Turbojet and turbofan engines for a mach 3 supersonic transport

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TURBOJET AND TURBOFAN ENGINES FOR A MACH 3 SUPERSONIC TRANSPORT by JAMES F. DUGAN, JR., ROBERT W. KOENIG, JOHN B. WHITLOW, JR. and TIMOTHY B. McAULIFFE National Aeronautics and Space Administration Cleveland, Ohio AIAA Paper No. 64-244

1st AIAA Annual Meeting

Washington,D.C. June 29 —July 2,1964

First publication rights reserved by American Institute of Aeronautics and Astronautics, 1290 Sixth Avenue, New York, N. V. 10019.

$1.00).

(Price—AIAA Member 50c, Non-Member Abstracts may be published without permission if credit is given to author and to AIAA.

TURBOJET AND TURBOFAN ENGINES FOR A MACE 3 SUPERSONIC TRANSPORT by James F. Dugan, Jr., Robert W. Koenig, John B. , Whitlow, Jr., Timothy B. McAuliffe Lewis Research Center National Aeronautics and Space Administration Cleveland, Ohio INTRODUCTION craft takeoff gross weight, which is considered to be the figure of merit in evaluating the different Two programs now exist with the aim of making engines.

commercial supersonic travel a reality - one in In calculating the design and off-design per- Europe and one in the United States. The subject of propulsion for such supersonic aircraft has been a formance of each engine type, the engine components popular one and with good reason. The economics of are matched so as to satisfy the relations involv- ing continuity of flow, engine rotational speed, the aircraft are strongly tied to the aircraft gross weight, and 60 percent or more of the aircraft gross and power balance between the compressor or fan and weight is comprised of the propulsion system and its its driving turbine. In the case of the afterburn- fuel. References 1 to 13 are among the many papers ing. turbofan engine, the fan air flow and turbine gas flow mix at equal static pressures. The proce- that have been written during the past several years on the subject of propulsion for supersonic trans- dures employed are similar to those discussed in ports. In these reports a variety of topics are reference 14.

discussed such as (1) why existing military engines should not be used, (2) the effects of engine de- Engine weight is calculated from empirical sign variables on engine perfthrnance, (3) why one curves that relate installed engine weight to type engine type is to be preferred over another, and of engine and design values of engine airflow, (4) numerous specific problem areas like the impor- overall compressor pressure ratio, fan pressure ra- tio, bypass ratio, and turbine inlet temperature.

tance of matching engine flow to inlet flow.

21000 For a turbine inlet temperature of F and The objective of this paper is to survey four near-optimum values of the remaining design param- types of gas turbine engines that are deemed suit- eters, the engine thrust-to-weight ratio ranges able for powering a Mach 3 supersonic transport: from 4.7 to 5.8 depending on the engine type. For dry turbojets, afterburning turbojets, duct-burning comparison, General Electric's YJ93 afterburning turbojet engine, which will power the Mach 3 RS-70 nonmixed-flow turbofans, and afterburning mixed- flow turbofans. Desirable features of each engine aircraft for the United States Air Force, has a 5)5 type are evolved, and the level of engine technol- thrust-to-weight ratio of above In the term ogy required to result in an attractive ratio of engine thrust-to-weight ratio, engine thrust is payload to aircraft gross weight is estimated. maximum thrust at sea-level static (SLS) condition Other topics considered in the paper are engine siz- and engine weight includes the gas generator, ing criteria, the effect of sonic boom limits on thrust reverser, and exhaust nozzle.

engine size and aircraft gross weight, inlet and ex- haust nozzle performance, airflow scheduling in tur- Aircraft takeoff gross weight (TOGW) is the sum of operating weight empty, fuel weight includ- bofan engines, and the use of variable turbine sta- tors in turbojet engines. ing reserves, and the fixed payload of 26,000 pounds. Operating weight empty less installed en- gine weight (expressed as percent of OGW) was as- NETHOD OF ANALYSIS sumed to vary with 10GW from 36 percent at a TOGW of 308,000 pounds to 29 percent at a TOGW of The four types of gas turbine engines consid- 540,000 pounds.

ered are shown in Fig. 1. The dry turbojet and the A typical variation in maximum lift-to-drag afterburning turbojet are one-spool engines, while the duct-burning turbofan and the afterburning tur- ratio with flight Mach number for the fixed-wing bofan are two-spool engines. Except where noted in aircraft used in this study is shown in Fig. 2.

Also shown in Fig. 2 is the mission profile used in the study, the duct-burning turbofan engine has a fixed primary nozzle downstream of the turbine and the study. The flight path up to the initial Mach 3 cruise altitude is fixed by scheduling a variable primary nozzle downstream of the duct flight Mach number with altitude. Some points burner. In the afterburning turbofan, the air flow from the fan is mixed with the gas flow from the along the path are Mach 1.05 and 40,000 feet, Mach 1.4 and 50,000 feet, and Mach 3 and 60,000 feet.

turbine prior to augmentation of the total flow in The initial cruise altitude is greater than 60,000 the afterburner. The geometry of the mixer is con- feet and is selected in bach case so as to minimize sidered to be fixed.

aircraft TOGW.

Engine performance and weight are calculated for many engines with each characterized by spe- The Mach 3 cruise portion of the flight ends when a range of 2800 nautical miles is achieved.

cific des.gn values of turbine inlet temperature (18000 to 25000 F), overall compressor pressure ra- During letdown over the remaining 400 nautical tio (6 to 13), augmentation temperature (2740 0 to miles, thrust setting and fuel consumption are at a 3100 0 F), fan pressure ratio (2 to 3), and bypass low level. Fuel reserves are sufficient to allow for(l) an extension of supersonic cruise for a ratio (0.5 to 1.5). A fixed-wing aircraft powered by four podded engines and carrying a payload of period equal to 10 percent of the elapsed time from 26,000 pounds is flown over a specified mission with takeoff to end of cruise, (2) cruise at Mach 0.9 a 3200-nautical-mile range. On each flight the en- for a distance of 250 nautical miles at the best 'altitude between 36,000 and 45,000 feet, and gines are sized to result in a minimum value of air-

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quired. The corresponding value of SLS thrust to (3) holding at 1500 feet and 180 knots for 1/2 hour engine weight ratio is 9.3.

RESULTS AND DISCUSSION Afterburning turbojets. - In Fig. 4(a) is shown the effect of design compressor pressure ratio on the ¶10GW of aircraft powered by afterburning turbo- Effect of Design Variables jet engines with a maximum augmentation temperature 31000 F.

of Again, a design compressor pressure ra- Each of the four engine types. is considered tio of about 10 results in a minimum value of air- separately to determine the effects of the major engine design parameters on aircraft TOGW. In the craft ¶10GW.

figures of this section, aircraft TOGW is expressed Based on the input of this study, For engines with a near-optimum design compres- as relative TOGW.

sor pressure ratio of 10, a design turbine inlet a relative TOGW of 100 corresponds to a payload 23000 F results in minimum air- temperature of about ¶10GW of 80 cor- fraction of percent. A relative craft gross weight (Fig. 4(b)). As turbine inlet 19000 responds to a payload of slightly more than 8 per- temperature increases from to 25000 F, the cent and is adopted as a reasonable goal. In striv- amount of fuel decreases continuously. Engine ing for the relative TOGW of 80, all the burden is weight, however, decreases, reaches a minimum around 2300 0 F, and then increases. The two factors that placed on the engine; that is, no improvements in aircraft weight or aerodynamics are hypothesized. influence engine weight are the turbine inlet tem- perature and design airflow. Engine weight per unit Dry turbojets. - The principal design variables design airflow increases continuously with rising of a dry turbojet engine are considered to be design turbine inJt temperature. Design airflow, however, compressor pressure ratio, design turbine inlet tem- decreases and then changes only slightly above about perature, and engine weight. 2300 0 F with further increases in turbine inlet tem- perature. The combined effects of temperature and The effect of design compressor pressure ratio engine size at turbine inlet temperatures above 2300 0 F result in an increase in engine weight. The is shown in Fig. 3(a). Design compressor pressure ¶10GW chiefly through its increase in engine weight overrides the decrease in ratio influences aircraft effect on engine weight and cruise specific fuel fuel so that the net effect is an increase in TOGW.

consumption (SFC). As design compressor pressure ratio increases from 6 to 13, engine weight in- At a turbine inlet temperature of 2300 0 F, the creases continuously. Depending on the level of assumed 6.6-percent cooling air for the turbine ac- turbine inlet temperature, cruise SFC decreases con- counts for an increase in TOGW of about 4 percent.

tinuously or else decreases, reaches a minimum, and The increase would be even more if the use of tur- then increases. In both cases there is a value of bine cooling air resultsin less efficient turbine design compressor pressure ratio that minimizes air- operation. An engine was considered in which the craft TOGW. For any level of turbine inlet temper- use of 6.6 percent of the compressor air to cool the ature in the range 1900 0 to 2500 0 F, a design pres- turbine caused the turbine to operate at 81-percent sure ratio of about 10 results in the minimum value efficiency. Data for this engine and an engine of ¶10GW. whose turbine operates at 88 percent efficiency with 6.6-pecent turbine cooling air are presented in The desirability of being able to build reli- Table I. The drop in turbine efficiency degraded able engines that operate at high values of turbine both thrust and specific fuel consumption throughout inlet temperature is illustrated in Fig. 3(b). As the flight range. The 7-percent drop in turbine ef- 19000 to turbine inlet temperature increases from ficiency caused ¶10GW to increase 9 percent. Engine 25000 F, both cruise SFC and engine weight decrease and fuel weights are tabulated as percentages of continuously. The reduction in engine weight comes ¶10GW. Engine weight remains the same, but total about from the decrease in engine size that results fuel weight increases. The breakdown in total fuel from the higher values of thrust per unit airflow. indicates increases in both useful fuel and reserve 25000 F it At 19000 F relative ¶10GW is 150, while at fuel. Most of the increase in reserve fuel is due has dropped to 115. Each of the dry turbojets had a to poor part-power performance during the 30-minute near-optimum design compressor pressure ratio of 10. hold where specific fuel consumption increased about The solid line corresponds to a selected turbine 11 percent. The increase in useful fuel is due cooling airflow schedule (2.6 percent of compressor mainly to higher fuel consumption prior to cruise.

airflow at 1900 F and 8.6 percent at 2500 F). The Specific fuel consumption is higher and the time re- dashed line is for zero turbine cooling airflow and quired to climb and accelerate to the initial cruise is of academic interest only since materials suit- conditions is longer because of the lower thrust able for operation in this temperature range without level. The increase in TOGW that accompanied the cooling are not presently available. Notice that drop in turbine efficiency illustrates that in order the required cooling airflow degrades the perform- to realize the potential gains of high turbine tem- ance potential appreciably. At 2500 0 F, the cooling perature operation, turbine efficiency must not be airflow requirement of 8.6 percent of compressor degraded appreciably by the use of cooling air. Ex- airflow is responsible for an increase in ¶10GW of cept for the case just described, no penalty in tur- about 9 percent. bine efficiency was assessed for high temperature turbine operation.

Figure 3(c) shows the reduction in aircraft TOGW that would result from being able to build The effect of engine weight reduction on after- lighter dry turbojets. The results are for a dry burning turbojets is shown in Fig. 4(c). An engine turbojet having a near-optimum design compressor weight reduction of 10 percent reduces aircraft ¶10GW pressurd ratio and a maximum turbine inlet tempera- about 4 percent. To attain a relative TOGW of 80, ture of 2500 0 F. Each 10-percent reduction in en- a weight reduction of 48 percent is required. The gine weight results in about a 5-percent reduction corresponding engine thrust-to-weight ratio is 11.3.

in aircraft ¶10GW. To achieve a relative TOGW of 80, Mterburning turbofans. - In Fig. 5(a), after- an engine weight reduction of 55 percent is re- -2-

burning turbofan engines having a design bypass ra- section was sized to minimize aircraft ¶10GW. In tio of 1.0 are considered. For a design turbine sizing an engine for a supersonic transport, how- 18000 F. a design overall com- inlet temperature of ever, it is possible that design engine airflow will pressor pressure ratio of about U results in mini- be dictated by something such as takeoff distance mum aircraft ¶10GW. At 2400 0 F, the optimum pres- or noise. These and other criteria that could be sure ratio rises to about 13; however, at a pres- critical in engine sizing are considered for the sure ratio of U, ¶10GW is only 1/2 percent greater case of a specific afterburning turbofan (Fig. 7).

than minimum.

From Fig. 7(a), the design engine airflow that mini- mizes aircraft ¶10GW is 460 pounds per second. If In Fig. 5(b), design overall compressor pres- some other criterion requires a design engine air- sure ratio is fixed at 11. Aircraft gross weight flow larger than 460 pounds per second, some in- is seen to be relatively insensitive to design by- crease in aircraft ¶10GW must be accepted.

pass ratio.

In Figs. 7(b), the noise level 1000 feet from The decrease in aircraft gross weight that re- the runway is plotted against design engine airflow sults from operating afterburning turbofans at with afterburner setting as a parameter. The pro- higher turbine inlet temperatures is shown in Fig.

cedures of Ref. 16 were used to calculate engine 5(c). Each engine considered has a design overall noise. At Los Angeles International Airport, noise compressor pressure ratio of 11 and a design bypass 1400 feet from the runway is limited to 120 per- ratio of 1.0. Raising turbine inlet temperature ceived noise decibels (PNdb); at 1000 feet, the limit from 18000 to 24000 F reduced aircraft TOGW about would be about 123 PNdb. To satisfy this limit on 7 percent.

a hot day, the afterburner temperature during take- off should not exceed about 1940 F. The value of Engine weight reduction is considered in Fig.

design engine airflow has only a slight effect on 5(d) for an afterburning turbofan having the fol- runway noise. At the 3-nautical-mile point, how- lowing design values: overall compressor pressure ever, it has an appreciable effect (Fig. 7(c)).

ratio, 11; bypass ratio, 1; turbine inlet tempera- This is because larger engines enable the aircraft ture, 2400 F; and afterburner temperature, 27400F.

to reach a higher altitude at the 3-nautical-mile An engine weight reduction of 10 percent results in point. On a hot day with the afterburner tempera- ture set at 1940 0 F, the noise level at the 3- about a 3k-percent reduction in aircraft TOGW. A nautical-mile point can be limited to a value of relative ¶10GW of 80 requires an engine weight re- 112 FNdb by installing engines with design airflows duction of about 40 percent. Such an engine would of 504 pounds per second. The lift-off distance on have a SLS thrust to engine weight ratio of 10.3.

a hot day for this engine size and afterburner setting is 4300 feet or 45 percent of a 9500 foot Duct-burning turbofans. - In Fig. 6 the ef- balanced field length (Fig. 7(d)) and the lift-off fects that the major design variables of duct- speed is 165 knots. These values of lift-off speed burning turbofans have on aircraft ¶10GW are shown.

and distance are considered to be acceptable.

All engines in Fig. 6(a) have a design bypass ratio of 1, a design fan pressure ratio of 2.5, and a 'The engine sizing considerations can be re- design duct-burner temperature of 3100 0 F. The op- peated for standard-day operation. To limit runway timum value of overall compressor pressure ratio is noise to 120 PNdb, afterburner temperature should seen to increase from a value of about 8 at a de- not exceed about 1340 F (Fig. 7(b)). With this sign turbine inlet temperature of 1800 0 F to a power setting, noise at the 3-nautical-mile point value of about 11 at a turbine inlet temperature of can be limited to 112 PNdb by selecting design en- 24000 F.

gine airflow per engine to be 486 pounds per second (Fig. 7(c)). Lift-off speed would be 165 knots and The engines of Fig. 6(b) have a design overall lift-off distance 4400 feet (Fig. 7(d)). Standard- pressure ratio of 10 and a design bypass ratio of day operation is less critical than hot-day opera- 1. The optimum value of fan pressure ratio is tion, so the hot-day engine size of 504 pounds per about 2.5 for turbine inlet temperatures in the second is selected to test two other possible siz- range 1800 0 to 24000 F. ing criteria.

The engines of Fig. 6(c) have a design fan In Fig. 7(e), the climb path angle after an pressure ratio of 2.5 and a design overall pressure engine failure is plotted against design engine air- ratio of 10. The optimum design bypass ratio is flow per engine. The Civil Air Regulation lower about 1.2 for turbine inlet temperatures of 18000 limit on second-segment climb gradient is 0.03, and 2100 0 F. For a turbine inlet temperature of which corresponds to a climb path angle of 1.720.

24000 F, optimum design bypass ratio is about 1.

For a design engine airflow of 504 pounds per sec- ond, the climb path angle is 3.95 0 . Thus, the ten- The effect of design turbine Inlet temperature tative engine size satisfies the one-engine out- is shown in Fig. 6(d). Airplane weight drops climb requirement. There is no regulation govern- 20 percent as turbine inlet temperature is raised ing minimum transonic thrust margin, but it has from 1800 0 to 24000 F.

been suggested that the minimum thrust margin should be 0.3 on a standard day in order that ade- Figure 6(e) shows that for the specific duct- quate thrust be available for hot-day acceleration burning turbofan considered, a relative aircraft to cruise speed in a reasonable time. From Fig.

¶10GW of 80 requires an engine weight reduction of 7(f), minimum transonic thrust margin on a standard about 35 percent. Such an engine would have a day is 0.46 for a design engine airflow of 504 thrust-to-weight ratio of 9.6.

pounds per second. For this particular afterburn- ing turbofan engine, then, the critical engine siz- ing criterion was that noise level at the 3- Engine Sizing Considerations nautical-mile point on a hot day should not exceed 112 PNdb. This required an engine about 10 percent Each of the engines discussed in the previous larger than that for minimum TOGW, and the result- -3-

ant penalty in 10GW was about 1/2 percent (Fig. the higher thrust levels attained with the higher 7(a)). recovery inlet.

The sizable effect that exhaust nozzle perform- Effect of Sonic Boom Limits ance can have an aircraft gross weight is shown in Fig. 10. The schedule of gross thrust coefficient The calculations presented thus far are for a with flight Mach number is representative of a high fixed schedule of altitude with flight Mach number performance exhaust nozzle. If the gross thrust during climb up to initial Mach 3 cruise conditions.

coefficient at each flight condition could be in- As a result, the maximum sonic boom associated with creased 0.01, aircraft gross weight could be re- climb varied for each aircraft/engine combination.

duced an impressive 5 percent. While such a gain For a ¶10GW of 400,000 pounds, maximum sonic boom is indeed enticing, this extreme sensitivity of air- during climb and acceleration to initial cruise craft gross weight to exhaust nozzle performance is conditions was found to be about 2 pounds per a reminder of the serious consequences that would square foot (PSF). For a ¶10GW of 320,000 pounds, result from falling short in developing a high per- the maximum sonic boom was about 1.8 PSF. During formance exhaust nozzle. For the case considered Mach 3 cruise, the maximum sonic boom occurs at the here, if the exhaust nozzle gross thrust coefficient beginning of cruise. It ranged from 1.5 to 1.7 PSF is decreased by 0.01 at each flight condition, the for aircraft having a ¶10GW of 400,000 pounds, and effect on aircraft ¶10GW is an increase of 5 per- from 1.4 to 1.6 PSF for aircraft having a ¶10GW of cent.

320,000 pounds. These values of sonic boom are 20 to 28 percent greater than the values that would be obtained for a rubberized aircraft configuration Airflow Scheduling in Turbofan Engines having an optimum combination of lift and volume at each flight condition.17 One of the choices open to the turbofan engine designer is the location of the engine operating If a specific engine is selected and the climb line on the fan performance map. In Fig. 11, the path is varied from its nominal schedule of alti- fan performance map of a duct-burning turbofan en- tude with Mach number, engine size, aircraft ¶10GW, gine is shown with three arbitrarily selected engine and maximum climb sonic boom, all will vary. The operating lines - A, B, and C. The variation in results of such a calculation for a specific duct- aircraft ¶10GW is seen to be less than 2 percent.

burning turbofan engine are shown in Fig. 8.

Operating line A resulted in the minimum value of ¶10GW.

In generating engine performance for the many In Fig. 8(a), design engine airflow is plotted duct-burning turbofan engines of this study, no at- against maximum sonic boom overpressure. For a tempt was made to select the optimum fan operating sonic boom of 2.2 PSF, design engine airflow is line for each engine. Instead, a fan operating line 440 pounds per second. For a lower sonic boom, the was drawn for each engine so that it resembled oper- altitude flown by the aircraft must be raised, and ating line A in Fig. 11.

this demands a larger engine. For a maximum sonic boom of 2.0 PSF, design engine airflow has in- Similar considerations were given to the choice creased to 550 pounds per second. As engine size of engine operating line on the fan performance map and weight increase, the aircraft TOGW also in- of the afterburning turbofan engine.

creases (Fig. 8(b)). This increase in aircraft weight tends to increase sonic boom. Thus, in- The duct-burning turbofan engines of the study stalling bigger engines to fly at higher altitudes have a fixed primary nozzle downstream of the tur- becomes less and less effective in reducing maxi- bine. As a result, engine airflow is not well mum sonic boom overpressure. Figure 8(b) shows matched with inlet airflow and spillage drag is that the sonic boom of this particular aircraft quite high during operation in the transonic speed cannot be lowered beyond about 1.95 PSF. This ex- range. A scheme for improving the inlet-engine air- ample illustrates the very major effect that allow- flow match is to incorporate a variable area nozzle able sonic boom overpressure has on the propulsion downstream of the turbine. With such an engine, the system and the 10GW of the aircraft.

engine airflow can be reduced at high flight Mach numbers. This results in better inlet engine air- flow matching throughout the flight speed range.

Inlet and Exhaust Nozzle Performance In Fig. 12, two duct-burning turbofan engines The effect of inlet pressure recovery on the are compared. They are alike in most respects but TOGW of aircraft powered by duct-burning turbofan differ in the type of nozzle downstream of the tur- engines is shown in Fig. 9. Three schedules of in- bine. Engine A has a fixed-area nozzle that results let pressure ratio with flight Mach number are in an engine airflow schedule designated base flow.

shown. In schedule A, which is used as a refer- Engine B has a variable-area nozzle that results in ence, recovery is 95 percent at takeoff and 85 per- an engine airflow schedule designated low flow. Up cent during Mach 3 cruise. In schedule B, recovery to flight Mach numbers of 2.5 the'two engine airflow during Mach 3 cruise was raised to 90 percent. The schedules are identical. Above Mach 2.5 the low- effect of this change on aircraft gross weight is flow engine demands less airflow than the base-flow shown on the right. The decrease in ¶10GW is less engine. At Mach 3, the difference in airflow is than 1 percent. The third recovery schedule is 20 percent.

characterized by a pressure ratio of 1.0 at takeoff and 92 percent during Mach 3 cruise. With such an Since the inlet is sized by the Mach 3 cruise inlet, aircraft gross weight-decreased over 5 per- condition, the low-flow engine has a smaller lighter cent. This was due principally to a decrease in inlet, but more nacelle wave drag. This is shown at the amount of fuel consumed prior to cruise. Al- the left of Fig. 12 where the installation drag co- though specific fuel consumption changed only efficient is plotted against the flight Mach number.

slightly, the time required to . reach cruise condi- Over most of the speed range, however, the low-flow tions decreased appreciably. This resulted from - 4 -

engine has a lower engine drag. This is because the and the turbine efficiency must not be degraded ap- reduction in spillage drag more than compensates foi preciably by the cooling airflow.

the increase in wave drag. The figure on the right The effect of allowable sonic boom overpres- shows that aircraft LOGW is reduced about 3 per- sure on engine size and aircraft '10GW was such that sonic booms below a certain level were not attain- cent by using the low-flow turbofan rather than the able; improvements in the propulsion system and/or base-flow turbofan. This gain must be weighed the aircraft would be required to lower the limiting against the complication of building and control- value of sonic boom overpressure.

ling the engine with a variable nozzle downstream of the turbine.

It is concluded that worthwhile benefits to the concept of supersonic commercial air travel will re- sult if the gas turbine engine technology continues Variable Turbine Stators in Turbojet Engines to improve. The magnitude of the possible gains in- dicates that research efforts in propulsion should For subsonic flight, the turbofan engine cycle be continued and intensified.

gives lower fuel consumption than the turbojet en- gine cycle. This is relevant to the two reserve requirements calling for subsonic flight to an al- REFERENCES ternate airport and holding prior to landing. In most cases, the weight of reserve fuel for these 1. Jamison, R. H., and Lane, R. J.: Engines for requirements equaled or exceeded the weight of the Supersonic Air Liners. Jour. Roy. Aero. Soc payload.

vol. 64, no. 597, Sept. 1960, 507-534.

pp.

One means for improving the fuel consumption 2. Alford, J. S.: Power Plants for Supersonic of the turbojet engine during subsonic flight is to Transports. Jour. Roy. Aero. Soc., vol. 64, incorporate variable turbine stators. The improve- no. 598, Oct. 1960, pp. 617-628.

ment that this can lead to is illustrated in Fig.

13. Engine performance for the hold flight condi- Z. Sens, W. H., and Slaiby, T. G.: A Consideration tion is shown for two turbojet engines: one with of Powerplants for Supersonic Transport Air- fixed turbine stators and the other with variable craft. Paper 341E, SAE, 1961.

turbine stators. The required level of thrust is such that the variable turbine stator engine has a 4. Staff of Bristol Siddeley Engines, Ltd.: Power- 7-percent lower specific fuel consumption. A simi- plants for the Supersonic Transport. Some De- lar advantage prevails during subsonic cruise to an sign Problems. Conf. 14/WP-SY}'/6, Symposium alternate airport. The effect on aircraft gross on Supersonic Air Transport, International Air weight is shown on the right of Fig. 13. By power- Transport Assoc. Tech. Conf., Montreal ing the aircraft with turbojet engines having vari- (Canada), Apr. 1961.

able turbine stators, aircraft gross weight was re- 5. Staff of Rolls-Royce: An Engine Manufacturer's duced about 27 percent. Much larger benefits would Appreciation of the Engine Problems for a Su - result from the use of variable turbine stators if personic Civil Aircraft. Conf. 14/WP-SYMP/7, airline operations required considerable flying Symposium on Supersonic Air Transport, Inter- time at subsonic speeds. Thus, the incentive for national Air Transport Assoc. Tech. Conf., developing the variable turbine stator concept de- Montreal (Canada), Apr. 1961.

pends very much on airline operational requirements.

6. Staff of General Electric Company: Factors In- fluencing the Selection of the Supersonic Transport Powerplant. Conf. 14/WP-SYMP/9, Symposium on Supersonic Air Transport, Inter- A Mach 3 transport aircraft with a fixed pay- national Air Transport Assoc. Tech. Conf., load and powered by various turbojet and turbofan Montreal (Canada), Apr. 1961.

engines was flown on a 3200-nautical-mile mission.

Minimum aircraft takeoff gross weight (TOGW) was 7. Flamand, C.: Turbo-Ramjet Powerplants and the used to indicate desirable values of engine design Supersonic Transport. Conf. 14/WP-S/12, parameters. For both dry and afterburning turbojet Symposium on Supersonic Air Transport, Inter- engines, a design compressor pressure ratio of national Air Transport Assoc. Tech. Conf., about 10 resulted in minimum 10GW. For the after- Montreal (Canada), Apr. 1961.

burning turbofan engine, the optimum overall com- pressor pressure ratio ranged from 11 to 13, while 8. Staff of Pratt & Whitney Aircraft: Power for a for the duct-burning turbofan engine, the range was Supersonic Transport. Conf. 14/WP-SYMP/32, from 8 to 11. Near optimum design values of bypass Symposium on Supersonic Air Transport, Inter- ratio and fan pressure ratio were 1.0 and 2.5, re- national Air Transport Assoc. Tech. Conf., spectively.

Montreal (Canada), Apr. 1961.

Engine weight reduction and high values of 9. Nichols, Mark H.: Supersonic Transport Propul - turbine inlet temperature resulted in lighter air- sion Requirements. Aerospace Eng., vol. 20, craft to carry the fixed payload. With each of the no. 9, pt. 1, Sept. 1961, pp. 12-13; 58-63.

four engine types, a payload equal to about 8 per- cent of the TOGW was attainable. The gas turbine 10. Walker, C. J.: Some Aspects of Supersonic engines had engine thrust-to-weight ratios in the Transport Propulsion. Preprint 427A, SAE, range 9.3 to 11.3 and operated at turbine inlet 1961.

25000 F. To temperatures in the range 2300 0 to realize the potential gains from high turbine tem- 11. Hawkins, H. C.: Powerplant Considerations for perature operation, the turbine must be adequately a Mach 3.0 Commercial Transport. Preprint cooled with only modest amounts of cooling airflow, 427B, SAE, 1961.

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TABLE I. - EFFECT OF TURBINE EFFICIENCY 12. Builder, C. H., and Cuadra, E.: An Elemental Approach to Propulsion Selection for the Su- [Afterburning turbojet design values: personic Transport. Paper 62-72, lAS, 1962.

compressor pressure ratio, 10; tur- bine inlet temperature, 23000 F; 13. Slaiby, T. G., and Stauback, R. L.: Propulsion afterburner temperature, 3100 0 F.

Supersonic Transports. Paper

Systems for '

586A, SAL, 1962.

88 81 Turbine efficiency 14. Dugan, James F., Jr.: Aerodynamic Design of 99.5 108.3 Relative TOGW Axial-Flow Compressors, Vol. 3. Ch. XVII - Weight, percent '10GW Compressor and Turbine Matching. NACA NM 9.56 9.56 Engine E56BO3b, 1956.

51.66 53.13 Total fuel 42.02 43.03 Useful fuel 15. Anon.: General Electric YJ93, Shell Aviation 10.10 9.64 Reserve fuel News, no. 305, Nov. 1963, P. 19.

16. Goodwin, W. M.: Procedure for Estimating Broad Band Noise of Jet Engine Exhaust. PWA-2045, Pratt and Whitney Aircraft, Feb. 13, 1962.

17. Carlson, H. W.: The Lower Bound of Attainable 14 - Sonic-Boom Overpressure and Design Methods of Approaching This Limit. NASA TN D-1494, 1962.

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Figure 1. - Gas turbine engines.

Turbine Turbine cooling air C it Without a 130— e i Design turbine Inlet temperature, OF Design compressor pressure ratio (b) Design turbine Inlet temperature. Design compressor pressure (a) Design compressor pressure ratio. Noturbine cooling air.

ratio, 10.

C .2' a e Engine weight reduction. percent (c) Engine weight. Design values: compressor pressure ratio, 1 turbine inlet temperature. 2500° F.

Figure 3. - Effect of design variables for dry turbojet.

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