the Class II level to confirmed by the current more detailed 1700K together with an increase of around evaluation. 3%in HP spool rpm.This corresponds to

been run on an engine to demonstrate a 10% flow increase. The choice for this study of a high temperature version of this engine rather than the basic Olympus 593 was dictated by a desire to meet a take-off thrust of around 500000 lb within a reheat temperature of 1800 K. - The relative weights of these engines are shown on Figure 4 as a function of flow increase over the Olympus 593. Also shown is the effect of scaling up the turbojets. It will be seen that to increase flow by increasing bypass ratio is a considerably lighter route than scaling up an engine. Fig 4. Estimated Weight of Engines for ICAO - Studies It is of interest to compare the results of powerplant plus fuel weight calculations for the family of engines studied with the results of earlier parametric studies such as Figure 3 which are based on simple cycle calculations and weight formulae. The engines have been scaled to the common cruise thrust requirement of the 700 000 lb TCGW; 4200 nm range, Mach 2.0 cruise aircraft model described in Reference 3. The internal engine,perfor- mance was evaluated at the flight condi￾tions listed in Table 4 to determine the fuel weight for a typical mission. The relative importance of each flight condi￾tion in determining mission fuel weight is also given in Table 4. The full line of Figure 5 shows the results of a parametric study similar to Figure 3 and in fact using identical assumptions to Figgre 32 of Reference 4, except that a 1600 K TET (155OoK stator outlet temperature) has been used. The individually plotted points show the ,results for the famil'y of engines here described. The datum is the current Olympus 593 scaled to suit the new aircraft model. It will be seen that the trends of the earlier parametric study are closely Eonfirmed by the current more detailed evaluation. J ~ Condition Mach number Altitude (ft) Assumed intake PRF Net thrust (lbf) Relative importance in mission fuel 73.8% 10.3% 6.6% ~ ILL 4 _____ 0.50 15 O( __ ~ 0.99 ~ 9940 ~ 9.2% .- Condition 1: sunersonic acceleration and cruise Condition 2: transonic acceleration Condition 3: diversion cruise Condi.tion 4: hold Table 4 - Fliqht Conditions for Mission Fuel Calculation (ISA + 50CL Fig 5. Comparison of ICAO Study Engines with Original Parametric Studies Take-off and flyover noise estimates have been made for the family of engines by a Rolls-Royce method. The results are shown for typical thrust levels on Figure 6 and Figure 7 as a function of mass flow relative to Olympus 593. Noteworthy is the diminish￾ing return of increasing mass flow at the flyover condition. The actual noise scales tiave been omitted, as this and other methods are being reviewed by the ICAO Working Group noise sub-group, and it would be undesirable to pre-empt the conclusion of the sub-group. By combining Figures 6 and 7 with the mission calculation it is possible tb relate powerplant plus fuel maight with total air flow and noise level, as shown on Figure 8. Figure 8 also .shows the benefit to be obtained from increasing cruise TET beyond the Class I1 level to 1700 K together with an increase of around 3% in, HP spool rpm. This corresponds to 5

MACH 03 ISA 10%C 6…40H公2必 DECREASING MOISE SdB ITERVLS- 名Edit时 SCALED 595 心老 62 X7a北RFLOW RELA公和PS53 Fig 6. AST Take-off (Sideline)Noise-Four Fig 8. Payload Total Airflow and Noise Engines,2310 ft Slant Range Comparison 175 ALTITUDE Integrated Aircraft/Engine Study POUR ENGINE FREE FIELD .Whilst the mission fuel plus powerplant weight studies outlined above can give the relative merits of self-consistent engine proposals,for accurate assessments there is no substitute for a combined aircraft/ engine study where the aircraft design can be modified and optimised in accordance with the characteristics of the powerplants, which can also be tailored to the specific aircraft. 25 As mentioned earlier,British Aerospace have been conducting such studies 20 EPNdB 30 as part of their contribution to the work +©% of the ICAO Parametric Study sub-group,and 2时 some preliminary results have been published in Reference 5. TAM NET7HRT=CO他E4落T包) The British Aerospace project assessment programme used for these studies can handle Fig 7.AST Cutback (Flyover)Noise a range of aircraft geometries,mission ranges,passenger payloads and engine Iturning up the tapt on each engine of the designs. Noise at each of three measuring family. The increased cruise thrust capa- points defined by ICAO 1971 Annex 16,based bility gives a smaller engine,but it on measured Concorde noise levels,is calcu- should be noted that the transonic regime lated, The noise calculations are in the will require a proportionately higher TET course of being checked by methods proposed increase to give the thrust. If this is by the ICAO noise sub-group,and the noise not practical,some reheat may be required scale has been omitted,l0dB bandwidth and the benefit of increased TET will be intervals being shown on the curves to indi- eroded,Furthermore,it must be remembered cate trends.No special silencing means, that the benefits of increased TET can also apart from the use of acoustic liners,has be eroded by a higher replacement cost for been assumed for these curves. hot end parts. Figures 9 and 10 reproduced from Reference 5 show the results of the cal- Figure 8 also shows the effect of changing the weight fraction of the in- culations for fixed range and fixed payload takes,nacelles,and secondary nozzles of respectively for the engines discussed above. The engines are sized to the cruise the datum engine powerplant from the thrust requirement of each aircraft,and Concorde value to 0,6 of the Concorde reheat boost is assumed for take-off,where value,which would be more appropriate to necessary.As expected,the lower noise second generation axisymmetric pods. The levels require higher bypass ratios,and penalty for increased mass flow is result in increasing gross weight for a diminished.However,Figure 8 ignores given range and mission, installation loss,(eg skin friction etc) which will tend to offset this trend,as is A very significant feature of these shown in Figure 3.It should be noted that curves is how the lower bound tends to if this change in datum weight fraction become vertical for a given range and could be achieved without any structural or payload.Thus there is a minimum noise other weight penalty,an improvement in level that cannot be bettered at a given payload of 27%would be obtained. level of technology. 6

Fig 6. AST Take-off (Sideline) Noise -Four Engines, 2310 ft Slant Range a. W6,NE FEE WCLD N75 AL77WDE Fig 7. AST Cutback (Flyover) Noise 'turning up the zap' on each engine of the family. The increased cruise thrust capa￾bility gives a smaller engine, but it should be noted that the transonic regime will require a proportionately higher TET increase to give the thrust. If this is not practical, some reheat may be required and the benefit of increased TET will be eroded. Furthermore, it must be remembered that the benefits of increased TET can also be eroded by a higher replacement cost for hot end parts. Figure 8 also showsthe effect of changing the weight fraction of the in￾takes, nacelles, and secondary nozzles of the engine powerplant from the .Concorde value to 0.6 of the Concorde value, which would be more appropriate to second generation axisymmetric pods. The penalty for increased mass flow is diminished. However, Figure 8 ignores installation loss, (eg skin friction etc) which will tend to offset this trend, as is shown in Figure 3. It should be noted that if this change in datum weight fraction could be achieved without any structural or other weight penalty, an improvement in payload of 27% would be obtained. Fig 8. Payload Total Airflow and Noise Comparison Intearated Aircraft/Enqine Studv .Whilst the mission fuel plus powerplant weight studies outlined above can give the relative merits of self-consistent engine proposals, for accurate assessments there is no substitute for a combined aircraft/ engine study where the aircraft design can be modified and optimised in accordance with the characteristics of the powerplants, which can also be tailored to the specific aircraft. As mentioned earlier, British Aerospace have been conducting such studies as part of their contribution to the work of the ICAO Parametric Study sub-group, and some preliminary results have been published in Reference 5. ~ The British Aerospace project assessment programme used for these studies can handle a range of aircraft geometries, mission ranges, passenger payloads and engine designs., Noise at each of three measuring points defined by ICAO 1971 Annex 16, ba~sed on measured Concorde noise levels, is calcu￾lated. The noise calculations are in the course of being checked by methods proposed by the ICAO noise sub-group, and the noise scale has been omitted, lOdB bandwidth intervals being shown on the curves to indi￾cate trends. No special silencing means, apart from the use of acoustic liners, has been assumed for these curves. Reference 5 show the results of the cal￾culations for fixed range and fixed payload respectively for the engines discussed above. The engines are sized to the cruise thrust requirement of each aircraft, and I-eheat boost is assumed for take-off,where necessary. As expected, the lower noise levels require higher bypass ratios, and result in increasing gross weight for a given range and mission. Figures9and 10 reproduced from A very significant feature of these curves is how the lower bound tends to become vertical for a given range and level that cannot be bettered at a given level of technology. payload. Thus there is a minimum noise \

ROLLS-ROYCe LyMPUS total air flow as turbofan B since both ERGN MACH-2.O DERNATIVE ENGNES engines would have almost the same weight x/O of intake,nozzles etc and drag and,poten- tially the same take-off noise. Figure 11 shows thrust and sfc of engine D compared with engine B as datum.Both engines have nearly the same thrust and sfc.If mixing is deleted from engine D the thrust will fall and sfc rise by some 210 g 4 as shown.Engine D,unmixed,is now scaled down to the same flow as engine B, and,as shown,the thrust falls by the same amount (at constant sfc).Duct burning is now added to the scaled down engine D,to restore the thrust,If the duct burner MACH 2.0 SA+5℃ STRATOSPHERE Fig 9.Gross Weight Community Noise,Fixed Mission Range/Variations in Payload 2驰 8 SFC RELATIVE e TO ENGINE ' DESISN MACH.2O 绿 XICOO o SDEKT BHRNINS EFFICIENCY Z· O5RBAD【MwhM2s -BHRN TO RESTORE THRLEST FOR WEIGHT 1地 9 4500 NM 饼 00%DIET BHRNING EPFICIENCY 3 SL正7 SAME FLOW AS P上2xW6 u2e@创 3a THRUST RELATIVE TO ENGINE 'B' r Fig 1l. Comparison of Duct Burning Turbofan with Dry Mixed Turbofan at Same SMMATED NOISE IO EPNdg ITERVALS Total Airflow efficiency,including profile losses,is 100%the sfc will fall slightly and rise Fig 10. Gross Weight Community Noise again,When the thrust of engine B is Fixed Payload/Variations in Mission reached it will be seen that the sfc is 8 Range above engine B.If a typical current duct burner efficiency of 85%is achieved the Reference 5 also shows a linear vari- sfc at datum thrust is shown to be nearly ation of DOC and TOC with design gross 16%above the engine B datum. weight,for a given range and payload. Cost estimation is the subject of a further The 2/3 scaled engine D with duct burn- ICAO study,and this subject will not be ing and the dry mixed turbofan B may,at pursued further in this paper,except to the same thrust level,be scaled up or down note that the viability of the aircraft is equally to meet a given thrust level. reflected by the take-off gross weight for a given mission. The same aircraft model used to compare the dry engine family,may be used to com- Included in Reference 5 is a list of the pare the duct burning engine with the dry sensitivities of gross weight,noise and turbofan.Table 5 compares the duct burn- costs to changes in component assumptions, ing engine sfc at the four flight conditions from which the risks involved in not meet- used for the mission fuel evaluation with ing design targets can be assessed. the dry turbofan for 100%and 85%duct burner efficiency,It will be noted that Duct Burning whilst the supersonic cruise and transonic acceleration sfc are higher for the duct The studies described so far have burning engine,the subsonic diversion and assumed no afterburning or duct burning at hold sfc are lower,due to the smaller core cruise,Duct burning will now be examined size and higher bypass ratio. as applied to turbofan D whose take-off bypass'ratio is 1.31,and a comparison will Table 6 gives the breakdown of the power- be made with turbofan B (take-off bypass plant plus fuel weight comparison. The ratio 0,46).The most instructive com- duct burning engine at 100%burner parison is between turbofan B and turbo- efficiency is shown to have a powerplant fan D scaled down (by 2/3)to give the same plus fuel weight penalty of only 1.1%TOGW >

.total air flow as turbofan B since both engines would have almost the same weight of intake, nozzles etc and drag and, poten￾tially the same take-off noise. Figure I1 shows thrust and sfc of engine D compared with engine B as datum. Both engines have nearly the same thrust and sfc. If mixing is deleted from engine D the thrust will fall and sfc rise by some 4% as shown. Engine D, unmixed, is now scaled down to the same flow as engine B, and, as shown, the thrust falls by the Same amount (at constant sfc). Duct burning is now added to the scaleddownengine D, to restore the thrust, If the duct burner Fig 9. Gross Weight - Community Noise,Fixed Mission Range/Variations in Payload SUMMATED NUS5 IO tP@,dB A'rEWAL5 Fig 10. Gross Weight - Community Noise * Fixed Payload/Variations in Mission Range Reference 5 also shows a linear vari￾ation of DOC and TCC with design gross weight, for a given range and payload. Cost estimation is the subj6ct of a further ICAO study, and this subject will not be pursued further in this paper, except to note that the viability of the aircraft is reflected by the take-off gross weight for a given mission. sensitivities of gross weight, noise and costs to changes in component assumptions, from which the risks involved in not meet￾ing design targets can be assessed. Included in Reference 5 is a list of the Duct Burning * The studies described so far have assumed no afterburning or duct burning at cruise. Duct burning will now be examined as applied to turbofan D whose take-off bypass'ratio is 1.31, and a comparison will be made with turbofan B (take-off bypass ratio 0.46). The most instructive com- parison is between turbofan B and turbo￾fan D scaled down (by 2/3) to give the same Fig 11. Comparison of Duct Burning Turbofan with Dry Mixed Turbofan at Same Total Airflow efficiency, including profile losses, is 100% the sfc will fall slightly and rise again. When the thrust of engine B is reached it will be seen that the sfc is 8% above engine B. If a typical current duct burner efficiency of 85% is achieved the sfc at datum thrust is shown to be nearly 16% above the engine B datum. The 2/3 scaled engine D with duct burn￾ing and the dry mixed turbofan B may, at the same thrust leve1,be scaled up or down equally to meet a given thrust level. The same aircraft model used to compare the dry engine family, may be used to com￾pare the duct burning engine with the dry turbofan. Table 5 compares the duct burn￾ing engine sfc at the four flight conditions used for the mission fuel evaluation with the dry turbofan for 100% and 85% duct burner efficiency. It will be noted that whilst the 'supersonic cruise and transonic acceleration sfc are higher for the'duct burning engine, the subsonic diversion and hold sfc are lower, due to the smaller core size and higher bypass ratio. Table 6 gives the breakdown of the power￾plant plus fuel weight comparison. The duct burning engine at 100% burner .' efficiency is shown to have a powerplant plus fuel weight penalty of~only 1.1% TCGW 7

Class III (beyond 1985)of the ICAO Working SFC ABOVE ENGINE B Group technology classification. FLIGHT CONDITION FOR DUCT BURNER EFFICIENCY Figure 12 shows the effect of changing the scale factor of the duct burning version Condition* Mach No 100% 85% of engine D on the powerplant plus fuel weight and payload comparison with engine B. 1 2.0 8.0 15.6 with zero duct burning the engine must be scaled up slightly to meet the required p 1.2 21.6 31.6 thrust level,due to the loss of thrust from deletion of mixing.There is a significant 3 0.93 -3.0 -3.0 payload penalty arising from the higher weight of the considerably higher total flow 4 0.5 -9.2 -9,2 engine,and from the sfc penalty due to not mixing.As the scale factor is reduced the See Table 4 duct burning must be increased to maintain the thrust,Initially the decrease in Table 5 -Duct Burning Turbofan SFC weight offsets any rise in sfc,and the pay- Comparison With Dry Mixed Turbofan at Same load penalty decreases.As scale is further Total Airflow and Thrust reduced,and duct burning increased,the increase in cruise and transonic sfc becomes significant,the payload penalty reaches a Duct Burner minimum and rises again. Thus there is an Item Efficiency optimum amount of duct burning,and hence 100% 85% engine scale,For 100%burner efficiency it will be seen that the scale to give the same flow as engine B is very near the Internal fuel Wt 19200 37120 optimum,Thus the first comparison at performance equal total flow is virtually at an optimised condition for the duct burning Skin friction fuel Wt 0 0 engine,The very large penalty of reducing duct burner efficiency is again evident. Total fuel 19200 37120 Dressed engine weight -11300 -11300 NCREAS是Pucr B依WNG Intake,nacelle, 0 0 2 ERO DHE可 secondary nozzle weight BURNING Total powerplant weight -11300-11300 r8) EL7 Total powerplant 1b 7900 25820 plus fuel weight %TOGW 1,13% 3.69% Payload penalty,datum 5 16,1% 52,7% payload Figures tabulated are weight changes from datum dry mixed turbofan B (1b wt or Table 6-Powerplant Plus Mission Fuel Weight Comparison,Duct Burning vs Dry SCALE FACTOR ON THRBOFAN 'D 6 Mixed Turbofan at Same Total Airflow and Thrust Fig 12. Comparison of Unmixed Duct Burning (16%datum payload)compared with the dry Turbofan (D) with Dry Mixed mixed turbofan,since the decrease in Turbofan (B) engine weight due to the smaller core nearly offsets the increase in mission fuel. In view of the fact that the duct burn- However if a burner efficiency of 85%can- ing engine has roughly the same mission not be exceeded the payload penalty is over performance as the dry mixed turbofan of 50%。 Small changes in assumptions such as the same total flow and cruise TET level, a longer subsonic element in the mission it is pertinent to ask twhat is the attrac- and a further optimisation of the duct tion of this cycle?t.The answer lies in burning cycle could change the small penalty the fact that this engine can easily be of the 100%efficient duct burner into a adapted to take advantage of any benefit small advantage compared with the dry mixed there may be from the co-annular silencing turbofan for the same flow and TET but the effect. comparison will remain marginal. Reference 6 claims a reduction of 6dB to It will be seen however that no appre- ciable departure from 100%burner efficiency 8dB in jet noise compared with a single (including profile losses)can be tolerated. stream nozzle with the same specific thrust. This poses a formidable development task, and it is noted that duct burning is in

FLIGHT CONDITION FOR DUCT BURNER EFFICIENCY - Condition* Mach No 100% 8 5% 1 2.0 8.0 1.5.6 2 1.2 21.6 31.6 3 0.93 -3.0 -3.0 4 0.5 -9.2 -9.2 ~ ~~~ .~ ~. ~ ~~~~ ~ ~~ ..~.. * See Table 4 I I Duct Burner I Item Efficiency I 100% 8 5% fuel Wt I 19 200 37 120 Internal performance Total fuel -11 300 -11 300 .~~. Intake, nacelle, secondary nozzle weight Total powerplant Payload penalty, % datum 52.7% payload Class I11 (beyond 1985) of the ICAO Working Group technology classification. Figure 12 shows the effect of changing the scale factor of the duct burning version weight and payload comparison with engine E. With zero duct burning the engine must be scaled up slightly to meet the required thrust level, due to the loss of thrust from deletion of mixing. There is a significant payload penalty arising from the higher weight of the considerably higher total flow engine, and from the sfc penalty due to not mixing. As the scale factor is reduced the duct burning must be increased to maintain the thrust. Initially the decrease in weight offsets any rise in Sfc, and the pay￾load penalty decreases. As scale is further reduced, and duct burning increased, the increase in cruise and transonic sic becomes significant, the payload penalty reaches a minimum and rises again. Thus there is an optimum amount of duct burning, and hence engine scale. For 100% burner efficiency it will be seen that the scale to give the same flow as engine B is very near the optimum. Thus the first comparison at equal total flow is virtually at an optimised condition for the duct burning engine. The very large penalty of reducing duct burner efficiency is again evident. of engine D on the powerplant plus fuel L Figures tabulated are wezght changes from datum dry mixed turbofan B (lb wt or %) Table 6 - Powerplant Plus Mission Fuel Weiqht Comparison.Duct Burninq vs Dry Mixed Turbofan at Same Total Airflow and Thrust (16% datum payload) compared with the dry mixed turbofan, since the decrease in engine weight due to the smaller core nearly offsets the increase in mission fuel. However if a burner efficiency of 85% can￾not be exceeded the payload penalty is over 50%. Small changes in assumptions such as a longer subsonic element in the mission and a further optimisation of the duct burning cycle could change the small penalty of the 100% efficient duct burner into a small advantage compared with the dry mixed turbofan for the same flow and TET but the comparison will remain marginal. In view of the fact that the duct burn￾ing engine has roughly the same mission performance as the dry mixed turbofan of the same total flow and cruise TET level, it is pertinent to ask ‘what is the attrac￾tion of this cycle71. The answer lies in the fact that this engine can easily be adapted to take advantage of any benefit there may be from the co-annular silencing effect. Reference 6 claims a reduction of 6dB to L ~~ ~ It will be Seen however that no apPre￾ciable departure from 100% burner efficiency Iincludina Drofile losses) can be tolerated. 8dB in jet noise compared with a single stream nozzle with the same specific thrust. .-- -. This poses a formidable development task, and it is noted that duct burning is in 8

By reducing the TET at take-off and Reductions in peak noise of up to 4 PNdB increasing duct burner temperature the have been obtained,but due to the limited desired inverted velocity profile is read- angular extent of the benefit this is worth ily obtained,although the cut-back thrust less than 2 EPNdB in flight.! condition at flyover may pose some difficulties. It has been suggested that,if Rolls- Royce had tested to higher jet velocities The co-annular silencing effect is than reported in Reference 1,higher values discussed in the next section,in the light of co-annular attenuation would have been of European work in this field. achieved,In this connection it is important to note that the critical case Noise Tests on Co-annular Jets for noise reduction on a second generation aircraft is at cut-back (as on Concorde). Measurements have been made by Rolls- The Rolls-Royce tests have covered the cut- Royce and SNECMA of the noise of inverted back jet velocity case where on Concorde velocity profile co-annular jets statically the noise level is around 119 EPNdB (at at model scale. These tests were fully certification measuring condition ) reported last Autumn in Reference 7,which Therefore a 2 EPNdB reduction would in no also discusses the application of the con- way be a break-through,on any solution to cept and compares its noise with that of the objective of achieving 108 EPNdB. other exhaust systems. Furthermore,with a larger aircraft having a bigger engine maximum flow the datum Some preliminary results of these tests noise level would be higher than 119 EPNdB. were given in Reference 1,which emphasises requiring an even bigger attenuation,which the importance of the basis of comparison, current co-annular data could not achieve. ie whether the noise reduction is quoted relative to a fully mixed jet,or relative The European tests have not covered the to the synthesised noise of two separate use of a central plug with co-annular round jets.The latter comparison always streams,which might give greater shows a higher noise reduction for a given attenuation. It is clearly necessary for set of measurements,since the higher the effect of forward speed to be experi- velocity jet,when treated separately,will mentally assessed,including testing at give a higher datum noise level than when real engine scale,before the co-annular fully mixed flows are used. Whilst the inverted velocity profile noise attenuation comparison with synthesised flows has can be used with confidence in a serious certain advantages for plotting test design results,the comparison with fully mixed flows is the valid one when comparing the Mechanical Silencing co-annular jet with other exhaust systems. The encouraging test results of the The Rolls-Royce tests were carried out McDonnell Douglas mechanical suppressor, on the open air noise test site at Derby, both statically and with forward speed on and the rig had the facility to heat both the Rolls-Royce spin rig,reported in streams independently, The SNECMA tests Reference 8,are supported by similar were carried out in the.anechoic chamber of results on Rolls-Royce ejector suppressor the Centre dEssais des Propulseurs (CEPr) models aimed primarily at subsonic near Paris,where only one stream can be applications.It would appear that the heated,Between them the tests covered a ejector/suppressor offers means whereby large range of velocity and temperature some 10dB reduction in the noise of a ratios,and the results showed good agree- supersonic transport can be achieved over ment with published Pratt and Whitney.data, 3 range of jet velocities corresponding showing a maximum reduction of 8kdB peak both to flyover and to the take-off (side- PNL compared with synthesis.The following line)condition. conclusions were reported. If the promise can be realised a very significant impact is made both on engine tIt has been found that an annular jet selection and the viability of a super- has both jet mixing and shock cell noise sonic cruise aircraft.Turning back to lower than a round jet of equal area and Figure.7,it will be noted that a point is that this noise benefit is increased by marked U on the datum thrust line at a adding flow to the centre of the annular flow relative to datum of 2.5.This point jet,For engine cycles where the velocity represents an engine configuration that difference between the stream is accompan- meets a certain stipulated noise level. If ied by a large temperature difference (eg it is now assumed that the noise level of a conventional turbofan with the flows an unsuppressed engine can be reduced by inverted)it has been found that the inver- ted-profile jet always produced less noise 10dB by fitting a suppressor,and that the accompanying net thrust loss is 6% than the conventional non-inverted jet,but (requiring an un-suppressed thrust 6%above rarely less noise than a single stream jet datum),the engine configuration,with such of the same thrust and mass flow. a suppressor that meets the same stipulatec noise level, is represented by the point wnere both streams are heated (eg a duct marked S,which has a flow relative to burning engine)there does appear to be a datum of about 1.65,a reduction in flow of noise benefit due to the inverted flow pro- 34%from point U.Figure 4 shows that a file compared to the fully mixed flow jet. reduction in relative airflow from 2.5 to 9

By reducing the TET at take-off and increasing duct burner temperature the desired inverted velocity profile is read￾ily obtained, although the cut-back thrust condition at,flyover may pose some - difficulties. The co-annular silencing effect is discussed in the next section, in the light of European work in this field. Noise Tests on Co-annular Jets Measurements have been made by Rolls￾Royce and SNECMA of the noise of inverted velocity profile co-annular jets statically at model scale. These tests were fully reported last Autumn in Reference 7, which also discusses the application of the con￾cept and compares its noise with that of other exhaust systems. Some preliminary results of these tests were given in Reference 1, which emphasises the importance of the basis of comparison, ie whether the noise reduction is quoted relative to a fully mixed jet, or relative to the synthesised noise of two separate round jets. The latter comparison always shows a higher noise reduction for a given set of measurements, since the higher velocity jet, when treated separately, will give a higher 'datum'noise level than when fully mixed flows are used. Whilst the comparison with synthesised flows has certain advantages for plotting test results, the comparison with fully mixed flows is the valid one when comparing the co-annular jet with other exhaust systems. The Rolls-Royce tests were carried out on the open air noise test site at Derby, and the rig had the facility to heat both streams independently. The SNECMA tests were carried out in the.anechoic chamber of the Centre d'Essais des Propulseurs (CEPr) near Paris, where only one stream can be heated. large range of velocity and temperature ratios, and the results showed good agree￾ment with published Pratt and Whitney,data. showing a maximum reduction of 8kdB peak PNL compared with svnthesis. The following conclusions were reported. Between them th? tests covered a 'It has been found that an annular jet has both jet mixing and shock cell noise lower than a round jet of equal area and that this noise benefit is increased by adding flow to the centre of the annular jet. For engine cycles where the velocity difference between the stream is accompan￾ied by a large temperature difference (eg a conventional turbofan with the flows inverted) it has been found that the inver￾ted-profile jet always produced less noise than the conventional non-inverted jet, but rarely less noise than a single stream jet ' 0.f the same thrust and mass flow. 'hnere both Streams are heated (eg a duct burning engine) there does appear to be a noise benefit due to the inverted f.low pro￾file compared to the fully mixed flow jet. 9 Reductionsinpeak noise of up to 4 PNdB have been obtained, but due to the limited angular extent of the benefit this is worth less than 2 EPNdB in flight.' It has been suggested that, if Rolls￾Royce had tested to higher jet velocities than reported in Reference 1, higher values of co-annular attenuation would have been achieved. In this connection it is important to note that the critical case for noise reduction on a second generation aircraft is at cut-back (as on Concorde). The Rolls-Royce tests have covered the cut￾back jet velocity case where on Concorde the noise level is around 119 EPNdB (at certification measuring condition). Therefore a 2 EPNdB reduction would in no way be a break-through, on any solution to the objective of achieving 108 EPNdB. Furthermore, with a larger aircraft having a bigger engine maximum flow the datum noise level would be higher than 119 EPNdB, requiring an even bigger attenuation, which current co-annular data could not achieve. The European tests have not covered the use of a central plug with co-annular streams, which might give greater attenuation. It is clearly necessary for the effect of forward speed to be experi￾mentally assessed, including testing at real engine scale, before the co-annular inverted velocity profile noise attenuation can be used with confidence in a serious design. Mechanical Silencing The encouraging test results of the McDonnell Douglas mechanical suppressor; both statically and with forward speed on the Rolls-Royce spin rig, reported in Reference 8, are supported by similar results on Rolls-Royce ejector suppressor models aimed primarily at subsonic applications. It would appear that the ejector/suppressor offers means whereby some lOdB reduction in the noise of a supersonic transport can be achieved over a range of jet velocities corresponding both to flyover and to the take-off (side￾line) condition. If the promise can be realised a very significant impact is made both on engine selection and the viability of a super￾sonic cruise aircraft. Turning back to Figure.7, it will be noted that a point is marked U on the datum thrust line at a flow relative to datum of 2.5. This point represents an engine configuration that meets a certain stipulated noise level. If it is now assumed that the noise level of an unsuppressed engine can be reduced by lOdB by fifting a suppressor, and that the accompanying net thrust loss is 6% (requiring an un-suppressed thrust 6% above datum), the engine configuration, with such a suppressor that meets the same stipulate? noise level, is represented by the point marked S, which has a flow relative to datum of about 1.65, a reduction in flow of 34% from point U. Figure 4 shows that a reduction in relative airflow from 2.5 to