Read A subsonic to Mach 5.5 subscale engine test facility text version

AIA A-87-2052 A Subsonic to Mach 5.5 Subscale Engine Test Facility

Earl H. Andrews, Jr., NASA Langley Research Center, Hampton, VA

AIANSAEIASMEIASEE 23rd Joint Propulsion Conference

June 29-July 2, 1987/San Diego, California

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NASA-Langley's involvement i n a i r b r e a t h i n g h y p e r s o n i c engine r e s e a r c h and development began i n t h e e a r l y 1960'9. This r e s e a r c h was focused around t h e Hypersonic Research Engine (HRE), a n a x i s m m e t r i c r a m j e t / s c r a m j e t t h a t was designed t o b e pod-mounted t o an X-15 f o r f l i g h t t e s t s up t o Mach 8. The x-15 program was terminated b e f o r e t h e f l i g h t tests could be conducted, b u t t w o eng i n e models were f a b r i c a t e d and t e s t e d i n ground t e s t f a c i l i t i e s as shown i n F i g u r e 1.1-3 These models were used t o s u c c e s s f u l l y demonstrate t h e r a m j e t / s c r a m j e t c y c l e s with v a r i a b l e engine geomet r y ( a t r a n s l a t i n g i n l e t spike1 and t o i n i t i a t e development of r e g e n e r a t i v e l y - c o o l e d f l i g h t hardware. A f t e r completion of t h e H E program, t h e R Hypersonic P r o p u l s i o n Branch IHPB) a t NASA-Langley i n i t i a t e d a program to s t u d y r a m j e t / s c r a m j e t prop u l s i o n systems t h a t were h i g h l y i n t e g r a t e d with the v e h i c l e t o o b t a i n a more e f f i c i e n t p r o p u l s i o n system f o r h y p e r s o n i c speeds which was designed f o r minimum d r a g and weight and t o p r o c e s s as much a i r as p o s s i b l e . This airframe-integrated propuls i o n system enconpasses t h e e n t i r e u n d e r s i d e of t h e v e h i c l e (Fig. 2 ) . The a i r c r a f t forebody i s used f o r part of t h e i n l e t compression p r o c e s s and t h e a f t b o d y f o r p a r t of t h e n o z z l e expansion p m cess. The p r o p u l s i o n system i n c l u d e s a c l u s t e r of i d e n t i c a l i n d i v i d u a l r e c t a n g u l a r engine modules of a s i z e and shape s u i t a b l e f o r ground t e s t i n g . An i n t e r n a l view Of one module o f the Langley f i r s t g e n e r a t i o n a i r f r a m e - i n t e g r a t e d s c r a m j e t is shown A c o n s i d e r a b l e amount of i n l e t and i n Figure 2. combustor component r e s e a r c h r e l a t e d to t h i s airf r a m e - i n t e g r a t e d engine concept was performed w i t h i n t h e H B i n t h e 1 9 6 0 - 8 0 ' ~ . ~ In t h e mid P 1970's two e x i s t i n g f a c i l i t i e s a t Langley were modified to a l l o w tests Of s u b s c a l e , componenti n t e g r a t i o n engine models a t s i m u l a t e d Mach 4 and 7 f l i g h t conditions. Four d i f f e r e n t engine conc e p t s have s i n c e been t e s t e d , and r e s u l t s from t h r e e of t h e s e engine tests are summarized i n Reference 5. The Mach 4 e n g i n e test f a c i l i t y c o n s i s t e d mostly of s u r p l u s hardware assembled i n an existing test c e l l . The t e s t gas was h e a t e d by aie/hydrogen combustion w i t h oxygen replenishment t o a l l o w p r o p u l s i o n t e s t i n g . The p r e s e n t f a c i l i t y h e a t e r h a s a p r e s s u r e l i m i t of 190 p s i a and a temperature l i m i t of 2250% because of S a f e t y considerations. A new h e a t e r w i l l i n c r e a s e these l i m i t s t o 600 p s i a and 3000% and a new f a c i l i t y n o z z l e and o t h e r o p t i o n a l d u c t i n g hardware have g r e a t l y extended t h e o p e r a t i n g range of t h e fac i l i t y . P l a n s are a l s o underway to i n c o r p o r a t e a 70-foot vacuum s p h e r e i n t o t h e f a c i l i t y system. In t h i s p a p e r , t h e o r i g i n a l and t h e modified f e a t u r e s of the f a c i l i t y are d e s c r i b e d and t h e extended o p e r a t i n g range of t h e f a c i l i t y i s documented. I n a d d i t i o n , S p e c i a l f a c i l i t y problems a s s o c i a t e d w i t h scramjet e n g i n e t e s t i n g are d i s cussed and t h e r e l a t i o n s h i p Of this f a c i l i t y t o t h e e n t i r e Langley Scramjet Engine T e s t Complex is described.

The Hypersonic Propulsion Branch a t NASALangley h a s focused i t s r e s e a r c h on t h e supersonic-combustion r a m j e t ( s c r a m j e t ) a i r f r a m e i n t e g r a t e d p r o p u l s i o n system concept s i n c e about 1969. Engine component ( i n l e t and cornbustor) = = s e a r c h was performed and documented i n t h e e a r l y 1970's. During t h i s same p e r i o d , two e x i s t i n g ground f a c i l i t i e s were modified t o a l l o w tests of s u b s c a l e , component-integration engine models a t s i m u l a t e d Mach 4 and 7 f l i g h t c o n d i t i o n s . The Mach 4 f a c i l i t y , i n which air-hydrogen combustion i s used t o h e a t t h e test gas, h a s been r e c e n t l y modified t o make i t s u i t a b l e for tests of r a m j e t t y p e models a t s u b s o n i c and t r a n s o n i c Mach numbers as w e l l as t h e p r e s e n t s u p e r s o n i c Mach number capability. These m o d i f i c a t i o n s are d e s c r i b e d , and the extended test c a p a b i l i t i e s of t h e f a c i l i t y are documented i n t h i s paper.

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n o z z l e exit e q u i v a l e n t diameter 114.97), i n . h e i g h t of s q u a r e nozzle, i n . free-stream t o t a l e n t h a l p y , BTU/lb. Mach number behind v e h i c l e bow shock or a t f a c i l i t y n o z z l e e x i t f a c i l i t y a i r e j e c t o r e x i t Mach

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f r e e - s t r e a m f l i g h t Mach number s t a t i c pressure, p s i a f a c i l i t y t e s t cabin pressure, p s i a f a c i l i t y nozzle e x i t w a l l pressure, psia t o t a l p r e s s u r e behind a normal shock ( t o t a l pressure a t Mex < 1.0). p s i a t o t a l p r e s s u r e behind v e h i c l e bow shock; f a c i l i t y n o z z l e e x i t , p s i a f a c i l i t y burner t o t a l pressure, p s i a free-stream f l i g h t t o t a l p r e s s . , psia f r e e - s t r e a m f l i g h t dynamic p r e s s u r e , psf f a c i l i t y nozzle e x i t radius, i n . f a c i l i t y b u r n e r or n o z z l e e x i t (= T ) , 'R t o t a l temperature free-stream f l i g h t t0t.d'" temperature, R ' d i s t a n c e i n f a c i l i t y exhaust d u c t (see f i g . 1 3 ) , i n . v e r t i c a l d i s t a n c e from f a c i l i t y nozzle c e n t e r l i n e , in. r a t i o of s p e c i f i c h e a t v e h i c l e precompeession a n g l e , deg. f l o w rate, l b s / s e c .

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B a s i c -roach BTtenSive i n l e t and combustor r e s e a r c h programs, both a n a l y t i c and experimental, have been and are being conducted a t LaRC i n s u p p o r t of t h e airframe-integrated scramjet ~ o n c e p t . ~ n l e t I and conbustor d e s i g n s generated i n this component r e s e a r c h are then a8sembled t o form componenti n t e g f a t i o n e n g i n e models. These Subscale engine models are t e s t e d i n ground f a c i l i t i e s t o d e f i n e and r e s o l v e i n t e r a c t i o n s between t h e v a r i o u s eng i n e components and t o determine t h e o v e r a l l eng i n e performance. The mcdels are of t h e heat-sink t y p e and are g e n e r a l l y r e p r e s e n t a t i v e of one mod u l e of a c l u s t e r shown on t h e a i r c r a f t i n Fiqllre 2. The ground f a c i l i t y t e s t s mnst s i m u l a t e t h e f l i g h t c o n d i t i o n s of an a i r f r a m e - i n t e g r a t e d scramj e t as c l o s e l y as p o s s i b l e . The s i m u l a t i o n l o g i c i s d e p i c t e d i n Figure 3. A v e h i c l e f l y i n g a t s u p e r s o n i c or f a s t e r speeds compresses t h e flow across t h e forebody bow shock. A t a p a r t i c u l a r a l t i t u d e , i n c r e a s e d a i r c r a f t speed r e s u l t s i n i n c r e a s e d S t a g n a t i o n c o n d i t i o n s such as t o t a l p r e s s u r e and total e n t h a l p y (or t e m p e r a t u r e ) . The flow c o n d i t i o n s downstream of the bow shock are d i f f e r e n t i n t h a t the v e l o c i t y (Mach number) a n d t o t a l p r e s s u r e are decreased, t h e S t a t i c p r e s s u r e i s i n c r e a s e d , and the t o t a l e n t h a l p y or t o t a l temperature remains t h e same. I t is t h i s f l o w c o n d i t i o n j u s t ahead of t h e p r o p u l s i o n system, d e p i c t e d i n Figure 31a) by M , , t h a t has t o be d u p l i c a t e d i n an engine test f a c i l i t y .

and parameters; i.e., P, T, D, M. Facility t u r b u l e n c e l e v e l which is almost c e r t a i n l y an unknown i n engine t e s t s , v a r i e s with each f a c i l i t y and Could have an e f f e c t on f u e l - a i r mixing. Engine size ( s c a l i n g e f f e c t ) is very important i f combustion is n o t mixing-controlled s i n c e chemical k i n e t i c s do n o t s c a l e (mixing does s c a l e ) . Therefore, poor t h r u s t performance i n s u b s c a l e t e s t s due t o k i n e t i c s problems does not n e c e s s a r i l y mean t h a t a l a r g e r engine would have F a c i l i t y model t h e same poor performance. i n t e r a c t i o n s can occur which w i l l negate or make d a t a i n t e r p r e t a t i o n d i f f i c u l t ; marginal f a c i l i t y Although t h e d i f f u s e r s can cause t h i s problem. f a c t o r s Outlined above Warrents c a u t i o n i n i n t e r p r e t a t i o n of s u b s c a l e s c r a m j e t ground f a c i l i t y d a t a , t h e s e t e s t s are i m p o r t a n t and informative. Subscale s c r a m j e t ground f a c i l i t y t e s t s can be performed r e l a t i v e l y i n e x p e n s i v e l y , can y i e l d l a r g e q u a n t i t i e s of d a t a , and t h i s d a t a i s v a l u a b l e i n t h e s t u d y of combustor-inlet i n t e r a c t i o n ? f u e l i n j e c t o r s i z e , s p a c i n g , and s t a g i n g ; flameholding; and t h r u s t performance. Engine and Component Test Facilities

?he Hypersonic Propulsion Ezanch h a s assembled a group of f a c i l i t i e s t h a t e n a b l e s such s c r a m j e t r e s e a r c h t o be conducted a t t h e NASA-Langley Research Center. Photographs of t h e f a c i l i t i e s are shown i n Figure 4.

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The arrangement shown i n Figure 3 ( b ) schematic a l l y r e p r e s e n t s an engine t e s t f a c i l i t y t h a t produces t r u e - v e l o c i t y , true-temperature, and t r u e - p r e s s u r e flow f o r f l i g h t s i m u l a t i o n . F l i g h t f r e e s t r e a m t o t a l e n t h a l p y or t o t a l temperature is d u p l i c a t e d by a h e a t e r (with oxygen replenishment, as r e q u i r e d ) . A f a c i l i t y contoured nozzle exh a u s t s the heated, simulated-air flow t o the eng i n e module i n l e t a t a Mach number l M 1 ) t h a t sirnul a t e s t h e v e h i c l e forebody p r e c w p r e s s e d flow. The a l t i t u d e . or dynamic pressure, simulated by t h e nozzle e x i t flow i s dependent upon t h e h e a t e r t o t a l pressure. Also, t h e s c r a m j e t engine nodules a r e mounted i n t h e f a c i l i t y so t h a t a l l or a p r t i o n of t h e f a c i l i t y - n o z z l e top-surface boundary l a y e r i s i n g e s t e d by t h e model i n partial sirnulat i o n of t h e i n g e s t i o n of a v e h i c l e forebody hound a r y l a y e r by a f l i g h t engine.

In an arrangement Such as presented i n Figure 3 ( b ) , f l i g h t c o n d i t i o n s are d u p l i c a t e d as c l o s e l y as p o s s i b l e , i.e., t r u e t o t a l e n t h a l p y , s t a t i c p r e s s u r e , s t a t i c temperature, model scale, t e s t gas composition, t u r b u l e n c e l e v e l , e t c . Although t o t a l e n t h a l p y can be d u p l i c a t e d i n s u b s c a l e ground engine t e s t f a c i l i t i e s ! t h e o t h e r f a c t o r s mnSt be examined more c l o s e l y . Heating t h e t e s t gas t o t h e c o r r e c t t o t a l e n t h a l p y t o s i m u l a t e t h e f l i g h t Mach number u s u a l l y i s accompanied by flow contamination, i.e., HZO, NOX, e t c , t h a t may e f f e c t engine f u e l combustion. Fxpansion of t h e heated a i r i n t h e f a c i l i t y nozzle l e a v e s t h e v i b r a t i o n a l mode i n a nonequilibrium s t a t e and, above Tt,l = 4000"R. chemical nonequilibrium (oxygen d i s s o c i a t i o n f i r s t ) can occur. These a f f e c t nozzle e x i t flow p r o p e r t i e s

Small-scale i n l e t experimental t e s t a f o r "screening" p o t e n t i a l i n l e t d e s i g n s are p e r f o r r e d i n the Mach 4 Blowdown Tunnel t h a t uses unheated a i r and has a 9 v %inch test s e c t i o n . Much can be learned from s u c h tests such as t h e i n l e t starting capability, r e s t a r t capability, i n l e t air mass c a p t u r e , and s u r f a c e p r e s s u r e s t h a t can be compared to computational f l u i d dynamic (CfD) results. The models are g e n e r a l l y r e l a t i v e l y inexpensive with c o n s i d e r a b l e model geometric v e r s a t i l i t y . Larger-scale i n l e t tests are a l s o performed i n l a r g e r aerodynamic wind t u n n e l s a t Langley. Small-scale d i r e c t - c o n n e c t combustor tests t h a t s i m u l a t e a p o r t i o n of t h e engine combustor are Conducted i n Test C e l l # 2 Direct-Connect Comb u s t o r F a c i l i t y (Fig. 41 t o provide b a s i c r e s e a r c h d a t a on s u p e r s o n i c f u e l - a i r mixing, i g n i t i o n , and combustion p m c e s s e s . The h o t t e s t gas is s u p p l i e d t o t h e combustor models by a hydrogen-airoxygen combustion h e a t e r , to s i m u l a t e a i r w i t h e n t h a l p l e v e l s ranging up t o Mach 7 f l i g h t Speeds.' V a r i o u s f a c i l i t y nozzles produce t h e d e s i r e d combustor e n t r a n c e flow c o n d i t i o n s . Much can be l e a r n e d from i n d i v i d u a l component experimental b a s i c r e s e a r c h , but when t h e s e compon e n t s are i n t e g r a t e d i n t o an engine c o n f i g u r a t i o n a new set of problems may be encountered. Flow c o n d i t i o n s s u p p l i e d t o d i r e c t - c o n n e c t conbustor t e s t s are g e n e r a l l y Shock f r e e uniform flow condit i o n s . The flow d e l i v e r e d by an i n l e t a t t a c h e d t o a combustor e n t r a n c e i n an engine c o n f i g u r a t i o n i n c l u d e s r e f l e c t e d shocks and "on-uniformity. It was t h e r e f o r e d e s i r a b l e t o provide a t Langley t h e c a p a b i l i t y t o conduct s u b s c a l e engine tests t o perform i n l e t and combustor b a s i c r e s e a r c h i n an enqine environment. To f u l f u l l t h i s requirement, two engine test f a c l i t i e s have been assembled by t h e lfypersonic Propulsion Branch (Fig. 4 ) . These

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f a c i l i t i e s permit inexpensive, h i g h l y p r o d u c t i v e , combustion and engine r e s e a r c h tests t o be cond u c t e d on s m a l l - s c a l e , ga8eou5-fuel-bu~ning. s c r a m j e t models. One is an e l e c t r i c - a r c - h e a t e d f a c i l i t y w i t h 11-inch square e x i t contoured nozz l e s t h a t e x h a u s t s throuqh a d i f f u s e r i n t o a 100f t vacuum Another f a c i l i t y u t i l i z e s a hydrogen-burning h e a t e r t o provide a h o t t e s t gas with the proper mygen Content t h a t i s expanded through a 13.26-inch sguare e x i t contoured nozzle; an a i r e j e c t o r is employed t o a i d e x h a u s t i n g t o Both f a c i l i t i e s y i e l d f r e e - j e t t h e atmosphere.' t u n n e l flow f o r s u b s c a l e s c r a m j e t tests a t s i m u l a t e d Mach 7 and Mach 4 f l i g h t c o n d i t i o n s i n t h e arc-heated and combustion-heated f a c i l i t i e s , r e s p e c t i v e l y . 5 The same s i z e models ( f r o n t a l view a b o u t 6 by 8 inches and 72 inches longl are t e s t e d i n b o t h of t h e s e f a c i l i t i e s . The combustionheated f e c i l i t y i s the t o p i c of t h e remainder of t h i s paper. F a c i l i t y D e s c r i p t i o n and Discussion

e x i t of t h e d i f f u s e r i s pumped upon by an a n n u l a r a i r ejector (about ZOO pounds per second) with an e x i t Mach number of 4.16 (can be manually changed

These flows mix i n a 25-inch d i a t o M = 3.721. meter mixer d u c t t h a t is a b o u t 5.5-diameters long and is then turned v e r t i c a l l y by the t u r n i n g elbow. The flow e x h a u s t s f r m t h e elbow as a f r e e - j e t and e n t e r s a & f o o t diameter d u c t i n the t e s t c e l l c e i l i n g exhaust tower. T h i s r e l a t i o n s h i p r e s u l t s i n a low-pressure, high-volume flow e j e c t o r t h a t e n t r a i n s test cell a i r i n t o t h e v e r t i c a l exhaust d u c t . This e j e c t o r a c t i o n , i n t u r n , induces flow t o e n t e r through t h e test c e l l c e i l i n g e n t r a n c e tower and "wash" over t h e f a c i l i t y hardware and p i p i n g t o ensure p u r g i n g from t h e t e s t c e l l any gaseous p r o p e l l a n t leakage t h a t may occur ( f o u r hydrogen d e t e c t o r s are s t r a t e g i c a l l y l o c a t e d i n t h e test c e l l , t e s t c a b i n , and e x h a u s t tower).

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Description. The Combustion-Heated s c r a m j e t T e s t F a c i l i t y (CHSTF) b a s i c c o n f i g u r a t i o n is rep r e s e n t e d i n Figure 5.' T h i s f i g u r e has a schernat i c s k e t c h of t h e b a s i c c o n f i g u r a t i o n along with a photograph and some S a l i e n t f e a t u r e s of t h e f a cility. A l l of t h e f a c i l i t y hardware i s contained i n a 16 Y 16 x 52 f o o t test c e l l as shown i n t h e s k e t c h with t h e flow d i r e c t i o n from l e f t t o A i r i s s u p p l i e d by a c e n t r a l i z e d 600 p s i right. d i s t r i b u t i o n System. (The a i r supply i n t o t h e T e s t C e l l Complex and i t s c o n t r o l s are shared a l o n g w i t h t h e g a s w v a p r o p e l l a n t S u p p l i e s and c o n t r o l s with t h e T e s t C e l l U2 Direct-Connect Combustor F a c i l i t y . ] Oxygen is i n j e c t e d t o mix with t h e a i r t o y i e l d an oxygen-rich a i r mixture. Hydrogen is then i n j e c t e d i n t o t h i s oxygen-rich a i r mixture i n the f a c i l i t y h e a t e r and i g n i t e d with a hydrogen/oxygen t o r c h i g n i t e r . The f a c i l i t y hydrogen and oxygen flow r a t e s are cont r o l l e d so t h a t t h e r e s u l t i n g combustion product mixture c o n t a i n s approximately 21 p e r c e n t free oxygen by volume to s i m u l a t e t h e oxygen c o n t e n t of a i r . The remaining t e s t gas is a mixture of n i t r o g e n and water vapor; the h i g h e r the s t a g n a t i o n temperature f o r h i g h e r Simulated f l i g h t speed, t h e g r e a t e r t h e hydrogen flow r a t e r e q u i r e d f o r comb u s t i o n and t h u s t h e l a r g e r t h e water vapor cont e n t i n the test gas. The mass f r a c t i o n of the t e s t gas c o n s t i t u e n t s a t simulated f l i g h t Mach number t o t a l temperatures are presented i n Figure 6. Mach 4 f l i g h t s i m u l a t i o n ( T - 1640'R) ? r e s u l t s i n a nominal test gas composition of 6 p e r c e n t water, 70 p e r c e n t n i t r o g e n , and 24 p e r c e n t oxygen by mass; 9 , 70, and 21 p e r c e n t , r e s p e c t i v e l y by volume. Expansion of t h e test gas i s accomp l i s h e d through a Mach 3.5 contoured nozzle t h a t h a s s q u a r e cross s e c t i o n s and is 13.264 i n c h e s Square a t t h e e x i t . The Mach 3.5 f r e e - j e t exhaust flow s i m u l a t e s t h e forebody precompressed flow of a v e h i c l e a t Mach 4.0 f l i g h t c o n d i t i o n s as d i s c u s s e d f o r F i g u r e 3. A test c a b i n c o n t a i n s t h e f r e e - j e t exhaust flow t h a t is received by an exh a u s t c a t c h cone t o y i e l d a free j e t l e n g t h (from . n o z z l e e x i t to cone e n t r a n c e 1 of 15 nozzle e x i t e q u i v a l e n t d i a m e t e r s . The c a t c h cone is connected t o a 19-inch diameter s t r a i g h t d u c t s u p e r s o n i c e x h a u s t d i f f u s e r ( a b o u t a-diameters l o n g l . The t u n n e l flow (30 t o 60 pounds per Second1 a t t h e

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e x i t of t h e f a c i l i t y nozzle as discussed f o r Figure 3 ( b ) . Engine model gaseous hydrogen f u e l is s u p p l i e d , up to 700 psia, through s i x i n d i v i d u a l l y c o n t r o l l e d systems. These s i x systems can be routed t o d i f f e r e n t engine f u e l i n j e c t i o n stat i o n s t h a t can be used as d e e i r e d d u r i n g a test. One of t h e s e systems g e n e r a l l y c o n t r o l s t h e s u p p l y Of a phyorphoric gas used f o r engine f u e l i g n i tion. This same System can a l s o b e e a s i l y conv e r t e d t o supply and c o n t r o l a h i g h e r p r e s s u r e ( u p t o 1000 p s i a l gaseous hydrogen f u e l . A 1000 psi a i r supply ( 1 i n c h p i p e ) i s a l s o a v a i l a b l e i n the t e s t c e l l t h a t may be connected t o one of t h e s i x c o n t r o l systems. Ale0 f o r f u e l i g n i t i o n , two high-voltage e l e c t r i c tranSformer8 and t h e i r cont r o l s are a v a i l a b l e a$ s t a n d a r d test Options i n the t e s t cell. F a c i l i t y instrumentation. The f a c i l i t y was h e a v i l y instrumented d u r i n g c a l i h r a t i o n t-ot- ana d u r i n g i n i t i a l engine model tests. The main comp u t e r of t h e d a t a a c q u i s i t i o n System ( D A s ) i s a multiprograrnmable, 3 2 - b i t , g e n e r a l purpose d i g i t a l computer w i t h p a r a l l e l 500.000 word paged memory. The D S is designed for high-speed d a t a A a c q u i s i t i o n and i n c l u d e s a console t e r m i n a l , two d i s k d r i v e s (a 24 and a 48 megabyte s i z e ) , a Card r e a d e r , two 800 b i t s / i n . t a p e d r i v e s , and a l i n e printer. A g r a p h i c s t e r m i n a l a t t a c h e d t o the main computer over an RS-232 i n t e r f a c e perrnitS d e s i r e d d a t a t o be p l o t t e d a t a r a t e of 2400 b i t s / s e c . A s e r i e s of e l e c t r o n i c a m p l i f i e r s allows up t o 192 analog s i g n a l s t o be digitized. An e l e c t r o n i c a l l y Scanned p r e s s u r e (ESP) measurement System is a t tached and c o n t r o l l e d over a Standard IEEE-488 i n t e r f a c e . The ESP i s used f o r some of t h e f a c i l i t y d u c t p r e s s u r e measurements, b u t i t s main purpose is t o provide model pressure measurements up t o 75 p s i a . Model t h r u s t and d r a g i s measured w i t h a one-component s t r a i n - g a u g e f o r c e balance.

D a t a are recorded a t a scan r a t e of 10 frames p e r second and are g e n e r a l l y p r i n t e d a t a rate o f The 2 p e r second for a nominal 20-second r u n . p r i n t e d d a t a are i n e n g i n e e r i n g u n i t s , r a t i o s , and/or c a l c u l a t i o n s u s i n g raw d a t a ; p r i n t i n g Starts w i t h i n 1 second a f t e r t h e end of a test. Data are immediately p l o t t e d on t h e g r a p h i c s terminal from which a s t a n d a r d set of p l o t s is obt a i n e d o n a hard c o p i e r f o r on-site p r e l i m i n a r y run a n a l y s i s . once s e l e c t e d d a t a a r e s t o r e d , the

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d a t a can be recovered f o r a d d i t i o n a l a n a l y s i s e i t h e r on s i t e or by remote terminals i n t e r f a c e d w i t h the computer.

operational conditions. A p o r t i o n of the O p e r a t i o n a l c o r r i d o r t y p i c a l O s c r a m j e t s i s shown f i n Figure l(a1. T h i s c o r r i d o r r e p r e s e n t s t h e o p e r a t i o n of an engine concept t h a t Would o p e r a t e as a r a m j e t or with mixed subsonic-supersonic cornbustion over t h e f l i g h t Mach number range of 3 t o 6 and as a s c r a m j e t f o r f l i g h t r a c h numbers above 6. A range of f l i g h t c o n d i t i o n s t h a t can be simulated ( w i t h o u t precompression c o n s i d e r a t i o n ) by t h e b a s i c f a c i l i t y c o n f i g u r a t i o n i s r e p r e s e n t e d by t h e hatched area of Figure 8Ial. h e v e r t i c a l bar r e p r e s e n t s t h e p r e s e n t f a c i l i t y nozzle with a nominal e x i t Mach number of 3.5. Ihe flight sirnulation r e g i o n (hatched area) is d i c t a t e d by t h e Mach 3.5 nozzle, t h e p r e s e n t maximum operat i o n a l h e a t e r p r e s s u r e l i m i t of 190 p s i a , and t h e pumping c a p a b i l i t y of the a i r - e j e c t o r which exhausts t o t h e atmosphere ( n o z z l e Mach number, rncdel/diffuser arrangement/size dependent).

Proper c o n t r o l of t h e h e a t e r oxygen and hydrogen allows t h e h e a t e r t e s t gas temperature t o be v a r i e d t o s i m u l a t e f l i g h t S t a g n a t i o n temperat u r e s over a range of Mach numbers r e p r e s e n t e d i n Figure 7 I b l ; t h e b a s i c c o n f i g u r a t i o n has a h e a t e r maximum t o t a l temperature l i m i t Of Z250DR or about s i n c e t h e f a c i l i t y nozzle e x i t Mach 5 s i m u l a t i o n . Mach number remains e s s e n t i a l l y c o n s t a n t ( s l i g h t change because of temperature c h a n g e ) , t h e corresponding range of a n g l e s (assumed wedge) t h a t t h e v e h i c l e u n d e r s u r f a c e , r e l a t i v e t o the h o r i z o n t a l , has to a t t a i n t o maintain a c o n s t a n t Mach number 13.5 or M1 i n Fig. )(a)) ahead of the i n l e t i s Associated a l s o r e p r e s e n t e d i n Figure 7 I b l . curves f o r f l i g h t f r e e - s t r e a m s t a g n a t i o n p r e s s u r e and dynamic p r e s s u r e assuming a c o n s t a n t t o t a l p r e s s u r e ahead of t h e engine i n l e t of 190 p s i a , which i s t h e p r e s e n t h e a t e r maximum p r e s s u r e l i m i t , are p r e s e n t e d i n Figure 7 I b ) . ( S i m i l a r curves could of course be r e p r e s e n t e d a t lowerh e a t e r t o t a l p r e s s u r e values.1 Rs t h e t o t a l tempe r a t u r e s i m u l a t i o n value i s i n c r e a s e d the sirnul a t e d precompression a n g l e may n o t n e c e s s a r i l y s i m u l a t e a r e a l i s t i c f l i g h t c r u i s e or a c c e l e r a t i o n c o n d i t i o n ; however, t h e v a r i a b l e t o t a l temperature c a p a b i l i t y allows engine environment combustion s t u d i e s t o be performed very e a s i l y a t v a r i o u s t o t a l temperatures t h a t s i m u l a t e those f o r various f l i g h t Mach numbers.

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more expensive r e v i s i o n is a replacement of t h e a i r e j e c t o r exhaust system w i t h a vacuum sphere exhaust system s u f f i c i e n t l y s i z e d t o a l l o w 20-30 Second h o t test times. Plans f o r such r e v i s i o n s are underway; r e q u i r e d f a c i l i t y hardware and the e f f e c t on t h e f a c i l i t y f l i g h t s i m u l a t i o n capabil i t y w i l l be d i s c u s s e d i n t h e following sections.

L?

Heater l h e r e l a t i v e p o s i t i o n of t h e mixer a n d h e a t e r assembly i n t h e b a s i c f a c i l i t y configur a t i o n is shown i n t h e schematic of Figure 5. Ihe schematic i n Figure 8(al shows some of the major A p r e v i o u s l y mens d e t a i l s of t h i s assembly. t i o n e d , t h e a i r and oxygen are premixed i n t h e mixer s e c t i o n between the oxygen and hydrogen b a f f l e plates; t h u s t h e hydrogen is i n j e c t e d i n t o oxygen-rich a i r . Both b a f f l e p l a t e s have 2 r i n g s of o r i f i c e s , 10 o r i f i c e s i n t h e i n n e r r i n g and 20 i n the o u t e r r i n g , through which i n j e c t o r tubes pass. I g n i t i o n of t h e h e a t e r p r o p e l l a n t s is provided by a hydrogen-oxygen t o r c h i g n i t o r t h a t is i n s t a l l e d as shown i n Figure 8Ial. Premixing t h e a i r and Oxygen r e s u l t s i n good mixing and thus a l l o w s t h e length-to-diameter r a t i o of 3.3 t o be s u f f i c i e n t f o r good combustion i n the h e a t e r . h e d e s i g n of t h e passages through t h e oxygen and hydrogen b a f f l e p l a t e s was such t h a t t h e a i r Mach number i s about 0.9 and t h e air-oxygen mixture Mach number i s about 0.7 through t h e r e s p e c t i v e baffle p l a t e s . Oxygen and hydrogen are i n j e c t e d through o r i f i c e s a t about M = 0.7.

The p r e s e n t test gas h e a t e r shorn i n Figure 8Ial has a l o n g i t u d i n a l - s l o t c o o l i n g water j a c k e t (450 gal./min. Water f l o w ) . This h e a t e r d u c t is an item of s u r p l u s equipment f o e which t h e d e s i g n documents were n o t a v a i l a b l e . After Safety t e s t s were performed, t h e d u c t was s a n c t i o n e d f o r use i n t h e secured t e s t c e l l a t maximum o p e r a t i n g l i m i t s of 190 p s i a and 2250% which c o n t r i b u t e to t h e p r e s e n t r e s t r i c t e d o p e r a t i o n a l range of the facility. The replacement t e s t gas h e a t e r t o be i n s t a l l e d i n 1987 w i l l extend the f a c i l i t y operat i o n a l range. The design of this h e a t e r d u c t i n c o r p o r a t e s a 0.75-inch t h i c k , heat-sink 201n i c k e l l i n e r w i t h an 18.0-inch i n t e r n a l diameter contained w i t h i n a 24.0-inch diameter schedule-40 carbon s t e e l pipe. P r o v i s i o n s are made t o a l l o w the o p t i o n of t h e l i n e r t o be back-side cooled w i t h a i r flowing from the downstream end toward t h e upstream end and i n t o t h e h e a t e r chanlber as shown i n Figure 8 ( b ) . An a d a p t o r f l a n g e w i l l be used t o a t t a c h the new h e a t e r t o the e x i s t i n g mixer section/hydrogen b a f f l e p l a t e Shown i n Figure 8(aI. A downstream end f l a n g e i s a rep l a c e a b l e a d a p t o r f l a n g e to a c c e p t v a r i o u s f a c i l i t y nozzles.

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Wif icatioos

The f l i g h t s i m u l a t i o n c a p a b i l i t y O t h e b a s i c f f a c i l i t y c o n f i g u r a t i o n i s being expanded t o a l l o w a l a r g e r range of t r u e s i m u l a t i o n i n s u b s c a l e engine t e s t 3 i n t h e CHSTF. obvious l i m i t a t i o n s for s c r a m j e t t e s t i n g are t h e test gas h e a t e r t o t a l p r e s s u r e and temperature limits Of 190 p s i a and 2250°R, r e s p e c t i v e l y . New design maximums are: (11 a h e a t e r t o t a l p r e s s u r e of 600 p s i g I t h e a i r supply l i m i t ) , and ( 2 1 a h e a t e r t o t a l ternperat u r e of about 3000-R. I n o r d e r to t a k e advantage of such expanded t e s t gas h e a t e r c a p a b i l i t i e s , new f a c i l i t y nozzles w i t h d i f f e r e n t e x i t Mach numbers are r e q u i r e d t o properly s i m u l a t e t h e f l i g h t precompressed flow c o n d i t i o n s e n t e r i n g the engine inlet. Heater and nozzle extended c a p a b i l i t i e s are r e l a t i v e l y inexpensive r e v i s i o n s t o expand t h e f a c i l i t y f l i g h t s i m u l a t i o n c a p a b i l i t y . Another

Teat setup m n f i g u r a t i o n . IWO f a c i l i t y t e s t c o n f i g u r a t i o n s are shown i n the schematics of Figure 9. The hardware arrangement f o r a superThis s o n i c test Setup i s shown i n Figure 9 ( a ) . s e t u p uses e i t h e r t h e p r e s e n t Mach 3.5 nozzle or t h e new Mach 4.7 nozzle which are connected d i r e c t l y t o t h e t e s t gas h e a t e r downstream f l a n g e . The engine model i s i n s t a l l e d a t t h e f a c i l i t y nozzle e x i t , as d i s c u s s e d f o r Figure 3 I b ) . suspended from a mounting beam by four f l e x beams A t h a t a l l o w l o n g i t u d i n a l movement of the engine. ereloaded one-component f o r c e balance senses the l o n g i t u d i n a l t h r u s t / d r a g f o r c e of t h e engine. h e nozzle e x t e n s i o n , t h e exhaust flow catch cone, and

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t h e i n i t i a l p o r t i o n of the s u p e r s o n i c d i f f u s e r are shown i n Figure 9(aI ( l a t t e r two shown as dashed lines I.

Figure 11. These r e s u l t s were o b t a i n e d w i t h multiprobe rake8 p o s i t i o n e d a t t h e nozzle e x i t p l a n e along t h e v e r t i c a l c e n t e r l i n e . P r o f i l e s f o r t h e Mach 3.5 nozzle a t 95 p s i a and 1600°R h e a t e r test gas s t a g n a t i o n c o n d i t i o n s are p r e s e n t e d i n Figure 1 1 ( a ] . ?he e x i t i n s t r e a m p i t o t p r e s s u r e s and total temperatures were obtained by use of a 13-probe (one-inch spacing1 S t a t i o n a r y rake with a l t e r n a t e p i t o t p r e s s u r e and t o t a l temperature probes (seven and six probes, r e s p e c t i v e l y ] . Boundary l a y e r p i t o t p r e s s u r e surveys were obtained w i t h a 7-probe s t a t i o n a r y rake. S t a t i c p r e s s u r e instream surveys were made b u t the r e s u l t s were the l e a s t r e l i a b l e of t h e The nozzle wall e x i t s t a t i c survey measurements. p r e s s u r e was assumed as an i n s t r e a m c o n s t a n t s t a t i c p r e s s u r e and used with t h e measured i n s t r e e m p i t o t pressures. This was n o t a v a l i d assumption because a t some l o c a t i o n s i n t h e stream near the boundary l a y e r edge, t o t a l p r e s s u r e s g r e a t e r than t h e t e s t g a s h e a t e r t o t a l p r e s s u r e were computed from t h e r e s u l t a n t Mach numbers. merefore, a computed t o t a l p r e s s u r e t r e n d a s s o c i a t e d with t h e wall s t a t i c and p i t o t p r e s s u r e measurements i n t h e boundary l a y e r was f a i r e d i n t o a c o n s t a n t v a l u e equal t o t h e h e a t e r test gas t o t a l p r e s s u r e . This t o t a l p r e s s u r e t r e n d was then used with t h e measured p i t o t p r e s s u r e s t o y i e l d t h e r a t i o of p i t o t t o t o t a l p r e s s u r e p r o f i l e shown i n Figure l l l a ) . The. t o t a l temperature p r o f i l e is n e a r l y c o n s t a n t and is f a i r e d through t h e boundary l a y e r toward an a d i a b a t i c wall recovery temperature. The Mach number p r o f i l e shown was d e r i v e d from the p i t o t to-total pressure r a t i o profile. s i n c e t h e square nozzle c o n t o u r s are o b t a i n e d from t h e s t r e a m l i n e t r a c i n g technique u s i n g axisymmetric nozzle f l o w s , some t o t a l p r e s s u r e l o s s is l i k e l y t o occur i n the a x i s r e g i o n of t h e nozzle e x h a u s t flow. Such reduced t o t a l p r e s s u r e s Would tend t o f l a t t e n the Mach number p r o f i l e i n t h e a x i s r e g i o n . The Mach number of 3.55 i n d i c a t e d on t h e Mach number f i g u r e was c a l c u l a t e d u s i n g t h e nozzle w a l l e x i t s t a t i c p r e s s u r e - t o - h e a t e r t e s t gas t o t a l p r e s s u r e r a t i o .

E x i t flow p r o f i l e s f o r t h e t r a n s o n i c m i s y m m e t r i c nozzle are p r e s e n t e d i n Figure l l ( b 1 . These p r o f i l e s were Obtained u s i n g the same 13probe rake which was used f o r t h e Mach 3.5 nozzle except a l l probes were converted t o measure p i t o t p r e s ~ w e . Boundary l a y e r surveys were n o t performed. The measurements of t h e two end probes were i n l i n e with t h e nozzle e x i t w a l l s 112.0 i n c h d i a m e t e r ) and may be i n a distltebed flow. The p r e s s u r e p r o f i l e s are a r a t i o of t h e nozzle w a l l e x i t s t a t i c p r e s s u r e t o measured p i t o t p r e s s u r e s . The Mach number p r o f i l e s shown are a f u n c t i o n of these pressure r a t i o s . Noted a t t h e t o p of t h e f i g u r e s are t h e t u n n e l chamber t o t a l p r e s s u r e , nozzle wall e x i t s t a t i c p r e s s u r e , and Mach number as a f u n c t i o n of t h e r a t i o of nozzle w a l l s t a t i c to-tunnel total p r e s s u r e . The p r o f i l e s shown for t h e t r a n s o n i c nozzle are r e p r e s e n t a t i v e of t e s t c o n d i t i o n s w i t h and without t h e o p e r a t i o n of t h e facility a i r ejector.

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The t e s t s e t u p f o r s u b s o n i c / t r a n s o n i c experimental i n v e s t i g a t i o n s is r e p r e s e n t e d by t h e Schematic of Figure 9 ( b ) . I n this s e t u p the test c a b i n , model i n s t a l l a t i o n , and exhaust System remain t h e same as shown i n Figure 9 ( a ) . The s u b s o n i c / t r a n s o n i c c o n f i g u r a t i o n is obtained by r e p l a c i n g t h e f a c i l i t y air/oxygen mixer s e c t i o n , t h e hydrogen-air-burning h e a t e r d u c t , and t h e contoured s u p e r s o n i c nozzle with d u c t i n g t h a t d e l i v e r s unheated a i r t o t h e engine medal. Two nozzle blocks can be e a s i l y exchanged t o y i e l d Mach 1.2 flow or subsonic flow with t h e convergent Mach 1.0 nozzle. Various Subsonic Mach number flows may be o b t a i n e d by a d j u s t i n g the chamber I h e nozzle e x i t s t a t i c p r e s s u r e s t o t a l pressure. f o r both nozzle b l o c k s is s l i g h t l y helow atmosp h e r i c p r e s s u r e w i t h o u t t h e use of t h e f a c i l i t y a i r e j e c t o r . when t h e f a c i l i t y a i r e j e c t o r is Operated a t l o w flow r a t e s , t h e nozzle e x i t p r e s sures are about ?ne-half atmosphere which simul a t e s an a l t i t u d e of about 17,500 f e e t . During subsonic t e s t s without the f a c i l i t y a i r e j e c t o r o p e r a t i n g , the test c a b i n p r e s s u r e changes s l i g h t l y as combustion occurs i n the engine. This i n t u r n a f f e c t s the tunnel operating pressure, which causes t h e f a c i l i t y nozzle e x i t Mach number t o change s l i g h t l y .

.d

Ilozrles. The p r e s e n t f a c i l i t y s u p e r s o n i c n o z z l e , shown i n t h e s k e t c h of Figure 1 0 ( a ) , is an uncooled contoured s q u a r e nozzle designed on the b a s i s of s t r e a m l i n e - t r a c i n g t h e flow of an miSymmetric nozzle. The t h r o a t is 4.976 i n . square ( t h r o a t area of 24.81 and t h e flow e x i t i s nominally 13 i n . square ( a c t u a l geometric nozzle e x i t dimensions are 13.264 i n . to account f o r boundary l a y e r displacement t h i c k n e s s ] . A t a t o t a l temperature of 1640PR, t h e n o z z l e - e x i t Mach number i s 3.50. The nozzle e n t r a n c e , which p r o t r u d e s i n t o t h e h e a t e r d u c t , makes a t r a n s i t i o n from a c i r c u l a r t o a square cross s e c t i o n . The t h r o a t s e c t i o n was Constructed with a l a r g e mass of s t a i n l e s s s t e e l f o r h e a t s i n k , and the downstream expansion s e c t i o n of the nozzle was Constructed of 0.183-in-thick carbon s t e e l with e x t e r n a l s t i f f e n i n g webs. The nozzle s i d e w a l l s and t o t t o m wall were extended a t the e x i t t o ensure t h a t shocks, g e n e r a t e d when the r a t i o of t e s t cabin to nozzle-exit s t a t i c pressure i n c r e a s e d from 1.0 t o 2.0, would n o t e n t e r t h e i n t e r n a l f l o w region of an engine model.

An a d d i t i o n a l nozzle, a v a i l a b l e i n t h e l a t t e r p e r t of 1987, is shown i n Figure 1 0 l b ) . This nozz l e is similar t o t h e Mach 3.5 nozzle b u t has an e x i t Mach number of 4.7 t o s i m u l a t e a i r f r a m e i n t e g r a t e d peeconpression a t a f l i g h t Mach number 0 E 5.5. a copper water-cooled t h r o a t S e c t i o n i s i n c o r p o r a t e d i n t h e design t o a l l o w o p e r a t i o n a t Hach 6.0 f l i g h t S t a g n a t i o n temperature (3000'R). The downstream contoured s e c t i o n i s ""cooled and i g f a b r i c a t e d by t h e same method as t h e p r e s e n t The t h r o a t i s 2.588 i n c h e s Mach 3.5 nozzle. square ( t h r o a t area of 6.70 i n . 2 ) and t h e e x i t is t h e same as t h e Mach 3.5 nozzle, 13.264 inches

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nozzle exhaust flows. - Exhaust flow p r o f i l e s of the e x i s t i n g f a c i l i t y n o z z l e s are p r e s e n t e d i n

The v a r i o u s p r o f i l e s shown i n Figure l l l c l f o r t h e subsonic nozzle are t h e r e s u l t of d i f f e r e n t t u n n e l t o t a l pressures as i n d i c a t e d a t t h e t o p of t h e f i g u r e ( t h e p i t o t measurements are of course i n s t r e a m t o t a l p r e s s u r e measurements). Most of t h e s e t e s t s were performed with the nozzles exh a u s t i n o i n t o t h e c l o s e d test c a b i n w i t h o u t the

5

f a c i l i t y air e j e c t o r i n operation. s l i g h t aspirat i o n of t h e t e s t c a b i n r e s u l t e d w i t h t h e nozzle e x h a u s t flow e n t e r i n g i n t o the e x h a u s t d u c t c a t c h cone; t h e nozzle e x i t s t a t i c pressures i n d i c a t e d a t t h e t o p of the f i g u r e are very n e a r l y equal t o t h e t e s t cabin pressure. I n o r d e r t o o b t a i n near s o n i c c o n d i t i o n s , t h e f a c i l i t y a i r e j e c t o r had t o be used; t h e elongated-diamond-symbol p r o f i l e s i n Figure l l ( c 1 are r e s u l t s obtained w i t h t h e a i r e j e c t o r i n operation.

erpanded Operational conditione. The a d d i t i o n of t h e new t e e t gas h e a t e r and Mach 4 . 1 n o z z l e r e s u l t s i n expanded f a c i l i t y o p e r a t i o n a l conditions that simulate t r u e f l i g h t conditions ( w i t h o u t precompression) as r e p r e s e n t e d by t h e hatched area of Figure 12. The h i g h e r p r e s s u r e and temperature l i m i t s of t h e new h e a t e r a l l o w b o t h lower a l t i t u d e and h i g h e r Mach number simulat i o n . A larger range of v e h i c l e precompression s i m u l a t i o n is a v a i l a b l e with t h e a d d i t i o n a l Mach 4.7 nozzle. Higher a l t i t u d e / l o w e r dynamic p r e s sure s i m u l a t i o n s remain l i m i t e d by t h e o p e r a t i o n a l c a p a b i l i t y of t h e a i r e j e c t o r as shown i n Figure 12. Also shown i n Figure 12 are the subsonic and t r a n s o n i c nozzle f l i g h t s i m u l a t i o n c a p a b i l i t i e s .

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Test Pxperience

m u s t a c t flow. Flow c o n d i t i o n s throughout t h e f a c i l i t y d u c t i n g can be determined from d u c t w a l l premure and p i t o t - p r e s s u r e rake measurements; t h e r e l a t i v e l o c a t i o n s of t h e wall and r a k e pressure ,neasurements are i n d i c a t e d i n Figure 1 3 ( a I . Two s p e c i a l problems are a s s o c i a t e d w i t h t e s t i n g scramjets i n the f a c i l i t y . First, i n s t a l l a t i o n of a s c r a m j e t model w i t h t h e i n l e t t o p s u r f a c e a l i g n e d w i t h or near t h e f a c i l i t y n o z z l e t o p w a l l c r e a t e s asymmetric flow p r o f i l e s i n t h e e x h a u s t d u c t ( g r e a t e r pressure l o s s e s o n t h e t o p s i d e of the d u c t i n g ) . Second, the f a c i l i t y e x h a u s t d u c t mixing l e n g t h had t o be made s h o r t e r than Optimum because of t h e containment of t h e f a c i l i t y w i t h i n an e x i s t i n g t e s t c e l l ; t h i s r e s u l t s i n t h e asymmetric flow s t i l l being i n e x i s t e n c e a t t h e e n t r a n c e of t h e t u r n i n g elbow. These flow c o n d i t i o n s are e v i d e n t i n Figure 1 3 ( b ) . ?he undisturbed t r e n d s ( c i r c l e symbols) are e v i d e n t a t the s t a r t of a test and a t moderate engine f u e l flow r a t e s . During t e s t s O poor engine f u e l f combustion or a t high f u e l flow r a t e s , some of t h e f u e l i s n o t burned e n t i r e l y w i t h i n t h e engine. This unburned f u e l a p p a r e n t l y i g n i t e s i n t h e fac i l i t y exhaust d u c t and g e n e r a t e e i n c r e a s e d p r e s sures i n t h e d u c t t h a t e v e n t u a l l y a f f e c t t h e tunn e l flow i n t h e region of the engine. This condit i o n is referred to as f a c i l i t y - e n g i n e i n t e r a c t i o n and d u c t p r e s s u r e measurements a s s o c i a t e d w i t h the d i s t u r b e d d u c t flow are r e p r e s e n t e d by t h e square symbol t r e n d s i n Figure 1 3 ( b l . ?he test c a h i n S t a t i c p r e s s u r e and the exhaust s t r a i g h t d u c t s u p e r s o n i c d i f f u s e r e x i t S t a t i c pressure and flow Mach number are good i n d i c a t o r s of f a c i l i t y flow breakdown ( f a c i l i t y - e n g i n e i n t e r a c t i o n ) and are g e n e r a l l y observed during a t e s t and/or d u r i n g p o s t t e s t o n - s i t e a n a l y s e s of t h e test d a t a . Large engine models canse more f a c i l i t y s e n s i t i v i t y to these facility-engine interactions.

A few techniques have been s u c c e s s f u l i n minimizing t h e s e i n t e r a c t i o n s . Water Sprayed i n t o t h e d i f f u s e r c a t c h cone and a t t h e elbow t u r n i n g vane l e a d i n g edges has been e f f e c t i v e . Also, an

i n c r e a s e d l e n g t h of t h e f a c i l i t y nozzle e x i t e x t e n s i o n decreased t h e f a c i l i t y f r e e - j e t l e n g t h ( a b o u t 0.75 e q u i v a l e n t d i a m e t e r s ) and made the nozzle flow less s e n s i t i v e t o i n c r e a s e s i n t h e exhaust-duct back p r e s s u r e . W i t h t h e s e techniques i n c o r p o r a t e d , the e x h a u s t d u c t flow t r e n d s remained undisturbed ( c i r c l e symbol curve, ~ i g .1 3 ( b ) ) a s engine f u e l f l o w r a t e was i n c r e a s e d t o much h i g h e r l e v e l s of flow r a t e s before t h e exh a u s t d u c t flows became d i s t u r b e d (similar t o square symbol curves) than was p o s s i b l e p r i o r to the r e s o l u t i o n s . P r i o r t o the r e s o l u t i o n of t h e i n t e r a c t i o n problem, the i n c r e a s e d p r e s s u r e s i n t h e exhaust system a f f e c t e d t h e e n g i n e measurementa and d i s t i n c t i o n was d i f f i c u l t between engine combustor-inlet i n t e r a c t i o n s and f a c i l i t y engine interactions. When t h e engine i n l e t flow """started," t h e f a c i l i t y nozzle flow was d r a s t i c a l l y affected. After t h e problem r e s o l u t i o n s , e n g i n e combustor-inlet i n t e r a c t i o n s , caused by model combustor fuel-burning p r e s s u r e r i s e , could be o b t a i n e d b e f o r e the occurrance of or without a facility-engine interaction. Also, a f t e r the changes were made, even w i t h t h e i n l e t u n s t a r t e d and a f a c i l i t y exhaust d u c t flow breakdown, the e j e c t o r could maintain t h e cabin pressure s u f f i c i e n t l y low so t h a t t h e f a c i l i t y nozzle flow was seldom a f f e c t e d .

Wels teated. Four d i f f e r e n t s u b s c a l e Scramjet engine models have been t e s t e d i n t h e p r e s e n t c o n f i g u r a t i o n of t h e CHSTF a t nominal s t a g n a t i o n c o n d i t i o n s Of 92.4 p s i a and 1640'R which simulated Mach 4 f l i g h t a t a dynamic pressure of 1000 p s f . The models were l a r g e r e l a t i v e t o t h e f a c i l i t y nozzle exit of 169 sq. i n c h e s . Some of t h e model c o n f i g u r a t i o n s could be v a r i e d r e s u l t i n g i n d i f f e r e n t i n l e t cowl f r o n t a l areas which, f o r t h e f o u r models, ranged from 12 t o 30 p e r c e n t OC t h e f a c i l i t y nozzle e x i t a m a ! maximum e x t e r n a l croes sectional areas ranged from 31 to 49 percent. when t h e s e models were i n s t a l l e d w i t h a c a t c h cone, nozzle e x t e n s i o n . and water s p r a y s a s diecussed p r e v i o u s l y , s u p e r s o n i c flow could be e s t a b l i s h e d through t h e exhaust d i f f u s e r ( p i t o t and w a l l S t a t i c pressures a t t h e d i f f u s e r e x i t i n d i c a t e d the flow t o be a t Mach number of about 1 . 4 a t zero engine f u e l f l o w ) . During such condit i o n s , t h e cabin pressure was very n e a r l y r q u a l t o t h e f a c i l i t y nozzle e x i t pressure. As t h e e n g i n e f u e l flow was i n c r e a s e d , t h e d i f f u s e r flow Mach number decreased and f o r 8ome high f u e l flow r a t e t e s t s , t h e d i f f u s e r e x i t flow Mach number became or was a b o u t to become subsonic. ?he c a b i n pressure f o r t h e s e cases i n c r e a s e d to a p o i n t t h a t a f a c i l i t y - e n g i n e i n t e r a c t i o n occurred. As ment i o n e d p r e v i o u s l y , l a r g e r m d e l blockage caused more s e n s i t i v i t y t o f a c i l i t y - e n g i n e i n t e r a c t i o n s . I t i s t h e r e f o r e recommended t h a t a c o n s e r v a t i v e approach should be taken and an a t t e m p t made t o d e s i g n f o r a mcdel cowl area blockage of 25 perc e n t or less and/or 35 p e r c e n t model maximum area blockage. Also, a f a c i l i t y should be designed w i t h d i f f u s e r and mixer d u c t l e n g t h s of optimum l e n g t h where p o s s i b l e . This approach should res u l t i n more s t a b l e exhaust d u c t flow c o n d i t i o n s .

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s F a c i l i t y w i t h M N U ~ phere

AS mentioned p r e v i o u s l y , the Simulated f l i g h t envelope f o r e n g i n e t e s t s i n t h e CHSTF is curr e n t l y l i m i t e d i n a l t i t u d e s i m u l a t i o n by t h e a i r e j e c t o r exhaust system. This e j e c t o r system

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consumes l a r g e q u a n t i t i e s of t h e f a c i l i t y a i r

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s u p p l y while p r o v i d i n g o n l y marginal a l t i t u d e simulation. In a d d i t i o n , facility-model i n t e r a c t i o n s have, d u r i n g some tests, i n t e r f e r e d w i t h o b t a i n i n g d a t a a t model f u l l t h r o t t l e s e t t i n g . TO s o l v e t h e s e problems, t h e i n s t a l l a t i o n of a vacuum s p h e r e system i s planned which w i l l be shared hy the CHSTF and the Test C e l l 112 Direct-Connect Combustor F a c i l i t y . This p r o j e c t is i n t h e f i n a l d e s i g n phase and is earmarked f o e 1988-89 cons t r u c t i o n funds. The n e w vacuum sphere w i l l provide t h e needed reduced back p r e s s u r e c a p a b i l i t y t o allow f l i g h t c o n d i t i o n s i m u l a t i o n f o r engine t e s t i n g i n t h e CHSTF With a l t i t u d e s i m u l a t i o n from 30,000 t o 120,000 f e e t and run t i m e s of 20 seconds. In a d d i t i o n , the vacuum sphere w i l l e l i m i n a t e t h e h i g h a i r flow r e q u i r e d by t h e p r e s e n t a i r e j e c t o r and e f f e c t i v e l y uncouple t h e test c e l l from t h e o t h e r f a c i l i t i e s with which it competes f o r a i r .

t h e p r e s e n t c o n f i g u r a t i o n that was shown i n Figure 5.

operational Conditions. me proposed vacuum sphere System w i l l a l l o w Mach 3.5 t o 6 f l i g h t c o n d i t i o n s ( w i t h o u t precompression) a t a l t i t u d e s from 30,000 t o 120,000 f e e t to he simulated as r e p r e s e n t e d by t h e hatched area of Figure 15. The low a l t i t u d e and low Mach number s i m u l a t i o n (high dynamic p r e s s u r e s i m u l a t i o n ) may r e q u i r e the opt i o n a l r e t u r n t o t h e f a c i l i t y a i r - e j e c t o r configuration. This a d d i t i o n Of t h e vacuum system W i l l g r e a t l y enhanoe t h e combustion-Heated Scramjet T e s t r a c i l i t y c a p a b i l i t y to b e t t e r serve t h e N S AA and n a t i o n a l ramjet/scrarnjet community.

R e l a t i o n e h i p to rmsn-rsnqlev scramjet l e s t Caplei. me enhanced c a p a b i l i t i e s of t h e Combustion-Heated s c r a m j e t lkst F a c i l i t y r e p r e s e n t s both an improvement t o the f a c i l i t y i t s e l f and a l s o c o n t r i b u t e s t o t h e enhancement of t h e o v e r a l l c a p a b i l i t y of t h e NASA-Langley Scramjet T e s t Complex.l0 This complex i n c l u d e s t h e Langley 8-Foot High Temperature m n n e l (8-ET HTT) which i s p r e s e n t l y part o f t h e Rerothernal Loads Complex. B 1989 t h e f a c i l i t y w i l l have an oxygen r e p l e y nishment system and new f a c i l i t y nozzles.11 These enhancements give t h i s t u n n e l t h e c a p a b i l i t y of engine t e s t s a t Mach 4, 5, and 7 f l i g h t c o n d i t i o n s (8-foot diameter nozzle e x i t s ) of l a r g e - s c a l e engine models (about 20 x 2 8 i n c h e s ) , a c l u s t e r of m u l t i p l e engine modules, or engine models t h a t have f u l l nozzle expansion s u r f a c e s , The ArcHeated Scramjet T e s t F a c i l i t y (AHSTF) now has a much expanded t e s t c a p a b i l i t y ' with two d i f f e r e n t nozzles ( M = 4.7 and 6.0) and i s p a r t of t h e NASALangley s c r a m j e t Test Complex. Upon completion of t h e m o d i f i c a t i o n s t o t h e 8-FI H1T and those d i s cussed h e r e i n f o r t h e CHSTF a l o n g with the extended t e s t c a p a b i l i t y of t h e AHSTF, t h e Scramjet Test Complex a t Langley w i l l be unique i n t h e w e s t e r n world. The f l i g h t s i m u l a t i o n envelope Of t h e t e s t gas h e a t e r s ( w i t h o u t precompression) f o r t h e 8-FT HTT and t h e AHSTF are shown i n Figure 16 superimposed on t h e o p e r a t i o n a l envelope of t h e CHSTF t h a t was p r e s e n t e d i n Figure 15. The potent i a l o p e r a t i o n a l envelope of t h i s complex would extend over a f l i g h t Mach number range from 3.5 t o 7. The lower Mach number l i m i t shown f o r each of t h e s e f a c i l i t i e s r e p r e s e n t s the lowest Mach number nozzle a v a i l a b l e f o r each f a c i l i t y .

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Description. I h e f a c i l i t y w i l l be connected t o t h e 70-foot vacuum s p h e r e by approximately 200 f e e t of 48-inch vacuum p i p e l i n e . A remotely o p e r a t e d b u t t e r f l y v a l v e w i l l be i n s t a l l e d i n t h e vacuum pipe near the sphere e n t r a n c e . A remotely o p e r a t e d b u t t e r f l y v a l v e w i l l a l s o be i n s t a l l e d a t the f a c i l i t y / s p h e r e s u c t i o n connection i n the t e s t c e l l . The vacuum sphere and connecting p i p i n g w i l l be evacuated by a f o u r S t a g e condensing steam e j e c t o r d i s c h a r g i n g to atmosphere through an exhaust silencer. An e x i s t i n g 6-inch stem l i n e o p e r a t i n g a t 350 p s i g and 26,000 pounds per hour w i l l supply the steam e j e c t o r . Condenser water w i l l be s u p p l i e d by an e x i s t i n g c o o l i n g Water system. S i n c e t h e system w i l l be shared w i t h T e s t C e l l X2. each f a c i l i t y w i l l be provided w i t h a remotely o p e r a t e d key-interlocked i s o l a t i o n butt e r f l y valve and remotely operated atmospheric purge valve. Piping from t h e test cells t o the s p h e r e w i l l be m n from t h e exhaust end Of the test c e l l s up, over, and across t h e roof t o t h e sphere. The schematic s k e t c h shown i n Figure 14 d e p i c t s t h e f a c i l i t y hardware/ducting arrangement i n the t e s t cell. m h a u s t d u c t s a s s o c i a t e d with t h e f a c i l i t y a i r e j e c t o r w i l l be removed ( s e e Fig. 5 ) and new d u c t i n g wi.11 be i n s t a l l e d as shown i n Figure 1 4 . For s a f e t y c o n s i d e r a t i o n s , a s m a l l amount ( a b o u t 5 l b s . per sec.) of COntinUOUSly flowing purge a i r is r e q u i r e d through t h e t u n n e l . During t h i s purge flow, t h e sphere w i l l be i s o l a t e d from t h e t u n n e l by t h e 48-inch b u t t e r f l y v a l v e near t h e t e s t c e l l c e i l i n g . The 30-inch d u c t , with i t s b u t t e r f l y valve open, d i r e c t s t h i s purge flow t o t h e v e r t i c a l exhaust d u c t to t h e atmosphere. The purge flow Sets up an e j e c t o r e f f e c t , as i n the p r e s e n t f a c i l i t y a i r e j e c t o r c o n f i g u r a t i o n s , t o c r e a t e a i r flow i n t h e t e s t c e l l t o "bathe" t h e e n t i r e hardware of the f a cility. This a c t i o n is a s s i s t e d by a f a n i n S t a l l e d i n t h e test c e l l c e i l i n g i n t a k e tower; t h e m a i n f u n c t i o n of t h e f a n , however, i s to provide t e s t c e l l b a t h i n g a i r flow during t u n n e l flow to t h e vacuum sphere. The vacuum sphere system is t o be i n s t a l l e d so t h a t , i f r e q u i r e d , the p r e s e n t a i r e j e c t o r system could be r e i n s t a l l e d . That i s , the 48-inch i s o l a t i o n b u t t e r f l y valve a t t h e t e s t c e l l c e i l i n g would i s o l a t e the vacuum sphere and t h e new d u c t i n g shorn i n Figure 14 would be replaced with

-

Concluding &marks

A t e s t f a c i l i t y h a s been assembled a t N S AA Langley Research Center t o provide t h e c a p a b i l i t y f o r s u b s c a l e a i r - b r e a t h i n g r a m j e t / s c r a m j e t engine tests. Hydrogen burned i n a i r w i t h oxygen r e p l c nishment y i e l d s simulated a i r with t h e proper oxygen c o n t e n t f o r combustion t e s t s . The p r e s e n t c o n f i g u r a t i o n is a f r e e - j e t wind t u n n e l with a three-dimensional Mach 3.5 nozzle (nominal 13-ins q u a r e e x i t ) and an exhaust a i r e j e c t o r f o r simul a t i o n of Mach 4 f l i g h t c o n d i t i o n s a t t h e i n l e t of an a i r f r a m e - i n t e g r a t e d s c r a m j e t . The f a c i l i t y is d e s i g n a t e d t h e Combustion-Heated Scramjet T e s t F a c i l i t y (CHSTFI.

' 2

D e t a i l s of t h e CHSTF were d e s c r i b e d h e r e i n and m o d i f i c a t i o n s t o enhance t h e t e s t c a p a b i l i t y range from s u b s o n i c / t r a n s o n i c and Mach 3.5 t o 5.5 were discussed. This f a c i l i t y allows r e l a t i v e l y i n e x p e n s i v e s u b s c a l e test of ramjet-type engines t o be conducted.

7

Referems

1.

7.

GUY,

Hypersonic Research Engine P r o j e c t Phase I1 S t r u c t u r e s and Cooling Development F i n a l T e c h n i c a l D a t a Rewrt. NASA CR-112087, M Y

T9T2.

~~ ~~~~

-

-~

~

Robert W.; Terrence, Marvin, G.; Sabol, Alexander P.i and Mueller, James N.: Operating C h a r a c t e r i s t i c s O the Langley Mach 7 s c r a m j e t f r r s t Facility. N S TM-81929, 1981. AA Thomas, S c G t t ~ R . and Guy, Robert w.: Ekpanded o p e r a t i o n a l C a p a b i l i t i e s of t h e Langley Mach 7 scramjet Test Facility. L AA lTr2186, 1983. US

8.

b

2.

Hypersonic Rese2r~h Engine P r o j e c t Phase I1 Aerothermodynamic I n t e g r a t i o n Model Development F i n a l Technical Data Report; NASA CR-132654. May 1975. Hypersonic Research Engine P r o j e c t Phase I1 Aerothemodynanic I n t e g r a t i o n Model ( A I M ) T e s t Report. N S CR-132655, May 1975. AA Northem, G. Burton and Anderson, G. Y . : s u p e r s o n i c Combustion & M e t Research at Langley. AIAA-86-0159: January 1986. Guy, Robert W.; Pinckney, S. Zanr; and Andrew$, Summary of N S Langley's AA E a r l H., Jr.: S c r a m j e t h g i n e Tests. 1985 JANNAF Propulsion Meeting, San Diego, C a l i f o r n i a , A p r i l 9-12, 1985. (Conf.)

d us sin, w i l l i a m Roger: Performance of a Hydrogen BWfier t o Simulate Air E n t e r i n g NASA TN 0-7576. 1974. Scramjet Conbustm-s.

-

9.

3.

-

Andrew*, M r l H a . JT.; Terrence, Marvin G.; Anderson, G r i f f i n Y; Northam, G. Burton; and . Mackley, E r n e s t A.: Langley Mach 4 s c r a m j e t Test Facility. N S 1)4-86277. 1985. AA

Scramjet T e s t Complex. Andrews, E a r l H.: Langley Aerospace T e s t H i g h l i g h t s , NASA TM87703. p. 34, 1985.

~ ~

10.

4.

11.

5.

K e l l y , H. N. and Wieting, A. R . : ~ Modification of N S Langley 8-Foot High Temperature l u n n e l AA t o Provide a Unique N a t i o n a l Research F a c i l i t y f o r Hypersonic Air-Breathing Propulsion Systems. NASA M 8 5 7 8 3 , May 1984.

~

6.

(b) S i m u l a t i o n i n ground f a c i l i t i e s .

Fig.

3.

F l i g h t simulation f o r engine t e s t s .

Fig. 2.

NASA-Langley Hyperemic P r o p u l s i o n ; Airframe I n t e g r a t e d Concept, 1970 to present.

Fig. 4.

Hypersonic Peopulsion Branch's Zlnnine and Component Test F a c i l i t i e s .

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(a)

F l i g h t / t e s t envelope; no precompression. Fig. 7.

(b) Variable Mach number s i m u l a t i o n .

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.

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(b)

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9

(a) supersonic test setup.

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Fig. 12.

Expanded f l i g h t / t e s t envelope; no precompression.

Fig.

15.

F a c i l i t y full F t e n t i a l f i i g h t / t e s t envelope; no precompression.

Srramjd lerl Facility

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Fig. 16.

Scramjet Test Complex h v e l o p e ; no precompression.

13.

Exhaust duct flow; supersonic setup.

Fig. 14.

Planned 70-foot V ~ C U U ~ sphere system.

11

Information

A subsonic to Mach 5.5 subscale engine test facility

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