NASA Technical Memorandum 88967 AIAA-87-1788

Full-scale Thrust Reverser Testing in an Altitude Facility

(NASA-IM-88967) EULL-SCALE IBEiE59 BEVEflSEB N87-1€53 5 ' I E S T l h G IN A i i A L l l T U D E EAClLllY ( h A S A ) 24 p

CSCL l 4 B Unclas

G3/09 43643

Charles M. Mehalic and Roy A. Lottig 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

FULL-SCALE THRUST REVERSER TESTING I N A N ALTITUDE FACILITY

Char les M. Mehal ic and Roy A. L o t t i g N a t i o n a l Aeronaut ics and Space A d m i n i s t r a t i o n

Lewis Research Center Cleveland, Ohio 44135

In m * m

I u

A b s t r a c t

A two-dimensional convergent-d ivergent exhaust n o z z l e designed and f a b r i c a t e d by P r a t t and Whitney A i r c r a f t was i n s t a l l e d on a PW1128 tu rbo fan engine and t e s t e d d u r i n g t h r u s t r e v e r s e r o p e r a t i n n i n an a l t i t u d e f a c i l i t y a t NASA Lewis Research Center. A unique c o l l e c t i o n system was used t o c a p t u r e t h e t h r u s t r e v e r s e r exhaust gas and t r a n s p o r t i t t o t h e p r i m a r y exhaust c o l l e c t o r . Tes ts were conducted a t t h r e e f l i g h t c o n d i t i o n s w i t h v a r y i n g amounts o f t h r u s t reve rse a t each c o n d i t i o n . Some r e v e r s e r exhaust gas s p i l l a g e by t h e c o l l e c t i o n system was encountered b u t engine performance was una f fec ted a t a l l f l i g h t c o n d i t i o n s tes ted . r e s u l t s o f t h i s t e s t program, t h e f e a s i b i l i t y o f a l t i t u d e t e s t i n g o f advanced m u l t i - f u n c t i o n exhaust n o z z l e systems has been demonstrated.

Based on t h e

I n t r o d u c t i o n

Opera t i ona l requi rements o f t h e n e x t genera- t i o n o f m i l i t a r y a i r c r a f t w i l l d i c t a t e t h e use o f advanced m u l t i - f u n c t i o n exhaust n o z z l e systems capable o f t h r u s t v e c t o r i n g and reve rs ing . ous ly , t h r u s t v e c t o r i n g p rov ided power a s s i s t d u r i n g s h o r t o r v e r t i c a l t a k e o f f s w h i l e t h r u s t r e v e r s i n g was used f o r ground r o l l r e d u c t i o n . The advanced n o z z l e systems c u r r e n t l y under considera- t i o n w i l l be used t o enhance a i r c r a f t performance b y combining a i r c r a f t aerodynamic c o n t r o l wi th nozz le p r o p u l s i v e c o n t r o l t o p r o v i d e increased i n f l i g h t maneuverab i l i t y . p r e v i o u s a p p l i c a t i o n s o f t h r u s t v e c t o r i n g and r e v e r s i n g o n l y r e q u i r e d sea- level t e s t stands, advanced n o z z l e systems w i l l r e q u i r e t e s t i n g under s imu la ted a1 t i t u d e cond i t i ons .

have been dev ised and f a b r i c a t e d . Y Many o f t h e concepts i n c l u d e separate exhaust streams f o r vec- t o r i n g and r e v e r s i n g . exhaust may be handled i n an a l t i t u d e f a c i l i t y i n a manner s i m i l a r t o p rev ious t e s t s w i t h axisym- m e t r i c nozz les, t h e requi rement t o c o l l e c t t h e r e v e r s e r exhaust p resen ts a unique problem. The supersonic r e v e r s e r exhaust gases must be t u r n e d and r o u t e d t o t h e p r i m a r y exhaust gas c o l l e c t i o n system w i t h o u t c r e a t i n g adverse e f f e c t s on engine performance o r t h e t e s t c e l l environment. Recently, compu ta t i ona l f l u i d dynamics has been a p p l i e d t o the problem o f des ign ing exhaust gas c o l l e c t i o n systems f o r a l t i t u d e t e s t i n g o f v e c t o r i n g l r e v e r s i n g exhaust n o z z l e system^.^-^ r e v e r s e r exhaust gas c o l l e c t o r duc ts w i t h o u t i n t e r - n a l vanes would be inadequate t o p r o v i d e s a t i s f a c - t o r y c o l l e c t o r system e j e c t o r performance because o f impingement on t h e o u t e r w a l l w i t h r e s u l t i n g f l o w r e v e r s a l and s p i l l a g e a t t h e r e v e r s e r p o r t / c o l l e c t o r d u c t i n t e r f a c e . t i o n system used f o r t h e t e s t s desc r ibed i n t h i s paper i n c l u d e d t u r n i n g vanes i n t h e c o l l e c t o r ducts t o t u r n t h e f l o w and p reven t reve rse f l o w caused b y impingement.

P rev i -

Whi le t e s t i n g f o r t h e

Severa l concepts f o r nhlti-f n c t i o n nozz les

Whi le t h e vec to red t h r u s t

These s t u d i e s i n d i c a t e d t h a t

The exhaust gas c o l l e c -

T h i s paper desc r ibes t h e hardware and d i s - cusses t h e f u l l - s c a l e t e s t s o f a two-dimensional convergent-d ivergent (2D/CD) exhaust nozz le d u r i n g t h r u s t r e v e r s e r o p e r a t i o n i n an a l t i t u d e f a c i l i t y a t NASA Lewis. These t e s t s were conducted as p a r t o f a j o i n t N A S A / A i r F o r c e / P r a t t and Whitney program t o t e s t a 2D/CD nozz le i n an a l t i t u d e f a c i l i t y and rep resen t t h e f i r s t t i m e t h a t f u l l - s c a l e e n g i n e l t h r u s t r e v e r s e r o p e r a t i o n was at tempted i n an a l t i t u d e f a c i l i t y . F o r t h e t e s t i n g , t h e nozzle, a PW1128 t u r b o f a n engine, and a unique system designed and f a b r i c a t e d s p e c i f i c a l l y f o r c o l l e c t i n g t h e r e v e r s e r exhaust gas were i n s t a l l e d i n t h e t e s t c e l l . a t t h r e e f l i g h t c o n d i t i o n s . A t each f l i g h t cond i - t i o n t h e amount o f t h r u s t reverse, de f i ned as per- cen t o f t h r u s t r e v e r s e r p o r t open area, was v a r i e d f rom 0 t o 100 pe rcen t a t seve ra l enqine power se t - t i n g s r a n g i n g f rom i d l e t o maximum nona f te rbu rn ing power. Resu l t s a r e presented t o document t h e oer- formance o f t h e r e v e r s e r exhaust gas c o l l e c t i o n system and v e r i f y t h e f e a s i b i l i t y o f pe r fo rm ing t h i s t y p e o f t e s t i n g i n a l t i t u d e f a c i l i t i e s .

T h r u s t r e v e r s e r o p e r a t i o n was demonstrated

A l t

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N

P

T

W

6

e

S u b s c r i p t s

a

9

no r e v e r s e

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s p i 11 age

t

1

Nomenclature

a l t i t u d e , m

t h r u s t , N

f a n t u r b i n e i n l e t temperature, K

r o t a t i o n a l speed, rpm

pressure, N/m2

temperature, K

f l o w r a t e , k g l s e c

r a t i o o f t o t a l pressure t o abso lu te p ressu re o f NACA s tandard sea- level c o n d i t i o n s

r a t i o o f t o t a l temperature t o abso lu te temperature o f NACA s tandard sea- level c o n d i t i o n s

a i r

gross

w i t h o u t t h r u s t r e v e r s e r o p e r a t i o n

s t a t i c c o n d i t i o n s

c o n d i t i o n o f r e v e r s e r exhaust gas s p i l l a g e i n t o t e s t c e l l

t o t a l c o n d i t i o n s

a i r f l o w measur ing s t a t i o n

1

2 fan inlet .

25H compressor i n l e t

25M fan discharge, mixed

3 compressor e x i t

6M low pressure turbine e x i t , mixed

Apparatus

F aci 1 i t y

The t e s t program was conducted in one of t h e two a l t i t u d e chambers o f the Propulsion Systems Laboratory (PSL) a t NASA Lewis. Each t e s t chamber i s 7 . 3 m in diameter a n d 11.9 m long. Conditioned a i r a t t h e proper pressure and temperature t o simu- l a t e the desired f l i g h t condition i s supplied by ducting through the forward bulkhead t o t h e engine i n l e t while t h e desired a l t i t u d e pressure i s main- ta ined i n t h e t e s t chamber.

A photograph of the t e s t i n s t a l l a t i o n i s shown i n Fig. 1. The engine was mounted on a multi-axis thrust stand with a f lex ib le seal between t h e i n l e t a i r ducting and the engine t o provide accurate thrust measurement while accounting f o r engine growth and movement. The e f f e c t i v e area of t h e seal used t o c o r r e c t the measured gross thrust was determined from pre-test thrust stand ca l ibra t ions . A gap of approximately 1.9 cm was maintained between t h e nozzle reverser por t s and the f ixed reverser c o l l e c t o r ducts t o attempt t o avoid any e f f e c t s on t h e thrust measurement.

Engine

The engine used for the thrust reverser tes ts with t h e 2D/CD nozzle was a P r a t t and Whitney Air- c r a f t PW1128 low-bypass, high-compression r a t i o twin-spool afterburning turbofan. This engine c o n s i s t s of a three-stage fan driven by a two-stage t u r b i n e , ten-stage compressor driven by a two-stage air-cooled turb ine , an annular combustor and a mixed flow augmentor. T h e augmentor was not used during t h e t h r u s t reverser t e s t i n g . instrumentation was instal led in t h e engine t o measure s t a t i c pressure, tot21 pressure and t o t a l temperature a t the engine i n l e t , fan discharge, compressor i n l e t and ex i t , and fan turb ine i n l e t and e x i t t o monitor engine performance during t h e t e s t s .

Research

2D/CD Nozzle

The 2D/CD nozzle has a c i r c u l a r i n l e t sect ion t h a t t r a n s i t i o n s t o a rectangular c ross sect ion nozzle with r i g i d sidewalls and moveable upper and lower horizontal f laps t h a t control t h r o a t area, e x i t a rea , and thrust d i rec t ion . The f l a p posi- t i o n s f o r a x i a l , par t ia l reverse, and f u l l reverse thrust operation a r e shown i n Fig. 2. The upper and lower convergent f laps pivot about r i g i d s h a f t supports and determine the convergent exhaust area i n t h e ax ia l thrust mode and t h e percent reverse thrust i n t h e reverse thrus t mode. These f l a p s a r e synchronized with external connecting l inks . A S t h e convergent exhaust area i s reduced t h e reverser por t s open t o maintain t o t a l discharge area approx- imately constant .

The divergent f l a p s a r e attached t o t h e con- vergent f l a p s and s e t area expansion r a t i o and thrust vectoring u p t o a maximum of 20' from t h e horizontal . I n the ax ia l thrust mode the upper and lower divergent f l a p s a r e positioned symmetrically; i n t h e vectored thrust mode one f l a p a c t s a s a d e f l e c t o r t o vector thrust.

The nozzle was actuated hydraul ical ly and had a control system t o provide manual control o f t h e thrust reverser posi t ion a t engine power s e t t i n g s u p t o intermediate. Thrust reverse with af terburn- ing was not permitted. allowed manual s e t t i n g of vector angle and nozzle area expansion r a t i o . During normal forward thrust operation t h e nozzle control scheduled throa t a rea as requested by t h e Digi ta l Electronic Enqine Con- t r o l ( D E E C ) . A more de ta i led descr ipt ion of t h e 2D/CD nozzle design and i t s operation during sea- level t e s t s i s contained in Ref. 4.

This control system a l s o

Reverser Exhaust Collection System

The reverser exhaust qas co l lec t ion system consjs ted of two uncooled rectanqular c ross sec t ion ducts constructed of 0.48 cm s t a i n l e s s s t e e l t o capture the reverser exhaust and t ranspor t i t t o t h e primary exhaust c o l l e c t o r , support hardware f o r t h e ducts , and a conical adapter sect ion t o t h e primary exhaust co l lec tor . The i n s t a l l a t i o n of t h i s hardware i n t h e t e s t c e l l i s shown i n Fig. 3. The support s t r u c t u r e held t h e reverser c o l l e c t o r ducts over the reverser por t s so t h e r e was no con- t a c t between t h e nozzle and ducts ( a qap of approx- imately 1.9 cm was maintained a t t h e nozzle/duct i n t e r f a c e ) . ing seal between the ducts and a cover p l a t e on t h e conical adapter (Fig. 4 ) was i n s t a l l e d t o e l imina te exhaust gas rec i rcu la t ion . conical adapter had t h r e e rectanqular openings t o accommodate the reverser exhaust ( top and bottom) and t h e nozzle exhaust during forward and p a r t i a l reverse thrust operation.

A schematic drawing of a reverser c o l l e c t o r duct i s shown i n Fiq. 5. A l imited amount of aero- dynamic instrumentation was i n s t a l l e d on the t u r n - inq vanes in te rna l t o the duct. Deta i l s of t h e instrumentation mounting a re shown i n Fiq. 6. I n addi t ion t o t h e instrumentation of Figs. 5 and 6, a i r temperatures were measured near t h e area of t h e qap between t h e ducts and reverser port openings on each s i d e of t h e i n s t a l l a t i o n t o determine i f exhaust qas s p i l l a q e occurred durinq reverser oper- a t ion . These thermocouples were mounted approxi- mately 20 cm away from t h e nozzle i n the plane of the reverser p o r t l c o l l e c t o r duct in te r face .

reverser c o l l e c t o r ducts was t o have them operate over a wide ranqe o f condi t ions without adversely a f fec t inq engine operation a r t e s t c e l l environ- ment. Also considered was ease of construct ion which accounts f o r the extensive use of c i r c u l a r a rc and s t r a i g h t sec t ions . Since a l imited amount of space was ava i lab le between t h e enqine center- l i n e and t h e t e s t c e l l f l o o r (1.8 m), i t was f e l t t h a t supersonic turning of t h e reverser exhaust could not be accomplished without compromisinq engine performance. Therefore, i t was decided t o reduce the Mach number of the exhaust qases below

A t the discharge of t h e ducts a s l i d -

The cover p l a t e on t h e

The primary considerat ion in t h e design of t h e

1.0 and t u r n t h e f l o w subson ica l l y . i n t h e area o f t h e r e v e r s e r p o r t s was p r e d i c t e d u s i n g NAP code, a two-dimensional E u l e r code t h a t hand les b o t h supersonic and subsonic f l o w and assumes no v iscous e f f e c t s . The des ign was based on a maximum expected r e v e r s e r exhaust f l o w o f 90.7 k g l s e c a t 811 K. f l o w f i e l d , t h e f i r s t row o f t u r n i n g vanes was p o s i t i o n e d approx imate ly a long a contour l i n e o f c o n s t a n t Mach number (2.0) w i t h t h e i n t e n t o f s e t t i n g up a s e r i e s o f shocks t o reduce t h e Mach number t o subsonic downstream o f these t u r n i n g vanes. The exhaust f l o w c o u l d t h e n be t u r n e d sub- s o n i c a l l y b y t h e n e x t two rows o f vanes and t r a n s - p o r t e d t o t h e p r i m a r y exhaust c o l l e c t o r . With t h e e x c e p t i o n o f t h e t h r e e rows o f t u r n i n g vanes, t h e remainder o f t h e i n t e r n a l d u c t s t r u c t u r e ( s t i f - feners , r i b s , and s t r e n g t h e n i n g r o d s ) was i n s t a l l e d t o s t r e n g t h e n t h e d u c t s t o opera te a t p ressure d i f f e r e n t i a l s up t o 172.4 kPa across t h e w a l l s . F i n a l l y , upon i n s t a l l a t i o n , t h e o u t s i d e o f b o t h d u c t s was covered w i t h i n s u l a t i o n t o reduce thermal g r a d i e n t s i n t h e d u c t s t r u c t u r e and min imize heat l o s s t o t h e t e s t c e l l d u r i n g t h r u s t r e v e r s e r o p e r a t i o n .

The f l o w f i e l d

Based on t h e a n a l y s i s o f the

P r oc ed u r e

T h r u s t r e v e r s e r o p e r a t i o n was c a r r i e d o u t a t t h r e e f l i g h t c o n d i t i o n s : 7315 ml0.8 Mach, and 9144 ml1.17 Mach. A t t h e 7315 m c o n d i t i o n , engine i n l e t and t e s t c e l l pres- sures were main ta ined a t t h e c o r r e c t va lues f o r a l t i t u d e / M a c h number s imu la t ion , b u t no a t tempt was made t o supp ly t h e c o r r e c t eng ine i n l e t tempera- t u r e ; ambient temperature (approx imate ly 292 K ) was used. F o r t h e o t h e r f l i g h t c o n d i t i o n s , t h e pres- sures and temperatures were those cor respond ing t o t h e a c t u a l f l i g h t c o n d i t i o n .

1 2 192 ml0.8 Mach,

A t each f l i g h t c o n d i t i o n , t h e amount o f t h r u s t reverse , d e f i n e d as percent o f t h r u s t r e v e r s e r p o r t open area, was v a r i e d f r o m 0 t o 100 p e r c e n t a t s e v e r a l eng ine power s e t t i n g s r a n g i n g f r o m i d l e t o i n t e r m e d i a t e (maximum n o n a f t e r b u r n i n g power). For each t e s t p o i n t , engine, nozz le, and t e s t c e l l con- d i t i o n s were e s t a b l i s h e d as requ i red , and t h e engine was a l lowed t o s t a b i l i z e a t a s teady-s ta te o p e r a t i n g c o n d i t i o n p r i o r t o r e c o r d i n g data. The d a t a were acqu i red b y r e c o r d i n g 1 0 consecut ive scans (1.5 sec i n t e r v a l between scans) and averag- i n g t h e i n d i v i d u a l measurements b e f o r e p e r f o r m i n g f i n a l c a l c u l a t i o n s . d a t a r a t h e r t h a n o n l y one ins tan taneous s e t o f data.

T h i s approach p r o v i d e d average

D u r i n g t h e t e s t s , CRT d i s p l a y s i n t h e c o n t r o l

A l l measured parameters and ca lcu-

I n a d d i t i o n , c r i t i c a l meas-

room were used t o m o n i t o r t h e t e s t c o n d i t i o n s , engine and n o z z l e performance, and c o n d i t i o n o f the t e s t hardware. l a t i o n s were a v a i l a b l e f o r m o n i t o r i n g and updated a t 1.5 sec i n t e r v a l s . urements were compared w i t h programmed l i m i t s and any o u t - o f - l i m i t c o n d i t i o n s were h i g h l i g h t e d on the d i s p l ays .

Discuss ion o f R e s u l t s

R e s u l t s f r o m t h e t h r u s t r e v e r s e r t e s t s a r e p resented t o show t h e performance o f t h e r e v e r s e r exhaust gas c o l l e c t i o n system and i t s a b i l i t y t o g a t h e r t h e f l o w and t r a n s p o r t i t t o t h e p r i m a r y

exhaust c o l l e c t o r . The e f f e c t o f t h e c o l l e c t i o n system on engine per formance d u r i n g t h r u s t r e v e r s e r o p e r a t i o n i s a l s o examined b y comparison o f s e l e c t e d engine parameters between 0 and 100 p e r c e n t reverse . t h r u s t r e v e r s e r t o reduce f o r w a r d a x i a l t h r u s t i s p resented f o r each f l i g h t c o n d i t i o n t e s t e d .

The e f f e c t i v e n e s s of t h e

Reverser C o l l e c t i o n System Performance

The r e v e r s e r exhaust c o l l e c t i o n system accom- p l i s h e d t h e t a s k o f c a p t u r i n g and t r a n s p o r t i n g t h e r e v e r s e r exhaust gas t o t h e p r i m a r y c o l l e c t o r q u i t e w e l l a t a l l engine power s e t t i n q s f o r r e v e r s e r s e t t i n g s up t o approx imate ly 60 percent . h i g h e r r e v e r s e r s e t t i n g s , exhaust gas s p i l l a g e occur red a t t h e n o z z l e l c o l l e c t o r i n t e r f a c e f o r a l l f l i g h t c o n d i t i o n s t e s t e d . t h e s p i l l a g e temperature r a t i o as a f u n c t i o n o f engine a i r f l o w f o r each t e s t c o n d i t i o n . age temperature r a t i o i s t h e average a i r tempera- t u r e measured by t h e thermocouples i n t h e p lane o f t h e gap between t h e n o z z l e r e v e r s e r p o r t s and c o l - l e c t o r d u c t s d i v i d e d by t h e temperature measured w i t h o u t t h r u s t r e v e r s e which was n o m i n a l l y 290 K. The amount o f r e v e r s e r f l o w was n o t measured b u t t h e inc rease o t s p i l l a g e temperature r a t i o may i n d i c a t e an inc rease o f exhaust qas s p i l l a g e as t h e amount o f t h r u s t r e v e r s e increases. opera t ions , t e s t c e l l temperatures a r e main ta ined a t o r below 340 K w i t h f a c i l i t y c o o l i n g a i r . F o r t h e r e v e r s e r t e s t s a c o n s t a n t t e s t c e l l c o o l i n g f l o w o f approx imate ly 23 k g l s e c was used and t h r u s t s tand c a l i b r a t i o n s were per formed i n an a t tempt t o account f o r d rag f o r c e s induced b y t h e c o o l i n g f l o w . To m a i n t a i n l o c a l t e s t c e l l temperatures below 340 K corresponds t o a maximum s p i l l a g e tem- p e r a t u r e r a t i o o f approx imate ly 1.2 i n F i g . 7. A t t h i s r a t i o cont.inuous r e v e r s e r o p e r a t i o n was pos- s i b l e a t 60 percent and below. Above 60 percent reverse, some damage t o i n s t r u m e n t a t i o n l i n e s was exper ienced on a l o c a l i z e d b a s i s near t h e n o z z l e / c o l l e c t o r d u c t i n t e r f a c e .

A t t h e

F i g u r e s 7 ( a ) t o ( c ) show

The s p i l l -

Dur ing normal

F i g u r e 8 presents an e s t i m a t e o f t h e f l o w f i e l d w i t h i n t h e r e v e r s e r d u c t s a t t h e 7315 ml Mach 0.8 c o n d i t i o n which i s r e p r e s e n t a t i v e of a l l t h e c o n d i t i o n s t e s t e d . Based on t h e l i m i t e d i n t e r - n a l r e v e r s e r c o l l e c t o r d u c t i n s t r u m e n t a t i o n which i n d i c a t e s l o c a l f l o w Mach numbers t h a t may d i f f e r f r o m t h e f r e e s t ream because o f t h e l o c a t i o n o f t h e s t a t i c pressures on t h e vanes, i t appears t h a t t h e r e v e r s e r f l o w (below 80 percent reverse) f o l l o w s t h e contour o f t h e w a l l w i t h a r e g i o n o f low f l o w as i n d i c a t e d . A t 80 p e r c e n t r e v e r s e and above t h e h i g h e s t Mach numbers e n t e r i n g t h e d u c t a r e a t t h e c e n t e r of t h e s t ream w i t h low v e l o c i t y ( h i g h pres- sure) f l o w near t h e f o r e and a f t p o r t i o n s of t h e d u c t ent rance. I t i s t h i s low v e l o c i t y , h i g h pres- sure exhaust gas, p r o b a b l y caused b y j e t impinge- ment on t h e i n t e r n a l s u r f a c e s o f t h e r e v e r s e r c o l l e c t o r duc ts t h a t causes t h e inc reased s p i l l a g e a t l a r g e r e v e r s e r s e t t i n g s . A t t h e h i g h r e v e r s e r s e t t i n g s , t h e t u r n i n g vanes may n o t be do ing an e f f e c t i v e j o b o f t u r n i n g t h e f l o w and impingement on t h e fo rward w a l l o f t h e d u c t may be c r e a t i n g an area o f back f low toward t h e d u c t i n l e t . I c c r e a s i n g t h e chord l e n g t h o f t h e t u r n i n g vanes may h e l p t o a l l e v i a t e t h e problem i n f u t u r e t e s t s . A d d i t i o n a l i n s t r u m e n t a t i o n i s a l s o r e q u i r e d t o more a c c u r a t e l y d e f i n e t h e f l o w f i e l d w i t h i n t h e ducts .

3

To subs tan t ia te the flow f i e l d est imate of Fig. 8, Fig. 9 indicates damage experienced by the strengthening rods internal t o the duct. staggered horizontal and ve r t i ca l c i r c u l a r cross sect ion rods were supported by welds a t t h e ends. As the t e s t i n g progressed, f a i l u r e of t he horizon- t a l rod welds caused by vibrat ion occurred as indicated in Fig. 9. Both the t o p and bottom ducts experienced s imi la r f a i l u r e pa t te rns indicat ing damage t o those horizontal rods submerged in the high flow along the outer duct wal ls while rod welds in the low flow region near t he inner walls did not f a i l . minimized in fu tu re co l lec tor duct designs by aero- dynamically shaping the strengthening rods o r pro- viding some type of support a t t he center of t he rods.

The

These f a i lu re s could possibly be

Engine Performance

As mentioned ea r l i e r , one of the prime con- s idera t ions in t h e design of t he co l lec t ion system was t o be able t o co l lec t the th rus t reverser exhaust gas without affect ing engine performance by crea t ing an unusual backpressure condition. Figures 10 t o 1 2 present comparisons of engine speed match, fan turbine i n l e t temperature, and fan and compressor operating l i nes without t h rus t reverse and a t 100 percent reverse. ind ica te t h a t over t h e range of f l i g h t condi t ions t e s t ed , t he engine performance was n o t affected by th rus t reverser operation or t he reverser exhaust gas co l lec t ion system.

The f igures

Measured Gross Thrust

Thrust measurements during reverser operation were made using a multi-axis t h rus t stand. The th rus t stand in s t a l l a t ion and c a l i b r a t i o n was sim- i l a r t o t ha t described in Ref. 5. Figure 13 shows measured gross thrust corrected t o engine i n l e t condi t ions as a function of nozzle pressure r a t i o f o r each f l i g h t condition tes ted . The e f fec t ive- ness of the th rus t reverser t o reduce axial gross th rus t i s readi ly apparent from the f igure . The la rges t reduction in gross thrust occurs in the s teps between 0 and 80 percent reverse with only a small axial t h rus t reduction between 80 and 100 percent reverse.

The performance of t he th rus t reverser a t 100 percent reverse i s shown in Fig. 14. The f ig - ure ind ica tes t h a t measured reverse th rus t s of 50 percent of intermediate forward th rus t were obtained over an engine power level range from 50 t o 100 percent. d a t a a t the higher power se t t i ngs ind ica tes t h a t reverse th rus t performance was unaffected by the reverser exhaust gas col lect ion system.

The good cor re la t ion of the

Concluding Remarks

The t e s t s of the 2 D / C D nozzle during th rus t reverser operation successfully demonstrated the

f e a s i b i l i t y of f u l l - s c a l e a l t i t u d e t e s t ing of advanced thrust vector ing/reversing exhaust nozzle systems and added s i g n i f i c a n t l y t o t he t e s t experi- ence of nonaxisymmetric nozzles. The hardware used t o c o l l e c t the reverser exhaust gas performed ade- quately over a wide range of engine operating con- d i t i ons even t h o u g h some reverser exhaust gas sp i l l age in to the t e s t c e l l was experienced. This hardware a l so successful ly s a t i s f i e d the c r i t e r i a t h a t engine performance remain unaffected during thrust reverser operation.

Based on t he r e s u l t s of these t e s t s , t h e following recommendations a re made f o r fu tu re t e s t ing of advanced vector inglreversing exhaust nozzle systems in s imi la r i n s t a l l a t i o n s :

1. Longer turning vanes should be used t o pro- vide more e f f ec t ive turning of t he flow and c r e a t e a more equally d is t r ibu ted flow f i e l d in the reverser c o l l e c t o r ducts.

2. Computational f l u i d dynamics should be used more extensively in the design of the reverser duct in te rna l s t ruc tu re along with vibrat ion analysis in an attempt t o minimize any damage t h a t may be caused by the flow through the duct.

used a t t he reverser por t lco l lec t ion system in t e r - face t o minimize exhaust g a s sp i l l age without compromising th rus t measurements.

4 . Additional instrumentation should be in s t a l l ed in the reverser c o l l e c t o r ducts t o b e t t e r def ine the actual flow f i e l d in the ducts.

3. Some type of seal arrangement should be

1.

2.

3.

4.

5.

Ref e renc es

Dusa, D.J., Speir , D.W. , and Dunbar , D.K. , "Multi-Functional Nozzles f o r Advanced Weapon Systems," Powered L i f t Systems Plus, SAE SP-555, SAE, Warrendale, PA, 1983, pp. 9-19.

Phares, W.J., Cooper, G . K . , Swafford, T.W., and Jones, R . R . , "Application of Computational Fluid Dynamics t o Test F a c i l i t y and Experiment Design," A I A A Paper 86-1733, June 1986.

Crane, S . N . , J r . , Jones, R . R . 111, and Maywald P.V. , "Exhaust Gas Management f o r Testing Vectoring and Reversing Turbine Engines,"

McLafferty, G . H . and Peterson, J.L., "Results of Tests of a Rectangular VectoringlReversing Nozzle on an F l O O Engine," A I A A Paper 83-1285, June 1983.

AEDC-TR-86-22 , Aug . 1986.

Wooten, W . H . , Blozy, J.T., Speir , D.W., and Lot t ig , R . A . , "Alt i tude Testing of a F l igh t Weight, Self-Cooled, 20 Thrust Vectoring Exhaust Nozzle," SAE Paper 841557, Oct. 1984.

ORiGlNAL PAGE IS OF POOR QLIALlTY)

FIGURE 1. - INSTALLATION OF ENGINE, NOZZLE AND REVERSER COLLECTOR SYSTEM I N TEST CELL.

5

--t-

HORIZONTAL THRUST

PARTIAL REVERSE THRUST

FULL REVERSE THRUST

FIGURE 2 . - SCHEMATIC OF 2D/CD NOZZLE OPERATION.

6

FIGURE 3. - REVERSER COLLECTOR SYSTEM INSTALLATION.

7

ORiGlNAL PAGE FS OF. POOR QUALITY

FIGURE 4. - SEAL BETWEEN REVERSER COLLECTOR DUCT AND CONICAL ADAPTER.

8

I

\ \ 0 0 0 0 0 I I

I 0 0 0

i

G 0 0

H L L 0

W LL

9

,r STIFFENER

TOTAL PRESSURE 7

I

- A I R TEMPERATURE THERMOCOUPLE

PRESSURE --

FIGURE 6 . - DETAIL OF TURNING VANE INSTRUMENTATION INSTALLATION.

10

2.2

2.0

1 . 8

1.6

1.4

1.2

1 . 0

20 25 30 35 (A) FLIGHT CONDITION: 12192 M.

0.8 MACH N U W R . 2.4 r-

W : 2.2 ’ W

W w

(B) FLIGHT CONDITION: 7315 H. 0.8 MACH NUMBER.

2.4 r

2.2 I 2.0

1 . 8

1.6

1.4

1.2

1 .o

20 30 40 50 60 70 80 ENGINE AIRFLOW. W A , l , KG/SEC

(C) FLIGHT CONDITION: 9144 M. 1.17 MACH NUMBER.

FIGURE 7. - EFFECT OF REVERSER OPERATION ON TEST CELL TEMPERATURES AS MEASURED

VERSER PORT/COLLECTOR DUCT INTERFACE . BY THERMOCOUPLES I N THE PLANE OF THE RE-

11

12

n

I- v) W

I- V a n PI 0 I- V W -I -I 0 V

PI W v) w

W PI

I

00

9

0 0 I I n -1

Y

\ L.

a a m \

13

10 000

9000

8000

7000

REVERSE. PERCENT

0 0 0 100

6000 10 OOC 11 000 12 000 13 000 14 000

(A) ENGINE SPEED MATCH.

COMPRESSOR SPEED, N2, RPM

1000 / 1200 llooL 1000

900

800

700 1 .o 1.5 2 .0 2.5 3 .0

ENGINE PRESSURE RATIO, PT,6~/P, ,2

(B) FAN TURBINE INLET TEMPERATURE.

FIGURE 10. - EFFECT OF THRUST REVERSER OPERATION ON ENGINE PERFORMANCE: FLIGHT CONDITION - 12192 M, 0.8 MACH NUMBER.

14

3.6

3.2

2

E I-

n \

2.8 n

2 s I-

W PI 2 2.4 W U Lz r U LL

2.0

1.6 .40

8.0

I In

I- ?

x

2

n

I- n . 7.0

I- < U

w PI 3 v) v) w @= CL

6.0 v) v)

W PI CL aI 0 0

REVERSE. PERCENT

0 0 0 100

60 80 100 120 FAN CORRECTED AIRFLOW. W,,l $5/b2, KG/SEC

(C) FAN OPERATING LINE.

5.0 14 18 22 26 30 COMPRESSOR CORRECTED AIRFLOW, W,,25H 6 / 6 2 5 H , KG/SEC

(D) COMPRESSOR OPERATING LINE.

FIGURE 10. - CONCLUDED.

15

10 000

9000

8000

7000

6000

5000

REVERSE, PERCENT

0 0 0 100

4000 9000 10 000 11 000 12 000 13 000 14 000

COMPRESSOR SPEED, N2, RPM

1200

1100

1000

900

800

700

600 .5 1 .o 1.5 2 . 0 2.5 3.0

ENGINE PRESSURE RATIO, PT,6~/PT,2

(B) FAN TURBINE INLET TEMPERATURE.

FIGURE 11. - EFFECT OF THRUST REVERSER OPERATION ON ENG- INE PERFORMANCE: FLIGHT CONDITION - 7315 M, 0.8 MACH NUMBER.

3.2

2.8

c'!

f N

; 2.4 n

s c 4 Ec

W LT a 2.0

W CL n.

z 4 U

1.6

1.2

REVERSE. PERCENT

0 0

P /

7.0

8 VI

c'1

M,

0

6 . 0

n

+ U CL

5.0 a v, 0 w LT n

8 v, v, W

4.0

3

60 80 100 120

FAN CORRECTED AIRFLOW, w ~ , ~ fi2/f$. KG/SEC

(C) FAN OPERATING L I N E .

3.0 10 14 18 22 26 30

COMPRESSOR CORRECTED AIRFLOW, MA,25H &/625H, KG/SEC

(D) COMPRESSOR OPERATING L I N E .

FIGURE 11. - CONCLUDED.

17

10 000

9000

8000

f a - r

n W W LL v,

L 4 U

Y

t C LL

W c* 3 c

W z B c C w -I z

W

m cc 3 c

L 4. U

- E

7000

6000

5000

REVERSE. PERCENT

0 0 100

4OOO 9000 10 OOO 11 000 12 000 13 000

COMPRESSOR SPEED. Np, RPR

( A ) ENGINE SPEED MATCH.

1200

1100

1000

500 .5 1.0 1.5 2 . 0 2.5 3.0

ENGINE PRESSURE RATIO. P,,6M/P,,7

(5) FAN TURBINE INLET TEMPERATURE.

FIGURE 12. - EFFECT OF THRUST REVERSER OPERATION ON ENGINE PERFORMANCE: FLIGHT CONDITION - 9144 M. 1.17 MACH NUMBER.

18

3 . 6

3 .2

”! n \ 5 2 . 9 N

+ n

0 5 2 .4 a W CL 3 v) In W

E 2 . t z 4 LL

1 .(

1 .:

8.0

REVERSE, PERCENT

0 0 0 loo

S In N

W CL 3 v) v) w g 5 . 0

5: CL

cn W a

4.0 8

(C) FAN OPERATING LINE.

14 18 22 26 30 3 . 0 ’ 10 COMPRESSOR CORRECTED AIRFLOW, w,,, 25H G H / b 2 5 H . KG/SEC

(D) COMPRESSOR OPERATING LINE. FIGURE 12. - CONCLUDED.

19

80 000

60 OOO

40 000

20 OOO

0

-20 OOO

REVERSE, PERCENT

0 0 A 2 0 0 40 D 60 V 80 0 1 0 0

(A) FLIGHT CONDITION: 1 2 1 9 2 M. 0.8 MACH NUMBER. 80 000

60 000

2 20 000 P W n W

W CL

V

L C

z - 2 0 ooc

-40 OM

80 000

60 000

40 000

20 000

a

- 2 0 ooc

-40 ooc

( B ) FLIGHT CONDITION: 7 3 1 5 M, 0.8 MACH NUMBER

P

2 3 4 5 6 NOZZLE PRESSURE RATIO

( C ) FLIGHT CONDITION: 9 1 4 4 M. 1 . 1 7 MACH NUMBER.

FIGURE 1 3 . - EFFECT ff THRUST REVERSER OPERATION ON MEASURED GROSS THRUST.

20

L . ‘ O r in =3 oi I I-

O

ALT, MACH

0 12192 0.8 0 7315 0.8

M NUMBER

20

0 20 40 60 80 100 PERCENT OF INTERMEDIATE THRUST

FIGURE 14. - REVERSE THRUST PERFORMANCE.

21

1 Report NO NASA TM-8896.1 A I A A - 87-1788

~

2 Government Accession No 3 Recipient's Catalog No

2. Sponsoring Agency Name and Address I Techn ica l Memorandum

4 Title and Subtitle

F u l l - s c a l e T h r u s t Reverser T e s t i n g i n an A l t i t u d e F a c i l i t y

7 Author($

Char les M. Mehal ic and Roy A . L o t t i g

5 Report Date

6 Performing Organization Code

505-62-30 8 Performing Organization Report No

E-3435

5. Supplementary Notes

Prepared f o r t h e 23rd J o i n t P r o p u l s i o n Conference, cosponsored by t h e A I A A , SAE, A M € , and ASEE, San Diego, C a l i f o r n i a , June 29 - J u l y 2, 1987.

9. Performing Organization Name and Address

N a t i o n a l Aeronaut ics and Space A d m i n i s t r a t i o n Lewis Research Center Cleveland, Ohio 44135

6 Abstract

A two-d imensional convergent-dfvergent exhaust n o z z l e designed and f a b r i c a t e d by P r a t t and Whitney A i r c r a f t was i n s t a l l e d on a PW1128 t u r b o f a n engine and t e s t e d d u r i n g t h r u s t reve rse r o p e r a t i o n i n an a l t i t u d e f a c i l i t y a t NASA Lewis Research Center. A unique c o l l e c t i o n system was used t o c a p t u r e t h e t h r u s t r e v e r s e r exhaust gas and t r a n s p o r t j t t o t h e p r imary exhaust c o l l e c t o r . Tests were con- ducted a t t h r e e f l i g h t c o n d i t i o n s w i t h v a r y i n g amounts o f t h r u s t r e v e r s e a t each c o n d i t i o n . Some reverser exhaust gas s p i l l a g e by t h e c o l l e c t i o n system was encountered b u t engine performance was u n a f f e c t e d a t a l l f l i g h t c o n d i t i o n s t e s t e d . Based on the r e s u l t s o f t h i s t e s t program, t h e f e a s i b i l i t y o f a l t i t u d e t e s t i n g o f advanced m u l t i - f u n c t i o n exhaust n o z z l e systems has been demonstrated.

10. Work Unit No

11. Contract or Grant No.

13. Type of Report and Period Covered

N a t i o n a l Aeronaut ics and Space A d m i n i s t r a t i o n Washington, D.C. 20546

Th rus t r e v e r s i n g ; l u r b i n e engine t e s t i n g ; Exhaust gas c o l l e c t o r ; Exhaust nozz les

14. Sponsoring Agency Code

U n c l a s s i f i e d - u n l i m i t e d S l A R Category 09

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

'For sale by the National Technical Inlormalion Service. Springfield, Virginia 221 61

9 Security Classif (of this report)

U n c l a s s i f i e d 20 Security Classif (of this page) 21 NO of pages 22 Price'

U n c l a s s i f i e d 23 A02