. NASA Technical Memorandum 100107 AIAA-87-1828

AVSCOM Technkal Report 87-C-8

Liner Cooling Research at NASA Lewis Research Center

(bASA-!€d-lOOlG’?) IINEB C C C L l i G EESEABCH AT 887-23624 L A S A LERIS EESEBBCB CEBXEB 0 A L E . A ) 18 p A v a i l : EXIS EC EO2/EF A01 CSCL 21E

Unclas G3/07 0076755

Waldo A. Acosta Propulsion Directorate U. S. Army Aviation Research and Technology Activity--AVSCOM Lewis Research Center Cleveland, Ohio

Prepared for the 23rd Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE San Diego, California, June 29-July 2, 1987

https://ntrs.nasa.gov/search.jsp?R=19870014191 2018-06-21T20:43:01+00:00Z

LINER COOLING RESEARCH AT NASA LEWIS RESEARCH CENTER

Waldo A. Acosta Propuls ion D i rec to ra te

Lewis Research Center Cleveland, Ohio 44135

U.S. Army Av ia t i on Research and Technology A c t i v i t y - AVSCOM

SUMMARY

This paper descr ibes recen t l y completed and cu r ren t advanced l i n e r research app l i cab le t o advanced small gas t u r b i n e engines. t o the e v o l u t i o n o f f u e l e f f i c i e n t s m a l l gas t u r b i n e engines capable o f meeting f u t u r e commercial and m i l i t a r y a v i a t i o n needs i s c u r r e n t l y underway a t NASA Lewis Research Center.’ As p a r t o f t h i s research, a reverse- f low combustor geometry was maintained w h i l e d i f f e r e n t advanced l i n e r w a l l coo l i ng techniques were i nves t i ga ted and compared t o a base l ine combustor. combustors f e a t u r i n g counter- f low f l lm-cooled l i n e r (CFFC) panels, t ransp f ra -

h d t i o n cooled l i n e r wa l l s (TRANS), and compl iant metal/ceramic (CMC) w a l l s was a M obta ined over a range o f s imulated f l i g h t cond i t i ons o f a 16: l pressure r a t l o w I gas t u r b l n e engine and f u e l a i r r a t i o s up t o 0.034. A l l t he combustors fea-

tu red an i d e n t i c a l f u e l i n j e c t i o n system, i d e n t i c a l geometric c o n f i g u r a t i o n o u t l i n e , and s i m i l a r designed i n t e r n a l aerothermodynamics.

Research r e l a t i n g

The performance o f t he

INTROOUCT I O N

Problems unique t o smal l combustors were reviewed dur ing a forum a t NASA Lewis conducted by A.D. L i t t l e , I nc . ( r e f . 1). The o b j e c t i v e was t o i d e n t i f y the R&D e f f o r t which must be considered i n the 1980 t o 1990 t ime frame t o meet the c r i t i c a l needs for smal l a l r c r a f t gas tu rb ines . Advancements i n gas t u r b i n e c y c l e e f f i c i e n c y and lowered s p e c i f i c f u e l consumption were pr imary considerat ions. These improvements can be r e a l i z e d by Increas ing c y c l e pres- sure r a t i o s and tu rb lne i n l e t temperature, and/or implementing e i t h e r regenera- t i v e o r recupera t ive cycles. Combustors opera t ing a t these cond i t i ons r e q u i r e increased amounts o f coo l i ng a i r . A t t he same time, the heat s ink c a p a b i l i t y of the a i r has been decreased because of h igher combustor i n l e t temperature r e s u l t i n g f rom increased pressure r a t i o and/or regenerat ive cyc le operat ion. This l oss i n heat s ink c a p a b i l i t y i s p a r t i c u l a r l y c r i t i c a l i n the smal l combus- t o r system due t o the inherent h igh surface-to-volume r a t i o as compared w i t h l a r g e combustion systems. forum t h a t l i n e r coo l i ng was one of the most impor tant areas t h a t should be considered. Advanced coo l i ng techniques must be incorporated i n the combustor i n order t o achieve the p o t e n t i a l for implementing advanced cycles.

I t was t h e concensus o f the p a r t i c i p a n t s i n the

The l i n e r coo l i ng approach predominately used i n present day smal l gas t u r b i n e combustors i s splash f i l m cool ing. The coo l i ng a i r enters the l i n e r through a row o f small diameter holes. The a i r j e t s impinge on a coo l i ng s k i r t , which then d i r e c t s the f l o w so as t o form a f i l m along the i n s i d e o f t he l i n e r w a l l . The t o t a l amount o f cool ing a i r requ i red depends on a number o f f a c t o r s ; however, t y p i c a l values range from 30 t o 50 percent o f t h e combustor a i r f l o w . Improvements i n l i n e r cool lng ef fect iveness can be achieve by us ing a more e f f i c i e n t a i r d i s t r i b u t i o n and t h e heat s ink a v a i l a b l e i n t h e coo l i ng

This paper is declared a work of the U.S. Government and is not subject to copy*ht proleclion in Ihe United Stales.

air (ref. 2). require little or no cooling.

Another alternative is the use of advanced materials that

A component research program for advanced small gas turbine engines was

The objective of the program is to identify and evaluate new initiated as a result of the forum's recommendations and is currently underway at NASA Lewis. and promising liner cooling techniques for advanced combustors.

liner designs operating at advanced cycle conditions is presented. The per- formance of the combustor was obtained over a range of simulated flight condi- tions of a 16:l pressure ratio gas turbine engine and fuel air ratios up to 0.034. This is an extension to the work previously reported in which the com- bustors were operated at fuel air ratios up to 0.024 (refs. 3 and 4).

In this paper, a reverse flow combustor equipped with different advanced

In addition to the reverse flow combustor liner research being conducted additional liner research currently in progress is described. These activities consist of a compliant metal/ceramic experimental combustor program with Allison Gas Turbine Engine Operations and a more fundamental liner test rig designed to screen advanced liner cooling techniques.

APPARATUS

Test Facility

lhe combustor was mounted in duct A of the test facility CE9B (fig. 1 ) located in the Engine Research Building (Bldg. 5) at NASA Lewis. Tests were conducted with inlet-air pressure ranging up to 16 atm with the air indirectly heated to about a temperature of 717 K (830 O F ) . The temperature of the air was automatically controlled by mixing the heated air with varying amounts o f cold by-pass air. Airflow through the heat exchanger and by-pass flow system and the total pressure of the combustor were regulated by remotely controlled valves.

Combustors

The basic reverse flow combustor used in this investigation was a full scale experimental design. The design was based on versatility so that the interchanging of fuel injectors and the modification or replacement of the swirlers, faceplate, and liner could be readily accomplished. The airflow distribution and hole sizing of the liner were based on 36 primary holes and 36 dilution holes. The design liner isothermal pressure loss is 1.5 percent and the dump loss is 0.24 percent. A cross section and isometric views of the combustor are shown i n figure 2.

Transpiration (TRANS) liner. - Lamilloy was used to simulate transpiration cooling (continuous mass flux of coolant over the surface). A Lamilloy 1s a commercial product composed of an electrochemically etched channel structure in multilayers which is diffusion-bonded forming a sheet. This channel and multilayer structures allow the use of the heat sink capability inherent in the coolant air prior to injection into the combustion gases. An isometric sketch of the material is shown schematically in figure 3;

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The Lami l loy combustor was fabr ica ted under government con t rac t t o be s i m i l a r t o the reverse f l o w f i lm-cooled geometry and t o i n t e r f a c e w i t h the e x i s t i n g NASA t e s t f a c i l i t y . bustor opera t ing pressure of 16 atm, 717 K (830 O F ) i n l e t temperature, and 1930 K (2500 O F ) e x i t w i t h ho t streaks up t o 1922 K (3000 O F ) .

are presented i n re ference 5.

The basic design cond i t i ons were f o r a peak com-

Design d e t a i l s

Counter-flow f i lm-coo led (CFFC) l i n e r . - The CFFC l i n e r design incorpo- ra ted t h e use o f f i n s t o increase the heat t r a n s f e r . t o determine among the e x i s t i n g and ava i l ab le o f f s e t - f i n t o o l i n g , a conf igura- t i o n t h a t would ensure acceptable wa l l temperature and minimum f r i c t i o n losses. The f i n a l f i n and channel he igh t was based on producing t h e most s i g n i f i c a n t w a l l temperature reduc t ion f o r a given pressure l oss and e x i s t i n g t o o l d ies . The o f f s e t f i n p l a t e s were d i s t r i b u t e d a long the i nne r and ou ter c y l i n d r i c a l l i n e r s . f e ren t coo l i ng techniqOes employed.

The prime o b j e c t i v e was

The cross sec t i on o f t he CFFC l i n e r shown i n f i g u r e 4 shows the d i f -

The CFFC combustor l i n e r was fabr ica ted under government con t rac t by the Garret Turbine Engine Company. f o r the TRANS l i n e r . re ference 6.

The basic design cond i t i ons were the same as More d e t a i l s about i t s design can be found i n

ComDliant metal/ceramic (CMC) l i n e r . - The o b j e c t i v e o f t he CMC l i n e r program was t o i n v e s t i g a t e advance l i n e r concepts capable of w i ths tand ing temperature l e v e l s o f 1922 K (3000 O F ) , w h i l e p r o v i d i n g improved c y c l i c dur- a b i l i t y w i t h l i t t l e o r no coo l i ng a i r . The CMC concept consis ted o f plasma spray ing an y t t r i a p a r t i a l l y s t a b i l i z e d z i r c o n i a (YPSZ) on a compl iant n i c k e l a l l o y subst rate. This compl iant metal subs t ra te was designed t o y i e l d a t r e l a t i v e l y low l e v e l s o f s t ress , thereby absorbing the d i f f e r e n t i a l expansion between the metal and the ceramic as the ma te r ia l i s heated. c ross-sec t iona l view o f the CMC l iner concept.

F igure 5 shows a

I n t h i s f i r s t order approach i t was e lec ted t o l i m i t the combustor o u t l e t temperature t o approximately 1789 K (2760 O F ) r a t h e r than s t r i v e f o r t he o r i g i n a l 1922 K (3000 O F ) goal. on ly backside convect ive coo l ing .

This al lowed the major s i m p l i f i c a t i o n o f us ing

INSTRUMENTATION

The combustor ins t rumenta t ion s ta t ions are shown i n f i g u r e 6. F i ve t o t a l pressure probes, two s t a t i c pressure taps, and f o u r Chromel-Alumel thermo- couples a re located a t s t a t i o n 2 t o measure the i n l e t pressure and temperature. A t s t a t i o n 3, a ser ies o f 18 t o t a l pressure probes a re i n s t a l l e d t o determine the i n l e t - a i r p r o f i l e and t o determine the ex ten t o f any f l o w d is turbance behind the s t r u t s which support t h e centerbody d i f f u s e r . p i t o t - s t a t i c probes a re pos i t ioned i n the c o l d - a i r passages between the com- bus tor l i n e r and combustor housing t o determine passage v e l o c i t y and d i s t r i b u - t i o n . A t s t a t i o n 5, o u t l e t temperature and pressure measurements were obta ined by means o f a r o t a t i n g probe. The probe conta ins th ree rakes spaced 120' apa r t a f i v e - p o s i t i o n r a d i a l rake conta in ing PT-PT 13 percent Rd thermocouples, a f i v e - p o s i t i o n t o t a l pressure rake, and a water cooled gas sampling rake. A 360' t r a v e l w i t h sampling a t 10" increments was used f o r the CFFC and CMC l i n e r programs.

A t s t a t i o n 4, s i x

The TRANS l i n e r program used a s t a t i o n a r y probe a t s t a t i o n 5 t o

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measure o u t l e t temperature and pressure. evenly spaced w i th 4 probes/rake con ta in ing PT-PT 13 percent Rd thermocouples and f o u r gas sample/pressure probes evenly spaced on the circumference.

The probe had 12 temperature rakes

PROCEDURE

Test Condi t ions

The experimental reverse f l o w combustor was operated a t t e s t cond i t i ons based on a gas-turbine engine c y c l e w i t h a compressor pressure r a t i o o f 16. t a b u l a t i o n of the t e s t cond i t ions used i n t h i s study i s g iven i n t a b l e I .

A

Slmulated f l i g h t data were obtained a t f u e l - a i r r a t i o s up t o approximately 0.034. The simulated combustor t e s t cond i t ions were based on a reference v e l o c i t y o f 5.49 m/sec'(18 f t / s e c ) . on u n i d i r e c t i o n a l t o t a l mass f l o w and the maximum cross-sec t iona l area o f t he housing p r i o r t o the reverse t u r n ( f i g . 6).

The reference v e l o c i t y quoted was based

The t e s t program was conducted us ing Jet-A f u e l w i t h 18 simplex pressure- a tomiz ing f u e l i n j e c t o r s w i t h a f l o w number o f 4.8.

Emission Measurements

Exhaust gas samples were obtained according t o t h e recommended procedures i n references 7 and 8. cooled probes mounted approximately i n the s t a t o r p lane and i n the center o f the exhaust duct a t s t a t i o n 5 ( f i g . 6) . The gas sample temperature was he ld a t approximately 423 K (302 O F ) i n the e l e c t r i c a l l y heated sampling l i n e . of the gas sample entered the analyzer oven, w h i l e the excess sample was bypassed t o the exhaust system. l i n e , a n i t rogen purge was used be fore and du r ing combustor i g n i t i o n .

Exhaust gases were withdrawn through the f o u r water

Most

To prevent f u e l accumulat ion i n the sample

A f t e r passing through the analyzer oven the gas sample was d i v ided i n t o th ree pa r t s , and each p a r t was analyzed. Concentrat ions o f oxldes o f n i t rogen , carbon monoxide and carbon d iox ide , and hydrocarbons were measured by the chemiluminescence, nondispersed- infrared, and flame i o n i z a t i o n methods, respec- t i v e l y . De ta i l s o f the gas ana lys is system a re presented i n reference 9.

RESULTS AND DISCUSSION

Combustor research r e l a t i n g t o the development o f f u e l e f f i c i e n t smal l gas t u r b i n e engines capable o f meeting f u t u r e commercial and m i l i t a r y a v i a t i o n needs i s cu r ren t l y underway a t NASA Lewis. The goal i s t o evolve the technol - ogy f o r reducing l i n e r coolant f lows by 50 percent o r more f o r increased opera t ing temperatures t o 978 K (1300 O F ) i n l e t , 1922 K (3000 OF) o u t l e t and pressure r a t i o s up t o 30: l .

A component research program f o r advanced smal l gas t u r b i n e engines was i n i t i a t e d towards t h a t goal . a c t i v i t i e s .

The f o l l o w l n g d iscuss ions summarize those

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Advanced Liner-Cooling Techniques

The objective of this program is to identify and evaluate new and promising liner cooling techniques for advanced combustors. The combustors evaluated incorporate advanced liner designs. The experimental study was performed under high pressure and temperature environments simulating those encountered in combustors operating under advanced cycle conditions. Previous experimental results have been reported in references 3 and 4. In reference 3 results between a transpiration (TRANS) cooled liner (continuous mass flux of coolant over the surface) and a liner incorporating counter-flow film-cooled (CFFC) finned panels are presented. These two liners are considered advance metallic designs and demonstrated reductions in liner coolant flow from 40 to 50 percent when compared to the reference splash film-cooled (SF) liner.

A new liner concept incorporating a cornpliant metal/ceramic (CMC) design is presented in reference 4. In that study a simplified compliant ceramic liner concept was selected to facilitate implementation into research hardware. This allowed using only backside convective cooling which was provided by the passage between the combustor wall and casing. The reduction in liner coolant for the CMC configuration is shown i n figure 7 and compared with the TRANS and CFFC configurations using as a reference the SF configuration. The CMC liner coolant flow was reduced by 80 percent.

As shown, the TRANS, CFFC, and CMC liners had similar reductions in liner temperature, but the CMC liner operated at exhaust temperatures 167 K (300 O F )

hotter than current practice. The most significant performance improvements resulting with the CMC liner are the reduction of liner coolant and the increase in turbine inlet temperature. The increase in turbine inlet tempera- ture will allow the engine to increase its cycle efficiency and the reduction of coolant flow frees some air that could be used for turbine cooling or exit temperature pattern improvements.

A new series of experiments were conducted to study the performance of those liners at off-design conditions. Each liner was designed for a maximum liner temperature of 1200 K (1700 O F ) at a fuel air ratio of 0.025 and sea level take-off conditions. For this experiments fuel air ratios up to 0.034 were used.

A comparison of liner temperatures is shown in figure 8. The panel selected for comparison is the inner wall prior to the turn. In general, this is one of the most difficult areas to cool due to the depletion o f the mass flow in the annular passage. The TRANS and the CFFC used about 1.4 percent of the total flow to cool this panel, while the CMC was uncooled. Even though the TRANS and the CFFC used the same amount of cooling the CFFC liner temperatures were an average of 139 K (250 O F ) cooler than the TRANS liner. This indicates a better cooling effectiveness for the finned panels than the transpiration panels assuming similar flame temperatures for this particular location.

The liner temperatures measured on the CMC liner were higher than the 'IRANS and CFFC liners for this particular panel up to a fuel air ratio of about 0.024. This could be explained by the fact that this was an uncooled liner and the convective heat transfer was reduced by the low annular velocities in that region. At fuel air ratios greater than 0.024 the CMC liner temperature is lower than the TRANS liner. At this point the thermal barrier capability of

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the ceramic became ev ident . None of l i n e r s exceeded the 1200 K (1700 O F ) l i n e r temperature l i m i t .

Compliant metal/ceramic combustor. - The r e s u l t s presented here showed the bene f i t s o f using the heat s ink c a p a b i l i t y a v a i l a b l e i n the coo l i ng a i r and ceramtc mater ia ls as thermal b a r r i e r s t o improve the performance o f smal l gas t u r b i n e engine combustor l i n e r s . obtained by combining both approaches.

A more s i g n i f i c a n t improvement could be

A j o i n t NASA/Army program was i n i t i a t e d and a con t rac t awarded t o A l l i s o n Gas Turbine Engine D iv i s ion . The purpose i s t o des ign and exper imenta l l y evaluate a compliant metal/ceramic reverse f low combustor t o operate a t 1922 K (3000 O F ) o u t l e t temperature. The f i r s t phase o f t he study was the s e l e c t i o n o f the ma te r ia l f o r t he composite m a t r i x l i n e r . The approach was t o conduct a l i t e r a t u r e survey, f l o w c a l i b r a t i o n , coo l i ng e f fec t i veness i n v e s t i g a t i o n , thermal s t ress calcula ' t ions, and mic ro s t r u c t u r a l l i f e p red ic t i ons t o f a c i l i - t a t e the pre l im inary design o f the combustor. The candidate ma te r ia l s have been reviewed and the ma te r ia l composi t ion selected. Two design models were formulated and hardware b u i l t t o exper imenta l ly con f i rm the model. models w i l l reso lve the optimum mate r ia l composition, th ickness, coo l ing , and cons t ruc t i on technique by cha rac te r i z ing f lows, and heat t r a n s f e r .

These

The second phase requ i res the implementat ion o f t he r e s u l t i n g ma te r ia l t o the ex ten t o f f u l l combustor p re l im ina ry design. The combustion performance o f t he combustor was analyzed us ing the three-dimensional aerodynamic combustor f l o w - f i e l d designated MARC-1. The gas temperature and f u e l a i r r a t i o pa t te rns as g iven by the three-dimensional code were used t o est imate the temperatures o f the i nne r and outer s h e l l s o f t he annular combustor f o r var ious combustion zones.

The compliant metal/ceramic concept se lected i s shown i n f i g u r e 9. This i s b a s i c a l l y the same concept used i n the in-house a c t i v i t y w i t h the except ion o f the coo lan t f l o w . The coo l i ng a i r f lows through holes on the metal support s t ruc tu re . These h o l e s must be o f a s u f f i c i e n t s ize , number, and d i s t r i b u t i o n t o a l l o w the a i r t o un i fo rm ly pass through t h e compl iant l aye r , p i c k i n g up heat u n t i l i t reaches an opening between t w o ad jacent ceramic t i l e s . I t i s impor- t a n t t h a t the cool ing a i r be p roper l y balanced as i t passes through the com- p l i a n t l aye r so t h a t the c o o l i n g be as un i fo rm as poss ib le no l eav ing uncooled ho t spots. Pre l iminary ana lys is i n d i c a t e t h a t on ly 20 percent o f t o t a l a i r f l o w w i l l be needed f o r coo l ing . This represents a 40 percent l i n e r c o o l i n g reduc- t i o n o v e r s ta te -o f - the -a r t combustors.

Ceramic mat r ix l i n e r . - I n recent years the re has been a r a p i d growth i n research and development aimed a t i n t r o d u c i n g h igh temperature ceramic mate- r i a l s and/or ceramic composites i n t o gas t u r b i n e engines which r e q u i r e l i t t l e o r no coo lan t and minimize o r e l im ina te the need f o r s t r a t e g i c ma te r ia l s .

Many promising candidate ma te r ia l s have been developed du r ing the pas t few years. Experimental eva lua t ion o f t h i s ma te r ia l s i n a f u l l s i z e gas t u r b i n e combustor l i n e r w i l l become very expensive. This creates the need f o r a l i n e r t e s t r i g i n whlch t h i s ceramic ma te r ia l s cou ld be exper imenta l l y evaluated under r e a l i s t i c combustor environments.

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A cont rac t% was awarded t o Teledyne CAE f o r the design and f a b r i c a t i o n o f a ceramic matr'lx combustor l l n e r t e s t r i g compat ib le w i t h the NASA Lewis f a c i l i t y shown i n f i g u r e 1. ma te r ia l p roper t i es and the rank ing of candidate ma te r ia l s f o r use i n gas t u r b l n e combustors. I n a d d i t i o n t o pure ceramic ma te r ia l s such as s i l i c o n n l t r a t e s and carbides o ther semimetal l ic ma te r ia l s w i l l be considered.

Inc luded i n t h i s o b j e c t i v e i s t he ca ta log ing o f

The t e s t r i g being fab r l ca ted i s shown i n f i g u r e 10. The ceramic speci - mens form a square annulus, approximately 10.2 cm ( 4 i n . ) on a s ide by 20.3 cm ( 8 i n . ) long. I n t e r n a l r i g hardware p e r m i t s f l e x i b l e t e s t i n g o f simple, f l a t ceramic specimens. From 8 t o 16 specimens, e i t h e r 8.9 by 10.2 cm (3.5 by 4 in . ) o r 8.9 by 5.1 cm (3.5 by 2 in . ) , in 4 independently c o n t r o l l e d quadrants can be tes ted simultaneously i n t h e r i g . Three panel th icknesses a re def ined: 0.25, 0.38, 0.64 cm. Panels i n any one quadrant must be o f t he same th ickness and a l l quadrants a r e , f i l l e d t o form the f l o w path.

The ceramic ma te r ia l s w i l l be exper imenta l ly evaluated i n an environment s imu la t i ng advanced engine cycles. The t e s t r i g w i l l p rov ide temperatures up t o 2200 K (3500 O F ) and pressures up t o 2068 kPa (300 ps la ) i n e i t h e r reducing o r o x i d a t i n g environments.

CONCLUSIONS

1. The advanced m e t a l l i c l i n e r designs studied, t r a n s p i r a t i o n and counter- f l o w f i lm-cooled, demonstrated reduct ions i n l i n e r coo lan t f rom 40 t o 50 per- cent when compared t o the reference splash f i lm-cooled l i n e r , and performed very w e l l under advanced cyc le condi t ions.

2. Pre l im inary ana lys is i n d i c a t e t h a t the i nco rpo ra t i on o f t r a n s p i r a t i o n coo l i ng i n the compl iant metal/ceramic concept w i l l pave the way f o r the a p p l i c a t i o n o f f u e l e f f i c i e n t , aggressive h lgh temperature cyc les t o small gas t u r b i n e engines.

3. The t e s t f a c i l i t y descr ibed w i l l p rov ide a r e a l i s t i c combustor env l ron- ment f o r the experimental eva lua t ion o f p o t e n t i a l ceramic ma te r ia l s f o r com- bus tor l i n e r s .

REFERENCES

1. Demetrl, E.P., Topplng, R.F., and Wilson, R.P. Jr., "Study o f Research and Development Requirements o f Small Gas-Turbine Combustors,18 ADL-83381-2, A r thu r D , L i t t l e Inc. , Cambridge, MA, Jan. 1980. (NASA CR-159796).

2. Col laday, R.S., "Analysis and Comparison o f Wall Cool ing Schemes f o r Advanced Gas Turbine Appl icat ions," NASA TN 0-6633, 1972.

3. Norgren, C.T. and Riddlebaugh, S.M., "Advanced L iner-Cool ing Technlques f o r Gas Turbine Combustors," A I A A Paper 85-1290, Ju ly 1985. (NASA TM-86952).

4. Acosta, W.A. and Norgren, C.T., ''Small Gas Turbine Combustor Experimental Study - Compliant Metal/Ceramic L iner and Performance .Evaluation,11 A I A A Paper 86-1452, June 1986. (NASA TM-87304).

i

5. P e t r a i t s , J.J., "Design and Fabr i ca t i on o f Laml l l oy Reverse Flow Combustor L iners, " EDR-9803, D e t r o i t Diesel A l l i s o n , I nd ianapo l i s , I N , Apr. 1979.

Referencea I n l e t I n l e t pressure temperature v e l o c i t y

kPa p s i a K 'F mlsec f t l s e c

1014 147 686 775 5.5 18 1358 197 703 805 5.5 18 1620 235 717 830 5.5 18 405 58.5 474 394 5.2 16.9 862 125 627 668 5.5 18 689 100 581 585 --- ---- 517 75 526 486 --- ---- 414 60 474 394 --- ----

6. "Design Documentation Report, Counterf low Film-Cooled Combustor Program," Report 21-4007-A, G a r r e t t Turbine Engine Co., Phoenlx, AZ, June 1982, (NASA CR-167922).

Simulated Comments compressor pressure

r a t i o

10 High-a1 t i tude c r u i s e 13.4 Low-alt i tude c r u i s e 16 Sea l e v e l take-of f (SLTO) 4 I d l e : f / a = 0.008 8.5 Simulated reduced power 6.8 5.1 4.1

7. "Control o f Air P o l l u t i o n from A i r c r a f t and A i r c r a f t Engines, Emission Standards and Test Procedures f o r A l r c ra f t , I l Federal Register, Vol. 38, No. 136, P t . 2, Tuesday, Ju l y 17 , 1973, pp. 19088-19103.

8. "Procedure f o r t h e Continuous Sampling and Measurement o f Gaseous Emissions from A i r c r a f t Turbine Engines," SAE ARP-1256, O c t . 1971.

9. Norgren, C.T. and Riddlebaugh, S.M., " E f f e c t o f Fuel I n j e c t o r Type on Performance and Emissions o f Reverse-Flow Combustor," NASA TP-1945, 1981.

condi- t i o n

H

To ta l a i rf 1 ow

1.23

aParametric v a r i a t i o n based on increase i n mass f low t o prov ide increases o f 33 and 66 percent i n re ference ve loc i t y .

ATMOSPHERIC EXHAUST

ALTITUDE EXHAUST ALTITUDE

EXHAUST

FIGURE 1. - SKETCH OF CLOSED-DUCT TEST FACILITY.

9

- FUEL INJECTOR

‘L REFERENCE AREA, 800

- EXHAUST

kL IN ST RUMENTAT ION STATION 5

-- - - ~- COMBUSTOR CENTE RLlNE

(a)

A I Rf

(a) Cross section. (b) Isometric view.

FIGURE 2. - REVERSE-FLOW COMBUSTOR. (ALL DIMENSIONS ARE I N CENTIMETERS.)

10

COMBUSTION GAS SIDE

COOLING AIRFLOW b---' tLtC' KuCH' C T P U C r l

FIGURE 3. - SCHEMATIC OF LAMILLOY LINER COOLING CONFIGURATION.

- 25.43%

(A) ENHANCED CONVECTION ( 6 ) COUNTER-FLOW FIUVI-COOLED (C) COUNTER-FLOW FILM-COOLED (D) PARALLEL-FLOW FILMCOOLED

FINNED FINNED FINNED

(E) PARALLEL-FLOW FILM-COOLED FINNED (F) PARALLEL-FLW FILMCOOLED FINNED TURNS : COUNTER-FLOVV FILMCOOLED

ENHANCED CONVECTION 0 THERMOCOUPLE LOCATIONS

FIGURE 4. - CROSS-SECTIONAL VIEW OF COUNTER-FLOW FILM-COOLED REVERSE FLOW COMBUSTION LINER INCLUDING CALCULATED AIR-FLOW DISTRIBUTION.

11

HOUSING -,

CERAMIC--

NiETAL SUB STRATE 7,

BRAZE- COMPLIANT

BOND - LAYER ---

r 0.1525 cm ( J- .-/+

/ THICKNESS ANNULUS EXTERNAL / CONVECTION FLOW , /

0635 cm

0.1524 cm I - O - D T 1 7 " z T

MATERIALS

HASTELLOY X

HOSKINS 875 ALLOY WIRE 35% DENSE

YTTRIA STABILIZED I ZIRCONIA

I FIGURE 5 . - CROSS SECTION OF THE COMPLIANT NETAL/CERAMIC LINER CONCEPT.

(NOT TO SCALE. )

OOQP

I

rd

13

L W u L W Q

- LL 0 . 1400 W er - = I- 4 g 1300 n z I- 1200

z 8 1100

g 1000

CL - w

J

I- v) a

V

900

I- z a

REDUCTION IN LINER TEMPERATURE, percent

FIGURE 7. - COMPARISON OF COOLING EFFECTIVENESS FROM A COM- PLIANT METAL/CERAMIC (CMC) LINER WITH A COUNTER-FLOW FILM- COOLED (CFFC) LINER, A LAHILLOY (TRANS) LINER, AND A SPLASH FILM (SF) LINER.

1 100

Y

1000 a !-

W n

I-

s 5 = 900 Y c.(

= W I- v)

800 B u

I 0 TRANS

.014 .018 ,022 .026 ,030 .034 ,038 700 L 800

FUEL A I R RATIO

FIGURE 8. - LINER TEMPERATURE COMPARISON FOR THE INNER LINER PRIOR TO THE TURN. INLET PRESSURE, 1620 KPA (235 PSIA): INLET TEMPERATURE, 717 K (830 OF): TOTAL AIRFLOW, 3.63 KG/SEC (8 LB/SEC).

14

y ;;;SF' I RAT 1 ON

METAL /

FIGURE 9. - COMPLIANT NETALKERAMIC LINER CON- CEPT FOR A 3000 % COMBUSTOR EXIT TEMPERATURE.

15

- - -

800 OF AIR 315 PSIA

0.15 TO 0.30 LB/SEC

PER QUADRANT 320 P S I A 1 0.75 LB M/sEc/RAIL

, -

- I

I -

- -

FIGURE 10. - CERAMIC MATRIX LINER TEST RIG LAYOUT. KEY TEMPERATURE, FLOW, AND PRESSURE CONDITIONS, UNDER DESIGN CONDITIONS. ARE INDICATED.

16

N81ma1 NASn Aeron8ulsr ana Report Documentation Page Sp8cr Adm8nts1r8tlon

2. Government Accession No I 1. Report No. TH-100107 AIAA-87-1828; AVScOn TR-87-C-8

3. Recipient's Catalog No

I I

4 Title and Subtitle 5. Report Date

Liner Cooling Research at NASA Lewis Research Center 6. Performing Organization Code

535-05-01 7. Author@) 8. Performing Organization Report No.

Waldo A. Acosta E-3647 10. Work Unit No.

9. Performing Organization Name and Address

National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135

11. Contract or Grant No.

2. Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration Washington, D.C. 20546

14. Sponsoring Agency Code I I

5. Supplementary Notes

Prepared for the 23rd Joint Propulsion Conference cosposonsored by the AIAA, SAE, ASME, and ASEE, San Dlego, California, June 29 - July 2, 1987.

6. Abstract

This paper describes recently completed and current advanced liner research applicable to advanced small gas turbine engines. lution of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently underway at NASA Lewis Research Center. maintained while different advanced liner wall cooling techniques were investi- gated and compared to a baseline combustor. featuring counter-flow film-cooled llner (CFFC) panels, transpiration cooled liner walls (TRANS), and compliant metal/ceramic (CMC) walls was obtained over a range of simulated flight conditions of a 16:l pressure ratio gas turbine engine and fuel air ratios up to 0.034. injection system, identical geometric configuration outline, a.nd similar designed internal aerothermodynamics.

Research relating to the evo-

As part of this research, a reverse-flow combustor geometry was

The performance of the combustors

All the combustors featured an identical fuel

7. Key Words (Suggested by Author(s)) 18. Distribution Statement

Gas turbine combustor; Liner cooling; Unclassified - unlimited Compliant layer; Ceramic combustor; Composite matrix

STAR Category 07

20. Security Classif. (of this page) 21. No. of pages 22. Price' 3. Security Classif. (of this report)

Unc lass 1 f i ed Unclassified 17 A02 ASA FORM 1626 OCT 86 .

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