Document
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COMBUSTION PERFORMANCE OF TWO EXPERIMENTAL TURBOJET ANNULAR COMBUSTORS A T CONDITIONS SIMULATING HIGH-ALT IT UDE SUPERSONIC FLIGHT By Eugene V. Zettle, C a r l T . Norgren, and H e r m a n Mark L e w i s F l i g h t Propulsion Laboratory C leve land, Ohio
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XEROX $ 2, +L, M I C R O F I L M $ "hb material contains information afiecang t t m National Defense oi the United States within Ur m e w of the esplorrap laws, ntle 18, U.S.C., Secs. 789 aed 794, t t m transmission or revelation of which in any manner t o rn unautbrteed person I s prohibited by law.
NATIONAL ADVISORY COMMITTEE
F O R AERONAUTICS
WASHINGTON March 26,1954 .........................
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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS AT CONDITIONS S - G HIGH-ALTITUDE SUPERSONIC FLIGHT By Eugene V . Zettle, Carl T. Norgren, and Herman Mark SUMMARY The perf o m m e of two experimental annular t u r b o j e t combustors w m i n v e s t i g a t e d a t operating conditions t y p i c a l of h i g h - a l t i t u d e supersonic Each combustor consisted of a o n e q u a r t e r s e c t o r of a s i n g l e an- f l i g h t .
n u l a r combustor designed t o fit i n a housing with an outside diameter of 1 5 255- inches, a n i n s i d e diameter of 1% inches, and a combustor l e n g t h of Liquid f u e l waa i n j e c t e d i n t o t h e combustion approximately 2 3 inches.
a f u e l - chamber from t h e upstream f a c e of t h e combustor; i n addition, s t a g i n g technique WBB investigated.
Combustion efficiencies near 100 percent were achieved i n both ex- perimental cornbustors operating with combustor reference v e l o c i t i e s of 200 f e e t per second and g r e a t e r at simulated supersonic f l i g h t c o n d i t i o m .
These high e f f i c i e n c i e s were maintained at t h e highest combustor-outlet temperatures investigated, namely, 1 8 0 0 ' F f o r one combustor and 2000' F f o r t h e o t h e r . Reasonably f l a t outlet-temperature p r o f i l e s were obtained The maximum t o t a l - p r e e s u r e l o s s e s were and were considered s a t i s f a c t o r y .
10.2 and 1 2 . 6 percent f o r t h e two combustors at a v e l o c i t y of 165 feet per second and a temperature r a t i o of about 1 . 7 . For combustor pressure l o s s e s of t h i s magnitude, c a l c u l a t i o n s indicated t h a t t h e increase i n engine s p e c i f i c - f u e l consumption r e a u l t i n g from combustor pressure l o s s e s would be no g r e a t e r i n t h e engine for supersonic propulsion t h a n i n cur- r e n t t u r b o j e t engines. These pressure losses t h e r e f o r e appear accepta- b l e for t h e supersonic f l i g h t conditions. Combustor-liner d u r a b i l i t y and carbon-deposition c h a r a c t e r i s t i c s of t h e combustors were not evaluated i n t h i s i n v e s t i g a t i o n .
Research on compressor and t u r b i n e aerodynamics has indicated that increases i n a i r flow Der u n i t f r o n t a l area of aa much a a 30 percent are possible for t h e s e comionents of t h e turbojet engine ( r e f s . 1 t o 4 ) .
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I n addition, t h e advancement of turbine-cooling techniques indicates that increases i n operating temperatures of as much 88 500' F over c u r r e n t p r a c t i c e are p o s s i b l e i n f u t u r e engines (refs. 5 and 6 ) . High supersonic f l i g h t speeds w i t h t u r b o j e t engines can be more e m i l y realized w i t h t h e The t u r - g r e a t e r power r e s u l t i n g from higher a i r flows and temperatures.
bojet combustor designed f o r u s e i n an engine incorporating t h e s e d- vancements and powering an aircraft a t high supersonic speeds (Mach num- bers of 2.0 t o 3.0) w i l l be required t o operate w i t h much higher air flows and at higher temperature l e v e l s . This means higher combustor v e l o c i t i e s , i f t h e combustor f r o n t a l area is n o t t o exceed t h a t of t h e o t h e r engine components. The high air flows a l s o indicate higher f u e l flows and higher h e a t - r e l e m e rates.
From t h e r e s u l t s of previous i n v e s t i g a t i o n s ( r e f . 7 ) , increased com- bustor flow v e l o c i t i e s t h a t are desirable i n supersonic propulsion would be expected t o r e s u l t i n decreased combustion e f f i c i e n c y . The higher pressures and combustor-inlet temperatures encountered a t supersonic f l i g h t conditions may, however, tend t o a l l e v i a t e t h e adverse e f f e c t of velocity on combustion e f f i c i e n c y . The increased flow v e l o c i t i e s w i l l a l s o increase combustor pressure loss, which adversely a f f e c t s engine f u e l comumption. The higher combmtor-inlet and - o u t l e t temperature l e v e l s (of the order of 400' t o 500' F above c u r r e n t combustors) w i l l i n c r e a s e t h e d u r a b i l i t y problems involved i n t h e combustor p a r t s .
The preliminary i n v e s t i g a t i o n s t h a t are reported h e r e i n are a p a r t of a general r e s e a r c h program a t t h e NACA Lewis l a b o r a t o r y t o determine design c r i t e r i a of combustors f o r t u r b o j e t engines operating a t high d- t i t u d e s and supersonic f l i g h t speeds. Performance c h a r a c t e r i s t i c s of two experimental single-annulus combustors were obtained a t combustor- i n l e t - a i r conditione approximating t h o s e of an engine with advanced de- sign components operating i n t h e range of a l t i t u d e s from 60,000 t o 80,000 f e e t and of f l i g h t Mach numbers from 2.0 t o 3.0. One-quarter s e c t o r s of t h e combustors were investigated i n a direct-connect system. Pressure- atomized l i q u i d f u e l w a s used i n both combustors. Hollow-cone spray nozzles injected f u e l a x i a l l y from t h e upstream f a c e of t h e cornbustors; i n addition, one combustor was equipped with f l a t spray nozzles i n j e c t i n g r a d i a l l y i n t o t h e combustor f o r t h e purpose of f u e l s t a g i n g a t high flow rates.
The performance of each combustor w a s evaluated a t a single i n l e t - a i r temperature of 870' F, a range of inlet-air pressures from 10 t o 30 , pounds p e r square inch absolute, and a range of combustor v e l o c i t i e s f r b m 125 t o 2 2 5 feet per second. Combustion e f f i c i e n c i e s , pressure l o s s e s , and combustor-outlet-temperature p r o f i l e s were determined a t t h e s e con- d i t i o n s . Combustor-liner d u r a b i l i t y and carbon d e p o s i t i o n were not evalu- ated during t h i s i n v e s t i g a t i o n .
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Combustors Two experimental annular combustors me h e r e i n describedj n e i t h e r of t h e s e combustors n e c e s s a r i l y represents an optimum design. Since t h e air flow per u n i t f r o n t a l a r e a of t h e engine is very high i n a t u r b o j e t engine f o r supersonic application, an annular configuration w a s s e l e c t e d f o r t h e combustors i n order t o maintain as low a f l a w v e l o c i t y 88 possible.
a one-quarter t l t s t o r of a ~i~@e-n_nni-ilas com- Each combustor consisted of bustor designed t o f i t i n t o a housing with a3 outside diameter of 252- in- ches, a n i n s i d e diameter of 1 % inches, and a combustor l e n g t h of approxi- mately 2 3 inches. The maximum combustor c r o s s - s e c t i o n a l area of t h e s e c - t o r w a s 105 square inchea, which corresponds t o 420 square inches f o r t h e complete combustor. I n each of t h e combustors t h e primary air w a s ad- mitted gradually and t h e secondary air, r a p i d l y through l a r g e rectangular s l o t s . Three-quarter cutaway views of the assembled combustors are shown i n f i g u r e 1. The combustor longitudinal c r o s s - s e c t i o n a l and a i r - e n t r y hole geometries a r e shown i n f i g u r e s 2 t o 4 .
A ( f i g . 2 ( a ) ) w a s s i m i l a r t o t h a t The geometric shape of combustor reported i n reference 8 i n t h a t t h e combustor occupied t h e same volume and p o s i t i o n w i t h i n t h e housing. The p r i m a r y zone w a s designed with a s e r i e s of c i r c u l a r holes ( f i g . 3(a)) which allowed primary air t o e n t e r between t h e f u e l nozzles, thus establishing alternate f u e l - and a i r - r i c h zones. Large secondary s l o t s ( f i g . 3 ( a ) ) were used t o provide adequate p e n e t r a t i o n of t h e secondary air and t o minimize flow r e s t r i c t i o n s .
Fuel was introduced through f i v e hollow-cone spray nozzles (10.5 gal/hr; 60' spray angle) located at t h e upstream f a c e of t h e combustor. The design of combustor A w a s t h e r e s u l t of t h e research described i n re- ference 8 aimed toward t h e development of a high-performance combustor f o r high-altitude, subsonic f l i g h t conditions.
Combustor B w a s s p e c i f i c a l l y designed t o meet t h e requirements of high-altitude, high Mach number f l i g h t of an engine with advanced design is encountered a t the design components. Since a high combustor velocity f l i g h t conditions, maintaining a m i n i m u m pressure loss i n combustor B w a s a primary consideration. A n attempt was made t o reduce t h e annular l o s s e s by designing combustor B with a somewhat smaller combustion space ( f i g .
Making t h e combustion space small, however, a l s o adversely a f f e c t s 2 ( b ) ) .
combustion efficiency, p a r t i c u l a r l y a t low pressures ( r e f . 7 ) . Analyti- c a l s t u d i e s of s e v e r a l f l i g h t missions of i n t e r e s t f o r supersonic turbo- j e t a i r c r a f t have indicated t h a t t h e combustor pressure w i l l be above about 1 atmosphere at all f l i g h t conditions considered i n t h e a n a l y s i s (unpublished d a t a ) . It w i l l t h e r e f o r e not be necessary t o meet t h e r e - quirements of providing high e f f i c i e n c i e s a t very low pressures; t h i s
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makes p o s s i b l e a compromise i n combustor design t o o b t a i n lower p r e s s u r e losses a t t h e expense of combustion e f f i c i e n c y a t low pressures.
Primary a i r was admitted i n t o combustor B through a number of inverted I n addition, a r e l a t i v e l y l a r g e propor- louvers, ae shown i n f i g u r e 2 ( b ) .
t i o n of t h e primary air w a s admitted i n t h e upstream half of cornbustor B ( s e e f i g . 4) which served t o f u r t h e r decrease t h e pressure-loss coef- Provision was made f o r f u e l staging i n combustor B s i n c e t h i s f i c i e n t .
technique waa indicated t o be p a r t i c u l a r l y advantageous at high heat- r e l e a s e conditions (ref. 9 ) . When t h e combustor w a a operated without f u e l staging, a l l t h e f u e l w a e i n j e c t e d through nine, hollow-cone, swirl-type nozzles (10.5 gal/hr; 60' spray angle) a t t h e upstream end of t h e com- bustor l i n e r . During operation with f u e l staging, two-thirds of t h e f u e l w a e injected through e i g h t fan-spr8y i n j e c t o r s located 6 inches downstream and spraying r a d i a l l y i n t o t h e combustor as shown i n f i g u r e l ( b ) .
Combust o r I n s t a l l a t i o n A schematic diagram of t h e combustor i n s t a l l a t i o n is shown i n f i g u r e 5. A i r of desired quantity, pressure, and temperature wae drawn from t h e laboratory air-supply system, passed through t h e combustor, and exhausted i n t o t h e altitude-exhaust system. Combustor-inlet temperatures were con- t r o l l e d by use of a g a s o l i n e - f i r e d preheater which burned a p o r t i o n of t h e a i r upstream of t h e combustor. The quantity of air flowing through t h e preheater, t h e t o t a l air flow, and t h e combustion-chamber s t a t i c pressure were regulated by three remote-control valves. Two observation windows were i n s t a l l e d i n t h e t e s t s e c t i o n i n order t o permit v i s u a l observation of the combusticn process.
Instrumentation T o t a l temperatures and p r e s s u r e s were measured at t h e t h r e e s t a t i o n s indicated i n f i g u r e 5. The p o s i t i o n of t h e instruments i n each of t h e t h r e e planes is shown i n f i g u r e 6. Combustor-inlet t o t a l temperatures were measured with t h r e e bare-junction, unshielded, iron-constantan thermo- couples a t s t a t i o n 1, as shown i n f i g u r e 6 ( a ) . S l i g h t l y upstream were located 1 2 t o t a l - p r e s s u r e tubes, t h r e e tubes i n each of f o u r rakes as shown in f i g u r e 6 ( a ) . Combustor-outlet t o t a l temperatures were measured w i t h 30 bare-junction, unshielded, chromel-alumel thermocouples; f i v e thermocouples i n each of f i v e rakes were located across t h e duct a t s t a t i o n A t s t a - 2 , 23 inches from t h e upstream end of t h e combustor (fig. 6 ( b ) ) .
t i o n 3 were located 15 t o t a l - p r e s s u r e tubes i n t h r e e rakes of f i v e pres- s u r e tubes each ( f i g . 6 ( c ) ) .
A l l instruments were located a t approximate centers of equal a r e a s . S t a t i c - p r e s s u r e o r i f i c e s were i n s t a l l e d a t t h e w a l l , as shown i n f i g u r e 6 ( c ) . Construction d e t a i l s of t h e pressure and temperature probes a r e shown i n f i g u r e 7 .
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NACA RM E54Al-y 5 Rotameters were used t o meaeure the f u e l flow; MIL-F-5624A grade Jp-4 f u e l w a s used throughout t h e investigation.
PROCEDURE The combustor w a s operated at conditions considered t o b e r e p r e - 7, f l i g h t Mach num- s e n t a t i v e f o r an engine with a compressor r a t i o of bers from 2.0 t o 3.0, and f l i g h t a l t i t u d e s from 60,000 t o 80,000 feet.
I I F l i g h t analyses such as shown i n reference 10 i n d i c a t e d t h a t t h e m i n i m i i m c~mbust.or-inl& p r ~ s s i c e g enccimt.ered would be &ox.re 1 E?+,~cs- phere. The minimum combustor-inlet pressure w a s 18.4 pounds p e r square inch absolute f o r a given i n t e r c e p t o r f l i g h t p l a n and w a s t h e r e f o r e chosen as a standard t e s t p o i n t . Combustor-inlet p r e s s u r e s of 10 and 30 pounds p e r square inch absolute were also included i n t h e t e s t schedule t o show t h e e f f e c t on performance of v a r i a t i o n s i n i n l e t pressure.
Com- bustor reference v e l o c i t i e s typical f o r t h e s e conditions of supersonic f l i g h t ranged from approximately 150 t o 200 f e e t p e r second. Data were obtained over a range of v e l o c i t i e s from 125 t o 225 feet per second t o determine t h e e f f e c t of v a r i a t i o n i n velocity on performance. Minimum combustor-inlet temperatures w e r e determined t o be about 870° F, and t h e r e f o r e t h i s value w a s chosen as a standard t e s t parameter. Turbine- i n l e t temperatures of 2000' F have been shown (ref. 10) t o be d e s i r a b l e f o r obtaining t h e high t h r u s t necessary f o r high supersonic speed. Be- cause of instrumentation l i m i t a t i o n s , average combustor-outlet tempera- t u r e s were maintained at 1 8 0 0 ' F f o r most of t h e runs; a s i n g l e r u n w a f ~ made a t a 2000' F o u t l e t temperature w i t h combustor B. The t e s t condi- t i o n s are shown i n t a b u l a r form i n t h e following t a b l e :
C ombus t or - Combustor - C ombk t o r Combustor -
i n l e t t o t a l inlet t o t a l reference o u t l e t temperature, velocity, temperature, pressure,
lb/sq i n . abs ft/s ec 9
OF (a) 165 1800 870 165 1800 870 1 2 5 1800 870 165 2 04 1800 87 0 1800 2 04 2 000
I 18.4 876
~ &Based on maximum combustor cross-sectional area of 105 sq i n , and combustor-inlet a i r density.
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Combustion e f f i c i e n c y , o u t l e t temperature p r o f i l e , and pressure losses were evaluated f o r each combustor. Combustion e f f i c i e n c y w a s com- puted as t h e percentage r a t i o of a c t u a l t o t h e o r e t i c a l increase i n en- t ha1py from t h e combust o r - i n l e t t o t h e c ombus t o r -out 1 e t instrument a t i o n planes by using t h e method of r e f e r e n c e 11. The a r i t h m e t i c mean of t h e 30 o u t l e t thermocouple readings w a s used t o o b t a i n t h e value o f t h e combustor-outlet enthalpy. The accuracy of t h e combustion e f f i c i e n c y calculated from t h e s e readings w a s considered t o be about ~3 percent.
The r a d i a l outlet-temperature d i s t r i b u t i o n w a s determined f o r an average o u t l e t temperature of approximately 1 8 0 0 ' F. The temperature a t each of f i v e r a d i a l p o s i t i o n s was computed as t h e average of s i x c i r c u m f e r e n t i a l thermocouple readings a t each p o s i t i o n . The pressure loss w a s computed as the percentage r a t i o of pressure loss through t h e combustor t o t h e i n l e t t o t a l pressure.
RESULTS AND DISCUSSION The performance of two experimental annular combustors, over a limited range of operating conditions t h a t a r e r e p r e s e n t a t i v e of super- are discussed subsequently. The performance c r i t e r i a con- sonic f l i g h t , sidered include combustion e f f i c i e n c y , combustor pressure loss, and o u t l e t -temperature p r o f i l e .
C ombus t i on E f f i c ienc y E f f e c t of v e l o c i t y . - The e f f e c t of combustor reference v e l o c i t y on combustion e f f i c i e n c y is shown i n f i g u r e 8 f o r each of t h e two combustors operating at a constant value of i n l e t - a i r temperature of 870' F, inlet- air pressure of 18.4 pounds per square inch absolute, and a n average combustor-outlet temperature of approximately 1800' F.
D a t a a r e shown f o r a range of reference v e l o c i t i e s from 1 2 5 t o 225 f e e t per second.
Combustor v e l o c i t y , as discussed herein, is based on t h e d e n s i t y of t h e combustor-inlet a i r and on t h e maximum c r o s s - s e c t i o n a l area of t h e com- bustor. The combustion e f f i c i e n c y of combustor A w a s e s s e n t i a l l y 100 percent a t a l l v e l o c i t i e s i n v e s t i g a t e d except t h e lowest v e l o c i t y (125 f t / s e c ) where t h e combustion e f f i c i e n c y w a s 97 percent.
For combustor B without f u e l s t a g i n g t h e combustion e f f i c i e n c y decreased from 100 p e r - cent a t a reference v e l o c i t y of 165 f e e t per second t o 88 percent a t Fuel s t a g i n g served t o improve t h e per- 225 f e e t per second ( f i g . 8).
formance of combustor B a t t h e higher v e l o c i t i e s . A t a reference velo- c i t y of 2 2 5 f e e t per second, t h e combustion e f f i c i e n c y of combustor B w i t h f u e l s t a g i n g w a s 97 percent. A s i n g l e d a t a point w a s obtained f o r combustor B with f u e l s t a g i n g a t a higher combustor-outlet temperature (2000' F ) .
This d a t a point is included i n f i g u r e 8 and shows t h a t t h e combustion e f f i c i e n c y remained high a t t h i s higher o u t l e t temperature.
. . % The d a t a presented i n f i g u r e 8 i n d i c a t e t h a t high combustion e f - f i c i e n c y can be obtained a t t h e high combustor r e f e r e n c e v e l o c i t i e s t h a t a r e a n t i c i p a t e d i n f u t u r e t u r b o j e t engines operating a t high a l - t i t u d e s and supersonic f l i g h t speeds. Moreover, t h e high combustion e f - w i t h a t l e a s t two experimental combustors of f i c i e n c i e s can be obtained s i g n i f i c a n t l y d i f f e r e n t des ign .
E f f e c t of pressure. - The e f f e c t of combustor-inlet pressure on
t h e combustion e f f i c i e n c i e s of each of t h e two combustors is shown i n fig12rp_ 9 f e r a c m r l t s p t cp?l?llJgter--in.let-BiT t e q e y s t u r e nf e700 F, average o u t l e t temperature of 1800' F, and a r e f e r e n c e v e l o c i t y of 165 f e e t per second. Above 18.4 pounds p e r s q u a r e inch absolute, t h e com- b u s t i o n e f f i c i e n c y was approximately 100 percent f o r both combustors; however, as t h e pressure w a s reduced t o 10 pounds per square inch absolute, t h e combustion e f f i c i e n c y of combustor A decreased t o 82.5 percent and t h a t of combustor B, t o 62.5 percent. The marked e f f e c t of low pressure on t h e cambustion e f f i c i e n c y of combustor B i s t h e r e s u l t of t h e design compromises previously noted (small combustion space and r a p i d e n t r y of primary a i r ) . The e f f e c t of low pressures on t h e ef- p a r t l y due t o the f a c t t h a t t h i s combustor f i c i e n c y of combuetor A is configuration w a s developed ( r e f . 6 ) f o r use with a f u e l prevaporizer, while i n t h i s i n v e s t i g a t i o n l i q u i d f u e l was used. Combustor pressures below about 1 atmosphere would n o t be encountered i n t h e t u r b o j e t - powered a i r c r a f t capable of f l i g h t at high supersonic Mach numbers which were considered i n a n a l y t i c a l s t u d i e s conducted a t t h i s laboratory; as shown i n f i g u r e 9, near 100-percent combustion e f f i c i e n c y w a s obtained with both combustors a t t h e s e conditions.
The i n d i c a t e d combustion e f f i c i e n c i e s a t p r e s s u r e s of 10 pounds p e r square inch absolute may be low by several percentages because of Oxygen oxygen d e p l e t i o n i n t h e i n l e t air due t o the g a s - f i r e d preheater.
depletion has been shown t o have a more severe e f f e c t at low p r e s s u r e s ( r e f . 12).
C ombus t or -Out l e t Temperature Prof il e Typical combustor-outlet isothermal contour p a t t e r n s f o r combustors A and B a r e shown i n f i g u r e 10, and t h e r a d i a l outlet-temperature pro- f i l e s i n f i g u r e 11. A maximum average temperature d e v i a t i o n from inner t o outer w a l l of 220' F w a s obtained with combustor A and a m a x i m u m de- v i a t i o n of 70' F vas obtained with combustor B. Preliminary a n a l y s i s has indicated t h a t uniform temperature d i s t r i b u t i o n s such as t h e s e are p a r t i c u l a r l y a p p l i c a b l e f o r cooled t u r b i n e blades t h a t may be used i n engines f o r high supersonic f l i g h t , inasmuch as t h e p r e f e r r e d gas- temperature p r o f i l e f o r cooled t u r b i n e blades i s r a d i a l l y more uniform than f o r uncooled t u r b i n e s . Previous stcLdies (ref. 13) describe methods of c o n t r o l l i n g outlet-temperature p r o f i l e s . It i s expected t h a t no ......................... . . . . . . .
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s i g n i f i c a n t s a c r i f i c e i n o t h e r performance c h a r a c t e r i s t i c s would be re- quired t o provide temperature p r o f i l e s d i f f e r e n t from those shown i n f i g u r e 11.
Combustor P r e s s u r e Losses The percent t o t a l - p r e s s u r e loss of each of t h e two combustors is The pressure shown BB a f u n c t i o n of reference v e l o c i t y i n f i g u r e 12.
loss of combustor A is approximately 20 percent as compared with 15 per- c e n t f o r combustor B a t a r e f e r e n c e v e l o c i t y of 204 f e e t per second and a temperature r a t i o across t h e combustor of about 1 . 7 . The lower pressure l o s s e s obtained with combustor B a r e t h e r e s u l t of t h e f e a t u r e s ( s m a l l combustion space and r a p i d e n t r y of primary a i r ) t h a t were in- corporated i n t h e design t o o b t a i n lower pressure-loss c o e f f i c i e n t s .
1 2 a r e not t o The pressure l o s s e s represented by t h e curves of f i g u r e b e considered t h e m i n i m u m required f o r high e f f i c i e n c y a t t h e conditione investigated, s i n c e t h e d e s i g n v a r i a b l e s were i n v e s t i g a t e d t o a very 1 i m i t ed e x t e n t .
The s p e c i f i c - f u e l consumption, as a f u n c t i o n of t h e t o t a l - p r e s s u r e l o s s e s , was c a l c u l a t e d by using t h e method of reference 1 4 f o r a rep- r e s e n t a t i v e subsonic and supersonic f l i g h t condition. The s p e c i f i c - f u e l consumption is p l o t t e d i n f i g u r e 13 as t h e r a t i o of t h e a c t u a l t o t h e i d e a l s p e c i f i c - f u e l consumption with no pressure loss i n t h e
combustor assumed. A p r e s s u r e loss of about 3 percent, which is ob-
t a i n e d i n many c u r r e n t combustors f o r subsonic f l i g h t conditions, re- s u l t s i n an i n c r e a s e i n s p e c i f i c - f u e l consumption of about 2 . 3 percent i n the 5:l p r e s s u r e - r a t i o engine, as shown i n f i g u r e 13. For t h i s same e f f e c t on specif i c - f u e l consumption, pressure l o s s e s of 9.2 percent a r e permitted at t h e supersonic f l i g h t condition i n a 7 : l compressor it is evident t h a t higher pressure l o s s e s pressure engine; t h e r e f o r e , can be t o l e r a t e d i n t h e engine f o r supersonic f l i g h t t h a n i n c u r r e n t engines for subsonic f l i g h t while equivalent performance l e v e l s a r e maintained. Pressure loss has a l e s s e r e f f e c t on s p e c i f i c - f u e l con- sumption a t t h e supersonic f l i g h t conditions mainly because t h e ram- temperature-rise r a t i o s encountered i n supersonic f l i g h t a r e high ( r e f .
1 4 ) . I n any application, however, it is obviously d e s i r a b l e t o design f o r a minimum value of pressure l o s s .
Carbon and D u r a b i l i t y D u r i n g t h e i n v e s t i g a t i o n which included operation a t pressures as high as 30 pounds per square inch absolute, no carbon d e p o s i t s were evident; however, with sustained high-temperature operation over several hours moderate t o s e v e r e liner d e t e r i o r a t i o n occurred.
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CONCLUDING RENARKS The performance r e s u l t s presented indicate t h a t combustion e f f i c i e n - c i e s over 95 percent and s a t i s f a c t o r y outlet-temperature p r o f i l e s can be obtained i n annular combustors operating a t simulated Supersonic f l i g h t conditions with combustor v e l o c i t i e s a8 high a8 225 feet per second.
These high e f f i c i e n c i e s were maintained t o t h e highest combustor-mtlet temperatures investigated, 1800° F f o r one combustor and 2000' F f o r t h e other. The t o t a l - p r e s s u r e l o s s e s of t h e two experimental cambustors xere acceptzble by prenent. nt.mda.rc-~ e i n w c a l c u l a t i o n s indicated t h a t t h e i n c r e a s e i n engine s p e c i f i c - f u e l consumption r e s u l t i n g from combustor pressure l o s s e s would not be s i g n i f i c a n t l y g r e a t e r i n t h e engine f o r supersonic propulsion than i n current t u r b o j e t engines. These pressure l o s s e s , t h e r e f o r e , appear acceptable f o r t h e supersonic f l i g h t conditions.
Further engine performance g a i n could be r e a l i z e d , however, i f t h e l o s s e s could be reduced. Liner d u r a b i l i t y may well prove t o be one of t h e l a r g e s t problems facing t h e combustor designer f o r combustor a p p l i c a t i o n s involving temperature levels of i n t e r e s t f o r high supersonic f l i g h t .
Lewis F l i g h t Propulsion Laboratory National Advisory Committee f o r Aeronautics Cleveland, Ohio, January 18, 1954 REFERENCES 1. Lieblein, Seymour, Lewis, George W., Jr., and Sandercock, Donald M.: Experimental I n v e s t i g a t i o n of an Axial-Flow Compressor Inlet Stage
Operating a t Transonic R e l a t i v e I n l e t Mach Numbers. I - Over-All
Performance of Stage with Transonic Rotor and Subsonic S t a t o r s up t o Rotor R e l a t i v e I n l e t Mach Number o f 1.1. NACA RM E52A24, 1952.
2 . Serovy, George I C . , Robbins, W i l l i a m H., and Glmer, Frederick W. : Experimental I n v e s t i g a t i o n of a 0.4 Eub-Tip Diameter R a t i o Axial- Flaw Compressor I n l e t Stage a t Transonic Inlet R e l a t i v e Mach Num-
b e r s . I - Rotor Design and Over-All Performance at Tip Speeds from
60 t o 100 Percent of Design. NACARM E53111, 1953.
I 3. Volt, Charles H. : I n v e s t i g a t i o n of a High-Pressure-Ratio E i g h t S t a g e I Axial-Flow Research Compressor w i t h Two Transonic Inlet Stages.
I I - Aerodynamic Design. NACA RM E53124, 1953.
4 . Geye, Richard P., Budinger, R a y E., and Voit, Charles H.: I n v e s t i - g a t i o n of a High-Pressure-Ratio Eight-Stage Axial-Flow Research
Compressor with Two Tramonic I n l e t Stages. I1 - Preliminary Amly-
sis of Over-All Performance. NACA RM E55J06, 1953.
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10 ' { A C A RM E54Al5 5. Schramm, Wilson B., Nachtigall, Alfred J., and Arne, Vernon L.: Analytical Comparison of Turbine-Blade Cooling Systems Designed f o r a Turbojet Engine Operating a t Supersonic Speed and High Al- t i t u d e . I - Liquid-Cooling Systems. NACA R M E52J29, 1953.
6. Schramn, Wilson B., Arne, Vernon L., and Nachtigall, Alfred J.: Analytical Comparison of Turbine-Blade Cooling Systems Designed f o r a Turbojet Engine Operating at Supersonic Speed and High Al- t i t u d e . I1 - Air-Cooling Systems. NACA RM E52J30, 1953.
7. C h i l d s , J. Howard, McCafferty, Richard J., and Surine, Oakley W.: Effect of Combustor-Inlet Conditions on Performance of an Annular Turbojet Combustor. NACA Rep. 881, 1947. (Supersedes NACA ! I ! N 1357.)
8. Norgren, C a r l T., and Childs, J . Howard: Performance of an Annular
Turbojet Combustor Having Reduced P r e s s u r e Losses and U s i n g Propane Fuel. NACARM E53G24, 1953.
9. Z e t t l e , Eugene V., and Mark, Herman: E f f e c t of Axially Staged F u e l I n t r o d u c t i o n on Performance of One-Quarter S e c t o r of Annular Turbo- j e t Combustor. NACA RM E53A28, 1953.
10. Gabriel, David S., m e b s , Richard P., Wilcox, E. Clinton, and Koutz, Stanley L.: Analysis of t h e Turbojet Engine f o r Propulsion of Supersonic F i g h t e r Airplanes. NACA RM E52F17, 1953.
11. Turner, L. Richard, and Bogart, Donald: Constant-Pressure Combustion Charts Including E f f e c t s of Diluent Addition. NACA Rep. 937, 1949.
(Supersedes NACA TN's 1086 and 1655.)
1 2 . Graves, Charles C . : E f f e c t of Oxygen Concentration of t h e Inlet Oxygen-Nitrogen Mixture on t h e Combustion Efficiency of a Single 533 Turbojet Combustor. NACA RM E52F13, 1952.
13. Mark, Herman, and Z e t t l e , Eugene V.: Effect of A i r D i s t r i b u t i o n on R a d i a l Temperature D i s t r i b u t i o n i n One-Sixth S e c t o r of Annular Turbojet Combustor. NACA RM E9122, 1950.
14. Pinkel, Benjamin, and Karp, I r v i n g M . : A Thermodynamic Study of t h e Turbojet Engine. NACA Rep. 891, 1947. (Supersedes NACA W R E-241.)
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Figure 1. - Cutaway view d experinental mnular turbojet oambuetors assembled in housing.
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