Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

NASA

Tech nica I

Paper

Comparison of Wind Tunnel

and Flight Test Afterbody

and Nozzle Pressures for a

Twinjet Fighter Aircraft

at Transonic Speeds

Jack Nugent

Ames Research Center

Dryden Flight Research Facility

Edwards, Calfornia

Odis C. Pendergraft, Jr.

Langley Research Center

Hampton, Virginia

National Aeronautics and Space Administration Scientific and Technical Information Branch CONTENTS Page

S U M M A R Y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

I N T R O D U C T I O N . e 1 NOMENCLATURE. s 2

DESCRIPTION OF APPARATUS . 3

16-Foot Transonic Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . 3

P r o p u l s i o n M o d e l . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

F - 1 5 A i r p l a n e . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

Comparison of P r o p u l s i o n Model and Airplane . . . . . . . . . . . . . . . 5

Flow V i s u a l i z a t i o n Model . . . . . . . . . . . . . . . . . . . . . . . . 5

INSTRUMENTATION 5

TESTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

P r o p u l s i o n Model T e s t s . . . . . . . . . . . . . . . . . . . . . . . . . 6

F l i q h t T e s t s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

W a t e r T u n n e l T e s t s . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

D A T A R E D U C T I O N 7

Axial Force C o e f f i c i e n t . . . . . . . . . . . . . . . . . . . . . . . . . 7

Boundary Layer P r o f i l e s . . . . . . . . . . . . . . . . . . . . . . . . . 8

Nozzle P r e s s u r e R a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . 8

UNCERTAINTY 8 M o d e l . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

A i r p l a n e . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

RESULTS AND DISCUSSION 9 Comparison of Afterbody P r e s s u r e C o e f f i c i e n t D i s t r i b u t i o n s . . . . . . . 9

E f f e c t of Mach Number . . . . . . . . . . . . . . . . . . . . . . . 10

E f f e c t of Angle of Attack . . . . . . . . . . . . . . . . . . . . . 11

E f f e c t of L e f t Nozzle Boattail Angle . . . . . . . . . . . . . . . . 11

summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Comparison of Boundary Layer P r o f i l e s and T h i c k n e s s e s . . . . . . . . . . 11 E f f e c t Of Angle O f Attack 1 2

Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

iii Page . . . . . . . . . . . . . . . . . . . . . . . . . . .

Flow v i s u a l i z a t i o n 13 Comparison of Afterbody and Nozzle P r e s s u r e C o e f f i c i e n t D i s t r i b u t i o n s . . . . . . . . . . . . . . . . . . 14

E f f e c t of Mach Number . . . . . . . . . . . . . . . . . . . . . . . 14

E f f e c t of Angle of Attack . . . . . . . . . . . . . . . . . . . . . 16

. . . . . . . . . . . . . . . .

E f f e c t of L e f t Nozzle B o a t t a i l Angle 16

Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

Comparison of P r e s s u r e C o e f f i c i e n t D i s t r i b u t i o n s

on t h e Nozzle E x t e r n a l S u r f a c e . . . . . . . . . . . . . . . . . . . . 16

E f f e c t of Mach Number . . . . . . . . . . . . . . . . . . . . . . . 16

E f f e c t of Angle of Attack . . . . . . . . . . . . . . . . . . . . . 17

E f f e c t of L e f t Nozzle Boattail Angle . . . . . . . . . . . . . . . . 17

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Summary 17

Comparison of Recompression on t h e Nozzle E x t e r n a l S u r f a c e . . . . . . . 17

E f f e c t of Mach Number . . . . . . . . . . . . . . . . . . . . . . . 18

E f f e c t of L e f t Nozzle B o a t t a i l Angle . . . . . . . . . . . . . . . . 19

Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

Comparison of Nozzle A x i a l Force C o e f f i c i e n t . . . . . . . . . . . . . . 19

E f f e c t of Mach Number and L e f t Nozzle B o a t t a i l Angle . . . . . . . . 19

E f f e c t of Angle of Attack . . . . . . . . . . . . . . . . . . . . . 20

E f f e c t of Nozzle P r e s s u r e R a t i o . . . . . . . . . . . . . . . . . . 20

E f f e c t of Reynolds Number . . . . . . . . . . . . . . . . . . . . . 21

E f f e c t of Mach Number on t h e S e n s i t i v i t y of A x i a l

Force C o e f f i c i e n t t o Reynolds Number . . . . . . . . . . . . . . . 22

Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

CONCLUSIONS . e . 22

APPENDIX -DERIVATION OF PITOT PRESSURE RATIO I N THE BOUNDARY LAYER

FOR AN ASSUMED VELOCITY RATIO DISTRIBUTION . . . . . . . . . . . 24

REFERENCES 26 i v S U M M A R Y proximity of t h e n o z z l e s , a f t e r b o d y , and t a i l s . V a r i a b l e geometry and s e n s i t i v - Afterbody and nozzle p r e s s u r e s i t y of t h e l o c a l flows t o Mach number, measured on a 1/12-scale model and i n Reynolds number, a n g l e of a t t a c k , and t h e f l i g h t on a t w i n - j e t f i g h t e r a i r c r a f t h o t j e t e f f l u x a r e a d d i t i o n a l f a c t o r s .

were compared a s Mach number v a r i e d from S i m i l a r flow c o m p l e x i t i e s e x i s t i n t h e 0.6 t o 1 . 2 , Reynolds number v a r i e d from i n l e t forebody r e g i o n . Accordingly, 17.5 m i l l i o n t o 302.5 m i l l i o n , and a n g l e e x p e r i m e n t a l r e s e a r c h programs a r e neces- of a t t a c k v a r i e d from l o t o 7O.

s a r y t o a c q u i r e a better understanding of p r o p u l s i o n system and a i r f r a m e flow A t Mach 0.6 and 0.8, nozzle p r e s s u r e i n t e r f e r e n c e .

c o e f f i c i e n t d i s t r i b u t i o n s and nozzle a x i a l f o r c e c o e f f i c i e n t s agreed and To improve t h e technology base f o r showed good recompression. f u t u r e f i g h t e r a i r c r a f t and t o a c q u i r e a better understanding of p r o p u l s i o n system A t Mach 0.9 and 1.2, flow complex- and airframe f l o w i n t e r f e r e n c e , t h e U.S.

i t y caused a loss i n recompression A i r Force and NASA have completed a n f o r both f l i g h t and wind t u n n e l nozzle e x p e r i m e n t a l program i n v o l v i n g wind tun- d a t a . The f l i g h t d a t a e x h i b i t e d less n e l tests u s i n g s u b s c a l e F-15 a i r p l a n e models and f l i g h t tests w i t h t h e f u l l - n e g a t i v e v a l u e s of p r e s s u r e c o e f f i c i e n t and lower a x i a l f o r c e c o e f f i c i e n t s t h a n scale F-15 a i r p l a n e . The objective w a s d i d t h e wind t u n n e l d a t a . Reynolds t o create d a t a b a s e s from which compari- number a f f e c t s were noted o n l y a t t h e s e s o n s of wind t u n n e l and f l i g h t t e s t d a t a Mach numbers. J e t temperature and mass c o u l d be made. Reference 1 p r e s e n t s f l u x r a t i o d i d n o t a f f e c t t h e compari- s e l e c t e d r e s u l t s of t h e o v e r a l l program, s o n s of n o z z l e a x i a l f o r c e c o e f f i c i e n t . and r e f e r e n c e 2 p r e s e n t s a comparison o f measurements made i n t h e i n l e t a i r - A t s u b s o n i c s p e e d s , t h e l e v e l s of frame region.

p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s on t h e upper f u s e l a g e and lower n a c e l l e s u r f a c e s For t h e nozzle-afterbody r e g i o n , wind f o r f l i g h t w e r e less n e g a t i v e t h a n t h o s e t u n n e l tests were conducted a t t h e NASA f o r t h e model.

Langley Research Center u s i n g a propul- s i o n model, and f l i g h t tests w e r e con- The model boundary l a y e r t h i c k n e s s a t ducted a t t h e Dryden F l i g h t Research t h e a f t r a k e s t a t i o n exceeded t h a t f o r F a c i l i t y of t h e NASA A m e s Research Center.

t h e forward r a k e s t a t i o n and i n c r e a s e d The tests were c a r e f u l l y c o o r d i n a t e d so w i t h i n c r e a s i n g a n g l e of attack. The t h a t t h e test c o n f i g u r a t i o n s , instrumenta- f l i g h t boundary l a y e r t h i c k n e s s a t t h e t i o n l o c a t i o n s , and test c o n d i t i o n s of t h e model and a i r c r a f t were as similar as a f t rake s t a t i o n w a s less than t h a t f o r t h e forward r a k e s t a t i o n and decreased possible. This w a s done t o reduce d i f - w i t h i n c r e a s i n g a n g l e of a t t a c k .

f e r e n c e s i n the t e s t r e s u l t s t h a t could be a t t r i b u t e d t o t h e s e f a c t o r s . Refer- INTRODUCTION ences 3 t o 7 p r e s e n t r e s u l t s from t h e model and f l i g h t tests of t h e nozzle and I n t e g r a t i n g a p r o p u l s i o n system i n t o a f t e r b o d y r e g i o n t h a t were p r e v i o u s l y o b t a i n e d from t h i s program. T e s t s were a new t w i n - j e t f i g h t e r a i r c r a f t c o n t i n u e s a l s o conducted i n a water t u n n e l u s i n g a t o r e q u i r e e x t e n s i v e development efforts.

separate flow v i s u a l i z a t i o n model t o Blending t h e i n l e t s and n o z z l e s w i t h t h e p r o v i d e a d d i t i o n a l t e s t d a t a .

a i r f r a m e f r e q u e n t l y creates flow i n t e r - f e r e n c e s n o t amenable t o a n a l y t i c a l pre- The purpose of t h i s r e p o r t i s t o com- d i c t i o n and f o r which experimental d a t a may be lacking. I n t h e nozzle-afterbody p a r e and a n a l y z e model and f l i g h t d a t a o b t a i n e d €or t h e nozzle and a f t e r b o d y r e g i o n , s e v e r a l f a c t o r s c o n t r i b u t e t o f l o w i n t e r f e r e n c e . They i n c l u d e close r e g i o n s of t h e F-15 a i r p l a n e . These d a t a l e f t nozzle a x i a l f o r c e c o e f f i - C a i n c l u d e s u r f a c e pressures on t h e a f t c i e n t r e f e r e n c e d t o wing area, f u s e l a g e and nozzle, boundary l a y e r s u r - veys on t h e upper n a c e l l e s , and flow v i s u a l i z a t i o n on t h e upper fuselage.

Excluding t h e low-speed water t u n n e l tests, t h e Mach number range of t h e cor- r e l a t i o n extended from 0 . 6 t o 1 . 2 over an pressure c o e f f i c i e n t , cP angle-of-attack range of l o t o 7O. The t e s t Reynolds number range extended from 1 7 . 5 m i l l i o n f o r t h e p r o p u l s i o n model t o 3 0 2 . 5 m i l l i o n i n f l i g h t , based on t h e l e n g t h of t h e f u s e l a g e . S i d e s l i p a n g l e p r e s s u r e c o e f f i c i e n t a t s o n i c w a s e s s e n t i a l l y zero.

CP* speed, To provide a range of test v a l u e s , 2 3.5 s e v e r a l nozzle b o a t t a i l a n g l e s were i n -

0 . 7 5 4 7 (1 + 0.2MW )

- 1.4286

cluded i n t h e comparison. A t subsonic

------------ I_--

speeds, t h e l a r g e s t boattail a n g l e w a s a M W nominal 18.4O, corresponding t o t h e d r y o r m i l i t a r y power t h r o t t l e s e t t i n g and average v a l u e of pressure coef- t h e smallest opening a t t h e nozzle e x i t . ‘Pavg f i c i e n t , f o r either f i r s t o r A l s o included w a s a nominal b o a t t a i l l a s t nozzle o r i f i c e a n g l e of 1 5 . 1 ° , corresponding t o a l o w a f t e r b u r n i n g t h r o t t l e s e t t i n g and a F.S f u s e l a g e s t a t i o n , c m larger n o z z l e opening. A t Mach number

M = 1 . 2, t h e b o a t t a i l a n g l e w a s a nom-

H h e i g h t of boundary l a y e r r a k e i n a l 7.7O, corresponding t o a h i g h e r probe above f u s e l a g e s u r - a f t e r b u r n i n g t h r o t t l e s e t t i n g and a f a c e , c m s t i l l l a r g e r n o z z l e opening. Nozzle p r e s s u r e r a t i o v a r i e d from u n i t y ( j e t p r e s s u r e a l t i t u d e , geopoten- HP o f f ) t o a b o u t 7. The e f f e c t s of t h e t i a l , m major test v a r i a b l e s on each compari- s o n were s y s t e m a t i c a l l y i n v e s t i g a t e d d i s t a n c e from a i r p l a n e nose t o L a and analyzed.

end of tailboom, 1 9 . 0 5 m NOMENCLATURE d i s t a n c e from model nose t o end Lm o f tailboom, 158.689 c m p r o j e c t e d area i n t h e a x i a l An d i r e c t i o n a s s i g n e d t o each of M Mach number i n boundary l a y e r a t t h e 42 n o z z l e s u r f a c e pres- h e i g h t H s u r e s , cm2 MFRA r a t i o of j e t mass f l u x t o f r e e - BTL b o a t t a i l chord a n g l e of the l e f t stream mass f l u x , a i r p l a n e n o z z l e f o r a i r p l a n e , o r ter- minal b o a t t a i l a n g l e f o r M F R M r a t i o of j e t mass f l u x t o f r e e - model, deg stream mass f l u x , model BTR b o a t t a i l chord a n g l e of t h e free-stream Mach number; i n t h i s M W r i g h t nozzle f o r a i r p l a n e , or r e p o r t , also assumed t o be t h e t e r m i n a l boattail a n g l e f o r Mach number a t t h e edge of the model, deg boundary l a y e r C c o n s t a n t l e f t nozzle t o t a l p r e s s u r e t o t a l temperature i n t h e f r e e T t Q Y d i v i d e d by P, stream, K U local v e l o c i t y i n t h e boundary s u r f a c e s t a t i c p r e s s u r e , N/cm2 l a y e r a t h e i g h t H , m/sec p i t o t p r e s s u r e i n boundary l a y e r f ree-s tream v e l o c i t y ; i n t h i s UQY a t h e i g h t H , N / c m 2 report, a l s o assumed t o be t h e local v e l o c i t y i n t h e boundary free-stream s t a t i c p r e s s u r e , l a y e r a t a given boundary N / c m 2 l a y e r t h i c k n e s s 6 , m/sec t o t a l p r e s s u r e i n t h e boundary X d i s t a n c e from a i r p l a n e nose, m, l a y e r a t h e i g h t H , N / c m 2 and d i s t a n c e from model nose, c m free-stream t o t a l pressure, nondimensional d i s t a n c e from x/L N / c m 2 nose f o r t h e a i r p l a n e ( X / L a ) and t h e model ( X / L m ) p i t o t p r e s s u r e r a t i o i n t h e boundary l a y e r a t h e i g h t H , Y t r a n s v e r s e d i s t a n c e from a i r p l a n e N / c m 2 p l a n e of symmetry, m, and d i s t a n c e from model p l a n e of free-stream dynamic p r e s s u r e , symmetry, c m N / c m 2 nondimensional t r a n s v e r s e dis- y/L gas c o n s t a n t t a n c e from f u s e l a g e plane of symmetry f o r t h e a i r p l a n e Reynolds number ( Y f i a ) and f o r t h e model ( Y f i , ) Reynolds number f o r a i r p l a n e , a a n g l e of attack, deg based on L a r a t i o of s p e c i f i c h e a t s Y Reynolds number f o r model, based on Lm A i n c r e m e n t a l change wing area ( a i r p l a n e , 55.839 m2; a measure of average recompres- AcP s i o n on t h e e x t e r n a l s u r f a c e model, 0.3923 m2) of t h e nozzle static temperature i n boundary boundary l a y e r t h i c k n e s s , c m l a y e r a t h e i g h t H , K c i r c u m f e r e n t i a l a n g l e around the

+

t o t a l temperature of j e t l e f t nozzle measured clockwise exhaust, a i r p l a n e , K when viewed from t h e rear, deg t o t a l temperature of j e t DESCRIPTION O F APPARATUS e x h a u s t , model, K 16-Foot T r a n s o n i c Wind Tunnel s t a t i c temperature i n f r e e stream; i n t h i s r e p o r t , also Wind t u n n e l tests compared i n t h i s assumed t o be t h e s t a t i c t e m - r e p o r t were conducted i n t h e NASA Langley p e r a t u r e a t t h e edge of t h e l6-Foot Transonic Wind Tunnel, which is boundary l a y e r , K a s i n g l e - r e t u r n , continuous-flow, atmos- d e t a i l s of t h e model are given i n r e f e r - p h e r i c wind t u n n e l w i t h a s l o t t e d o c t a - e n c e s 3 and 7.

g o n a l test s e c t i o n measuring 4.8 m d i a - m e t r i c a l l y t o m i d f l a t c e n t e r l i n e . With F i g u r e 4 shows geometric d e t a i l s of t h e a i d of a compressor system, which t h e t h r e e test nozzles used i n t h e cor- draws a i r o u t through s l o t s i n t h e test r e l a t i o n : t h e m i l i t a r y power nozzle s e c t i o n f o r Mach numbers g r e a t e r than ( f i g . 4 ( a ) ) with a t e r m i n a l b o a t t a i l a n g l e 1.05, the test s e c t i o n a i r s p e e d i s con- of 18.4O and t h e t w o p a r t i a l a f t e r b u r n i n g t i n u o u s l y v a r i a b l e from subsonic speeds power n o z z l e s w i t h t e r m i n a l b o a t t a i l a n g l e s up t o Mach 1.3. A d d i t i o n a l d e t a i l s of of 15.1O ( f i g . 4 ( b ) ) and 7.7O ( f i g . 4 ( c ) ) .

t h e wind t u n n e l and i t s o p e r a t i o n are The i n t e r n a l geometry and area r a t i o of given i n r e f e r e n c e 8. t h e s e n o z z l e s simulated t h o s e f o r t h e f u l l - scale nozzles.

P r o p u l s i o n Model F-15 A i r p l a n e F i g u r e 1 is a photograph of t h e 1 /12-scale F-15 a i r c r a f t p r o p u l s i o n The F-15 is an a i r - s u p e r i o r i t y model mounted on a s t r u t i n t h e NASA f i g h t e r a i r p l a n e capable of t r a n s o n i c Langley 16-Foot T r a n s o n i c Wind Tunnel, c r u i s e and s u p e r s o n i c dash t o speeds and f i g u r e 2 shows geometric d e t a i l s of g r e a t e r than Mach 2. D i s t i n g u i s h i n g air- t h e model. Room-temperature a i r a t high p l a n e f e a t u r e s i n c l u d e a high-mounted p r e s s u r e w a s used t o s i m u l a t e t h e j e t sweptback wing, twin v e r t i c a l s t a b i l i - exhaust. The a i r w a s ducted through t h e z e r s , and a large h o r i z o n t a l s t a b i l a t o r .

s u p p o r t s t r u t , r o u t e d i n t e r n a l l y i n t h e A photograph and a three-view drawing of model, and exhausted through t h e test t h e a i r p l a n e are shown i n f i g u r e s 5 and nozzles. There w a s no i n l e t flow through 6 , r e s p e c t i v e l y . The variable-geometry t h e model; f a i r i n g s w e r e placed on t h e i n l e t s use h o r i z o n t a l ramps mounted model where t h e i n l e t s were normally a t the wing l e v e l . Each i n l e t is l o c a t e d on t h e a i r p l a n e . The model w a s a u t o m a t i c a l l y c o n t r o l l e d u s i n g inde- t e s t e d w i t h and w i t h o u t a noseboom, pendent c o n t r o l l e r s ; however, t h e l e f t a f t e r b u r n e r f u e l v e n t f a i r i n g s , nozzle i n l e t w a s manually c o n t r o l l a b l e i n t h e f l a p a c t u a t o r f a i r i n g s , and a 0.127-cm t e s t a i r p l a n e .

s t e p upstream of each n o z z l e s i m u l a t i n g t h e engine bay v e n t on t h e a i r p l a n e . The powerplants f o r t h e F-15 a i r p l a n e The model c o n t a i n e d a six-component are two p r o t o t y p e P r a t t and Whitney F100- Pw-100 e n g i n e s t h a t are twin-spool, a f t e r b o d y b a l a n c e t h a t measured t h e n e t f o r c e and moment on a l l s u r f a c e s down- a f t e r b u r n i n g t u r b o f a n s i n t h e 110,000-N stream of t h e metric break ( f i g . 21, t h r u s t class. Both t h e f a n and the com- i n c l u d i n g t h e a f terbody, test n o z z l e s , p r e s s o r use v a r i a b l e geometry f o r h i g h and t a i l s u r f aces . performance and d i s t o r t i o n a t t e n u a t i o n .

The nominal bypass r a t i o i s approximately F i g u r e 3 is a photograph i l l u s t r a t - 0.7 f o r m i l i t a r y (maximum nonaugmented) i n g a f t e r b o d y d e t a i l s and t h e d r y , or power a t s e a - l e v e l , standard-day condi- m i l i t a r y , power nozzles. The a f t e r b o d y t i o n s . The engine c o n t r o l s c o n s i s t of a c o n s i s t s of c l o s e l y spaced n o z z l e s , hydromechanical u n i t and a s u p e r v i s o r y widely spaced tailbooms, and large ver- d i g i t a l c o n t r o l u n i t .

t i c a l and h o r i z o n t a l t a i l s u r f a c e s . A wedge-shaped i n t e r f a i r i n g is l o c a t e d F i g u r e 7 is a rear-view photograph between the t w o nozzles. The wedge h a s of t h e F-15 a i r p l a n e . Exhaust n o z z l e s are v a r i a b l e geometry w i t h a convergent- a 20° i n c l u d e d a n g l e and t e r m i n a t e s c l o s e d i v e r g e n t i n t e r n a l f l o w path. To v a r y t o t h e e x i t p l a n e of t h e nozzle. The t h e n o z z l e geometry, t h e engine c o n t r o l development of t h i s c o n f i g u r a t i o n is d i s c u s s e d i n r e f e r e n c e 9. A d d i t i o n a l c o n t i n u o u s l y maintains a scheduled rela- t i o n s h i p between t h e t h r o a t area and t h e t h e flow e x i t i n g from t h e vent and flow- I r a t i o of t h e e x i t area t o t h e t h r o a t i n g downstream over t h e nozzle. A s pre- area. Nozzle s u r f a c e s c o n s i s t of v a r i - v i o u s l y mentioned, t h e model jets simu- a b l e f l a p s and seals t h a t s l i d e i n a cir- l a t e d t h e a i r p l a n e ' s h o t j e t exhaust with c u m f e r e n t i a l d i r e c t i o n t o maintain a room-temperature air.

c i r c u l a r i n t e r n a l p a t h a t t h e t h r o a t and e x i t as t h e geometry is varied. Figure 11 compares t h e a i r p l a n e and model nozzle e x t e r n a l s u r f a c e s . The wind t u n n e l model nozzles were smoothly Details of t h e nozzle geometry are machined, whereas t h e a i r p l a n e nozzle shown i n f i g u r e 8. The three-dimensional had v a r i a b l e f l a p s and seals. The flaps f u s e l a g e s u r f a c e ends a t nondimensional and seals r e s u l t e d i n gaps n e a r t h e d i s t a n c e X/L = 0.886. A f t of t h i s sta- t r a i l i n g edge t h a t v a r i e d w i t h nozzle t i o n , t h e exposed nozzle is axisymmetric.

b o a t t a i l angle.

The nozzle i s f i x e d geometry forward of X/L = 0.900 where t h e t e r m i n a l a n g l e is Flow V i s u a l i z a t i o n M o d e l 13.2'. Nozzle geometry v a r i e s w i t h p o w e r s e t t i n g downstream of X/L = 0.900, T h i s B e c a u s e t h e d i f f e r e n c e s i n t h e i n l e t v a r i a t i o n r e s u l t s i n a d i s c r e t e change c o n f i g u r a t i o n s of t h e propulsion model i n nozzle e x t e r n a l shape a t X/L = 0.900.

and a i r p l a n e w e r e thought t o a f f e c t t h e The nozzle b o a t t a i l a n g l e ranges from boundary l a y e r p r o f i l e s on t h e upper a b o u t 18.4O a t m i l i t a r y p o w e r t o a b o u t f u s e l a g e , a flow v i s u a l i z a t i o n s t u d y w a s 1' t o 2' with f u l l a f t e r b u r n i n g power.

conducted. The 1/48-scale f l o w v i s u a l - I n determining t h e nozzle b o a t t a i l i z a t i o n model of t h e F-15 a i r p l a n e used a n g l e , t h e nozzle f l a p s are assumed t o i n r e f e r e n c e 10 was s u i t a b l y modified t o be s t r a i g h t l i n e s . A d d i t i o n a l F-15 air- s i m u l a t e t h e p r o p u l s i o n model and t h e p l a n e d e t a i l s are given i n r e f e r e n c e 6.

a i r p l a n e . Figure 1 2 ( a ) shows t h e f a i r e d - i n l e t c o n f i g u r a t i o n , and f i g u r e 1 2 ( b ) Comparison of P r o p u l s i o n Model shows t h e f l o w i n g - i n l e t c o n f i g u r a t i o n .

and Airplane With t h e i n l e t f a i r i n g i n s t a l l e d , t h e r e w a s no i n t e r n a l flow. The model w a s Although t h e overall program w a s equipped w i t h s u r f a c e o r i f i c e s through c l o s e l y c o o r d i n a t e d , t h e r e were some which v a r i o u s c o l o r e d dyes were e j e c t e d unavoidable d i f f e r e n c e s between t h e pro- from i n s i d e t h e model t o flow downstream p u l s i o n model and t h e a i r p l a n e . Figure 9 o v e r t h e f u s e l a g e s u r f a c e , t h u s i l l u s - compares t h e model and a i r p l a n e i n l e t s .

t r a t i n g t h e flow p a t t e r n s .

For t h e model ( f i g . 9 ( a ) ) , t h e i n l e t was r e p l a c e d w i t h a f a i r i n g t h a t extended INSTRUMENTATION forward on t h e forebody. Figures 9 ( b ) and 9 ( c ) show one o€ t h e airplane's F i g u r e 13 shows t h e l o c a t i o n s of t h e i n l e t s and its e x t e r n a l v a r i a b l e geom- s u r f a c e p r e s s u r e o r i f i c e s on t h e upper e t r y , which c o n s i s t e d of a v a r i a b l e - and lower f u s e l a g e of t h e model and air- a n g l e c o w l and a two-position bypass plane. The l o c a t i o n s are shown i n terms door. The model i n l e t simulated a n of nondimensional d i s t a n c e s X / L , measured u n d e f l e c t e d cowl; t h e r e f o r e , as shown downstream from t h e nose, and Y / L , meas- i n f i g u r e 9 ( c ) , f l i g h t tests were made ured from t h e p l a n e of symmetry. Table 1 w i t h t h e c o w l a l i g n e d with t h e n a c e l l e lists t h e l o c a t i o n s i n t a b u l a r form and s u r f a c e and w i t h t h e bypass door closed.

shows t h e good agreement between t h e model and a i r p l a n e f o r s u r f a c e p r e s s u r e o r i f i c e Photographs of t h e engine compartment l o c a t i o n s , N o t e t h a t on t h e upper na- celle, t h e a i r p l a n e had f o u r more o r i f i c e s v e n t on t h e a i r p l a n e and i t s s i m u l a t i o n t h a n t h e model. These were i n s t a l l e d f o r on t h e model are shown i n f i g u r e 10. The t h e companion i n l e t - a i r f r a m e i n t e r a c t i o n s model simulated t h e rearward-facing s t e p c r e a t e d by t h e v e n t b u t d i d n o t s i m u l a t e program ( r e f . 2 ) .

Figure 1 3 ( a ) a l s o shows t h e l o c a t i o n s tests y i e l d e d a m a t r i x of Mach number and of t h e boundary l a y e r rake on each of angle-of-attack test p o i n t s using s e v e r a l t h e upper n a c e l l e s . Drawings and photo- t e s t nozzles. These w e r e t o be matched graphs of t h e model and a i r p l a n e rakes d u r i n q t h e f l i g h t tests. The test p o i n t s shown i n f i g u r e 14 d i s p l a y d i f f e r e n c e s

were attempted i n f l i g h t , b u t several

i n t h e rake c o n f i g u r a t i o n s . The model could n o t be flown because of q l i m i t a - rakes ( f i g s . 1 4 ( a ) and 1 4 ( b ) ) used f i v e t i o n s and t h e i n a b i l i t y t o s t a b i l i z e p i t o t t u b e s , with t h e uppermost p i t o t t h e a i r p l a n e t r a n s o n i c a l l y because of t u b e of each rake about 2.5 c m above t h e e x c e s s t h r u s t a t t h e p a r t i a l a f t e r b u r n i n g model s u r f a c e . T h i s value scales t o power s e t t i n g s .

a b o u t 30.0 c m (when m u l t i p l i e d by 12) t o compare w i t h t h e a i r p l a n e rakes. Heiqhts Following the f l i g h t tests, t h e sec- of t h e i n d i v i d u a l p i t o t tubes v a r i e d ond model test series was conducted. I n s l i g h t l y between the two model rakes. t h i s series t h e e x a c t test p o i n t s t h a t The a i r p l a n e r a k e s ( f i g s . 1 4 ( c ) and were flown were d u p l i c a t e d i n t h e t u n n e l 1 4 ( d ) ) used 10 p i t o t t u b e s , with t h e u s i n g smaller nozzles. I n a d d i t i o n , uppermost tube 38.4 c m above t h e air- a noseboom w a s added and o t h e r model plane s u r f a c e . The model r a k e s used changes were made t o b e t t e r match the a tube t h a t measured s t a t i c p r e s s u r e a i r p l a n e c o n f i g u r a t i o n , The d a t a from a t t h e t o p of t h e rake; t h e tube w a s t h e s e tests were used i n t h e comparison l a c k i n g on t h e a i r p l a n e rake.

p r e s e n t e d i n t h i s paper.

Figure 15 shows t h e angular loca- Model d a t a were taken a t Mach 0.60, t i o n s of t h e e i g h t r o w s of s u r f a c e 0.80, 0.87, 0.90, and 1.2. D a t a w e r e p r e s s u r e o r i f i c e s on t h e l e f t nozzles o b t a i n e d with t h e boundary l a y e r r a k e s o f t h e model and a i r p l a n e . The a x i a l mounted and removed. Model a n g l e of l o c a t i o n s , i n terms of X / L , of t h e a t t a c k v a r i e d from - 2 O t o 7O. The r a t i o 42 p r e s s u r e o r i f i c e s are qiven i n o f j e t t o t a l p r e s s u r e t o free-stream t a b l e 2 f o r each of the angular rows. s t a t i c p r e s s u r e w a s v a r i e d from 1 , j e t The model and a i r p l a n e o r i f i c e loca- o f f , t o about 7. H o r i z o n t a l t a i l or t i o n s show e x c e l l e n t agreement.

s t a b i l a t o r a n g l e v a r i e d from 2.5O to -4.0° l e a d i n g edge down. For t h e com- I n a d d i t i o n t o t h e f u s e l a g e , bound- p a r i s o n , d a t a a t h o r i z o n t a l t a i l a n g l e a r y l a y e r , and nozzle p r e s s u r e s , numer- s e t t i n g s of -2O, O o , and 2 O were used.

ous o t h e r parameters were measured Reynolds number v a r i e d from 17.5 m i l l i o n d u r i n g t h e model and f l i g h t tests. For t o 21 m i l l i o n f o r t h e Mach number range t h e model tests, t h e s e included t h e t e s t e d . Force and moment d a t a were n o t free-stream parameters, h o r i z o n t a l t a i l obtained. A d d i t i o n a l t e s t information d e f l e c t i o n , and j e t nozzle parameters. i s given i n r e f e r e n c e s 3 and 7.

For t h e f l i g h t tests, free-stream parameters, s u r f a c e d e f l e c t i o n s , i n l e t F l i g h t T e s t s parameters, and engine parameters were measured. A d e t a i l e d d i s c u s s i o n of Although t h e Mach number range f o r t h e s e measurements is presented i n t h i s comparison i s 0.60 t o 1.2, t h e f l i g h t test Mach number range extended r e f e r e n c e s 3, 6, and 7.

from 0.60 t o 2.0 ( r e f . 6 ) . The f l i g h t tests were conducted t o o b t a i n q u a s i - TESTS s t e a d y - s t a t e d a t a t o match t h e model Propulsion Model T e s t s d a t a . To a i d t h e p i l o t i n s t a b i l i z i n g on a test p o i n t , i n - f l i g h t measurements Two test series were planned and of a i r p l a n e parameters were t r a n s m i t t e d conducted t o o b t a i n p r e s s u r e d i s t r i b u - i n real t i m e by way of a t e l e m e t r y down- t i o n s f o r t h e model d a t a base. The l i n k t o t h e ground-station computer.

f i r s t series conducted b e f o r e t h e f l i g h t Ground-computed parameters i n real t i m e were uplinked t o t h e p i l o t ' s c o n s o l e , wind t u n n e l tests, t h i s was n o t p o s s i b l e allowing him t o make r a p i d c o r r e c t i o n s f o r t h e f l i g h t tests. The r a t i o was t o h i s f l i g h t p a t h and t h u s s t a b i l i z e on g e n e r a l l y f i x e d by t h e t e s t p o i n t . Fig- t h e d e s i r e d t e s t p o i n t . This technique u r e 18, o b t a i n e d from r e f e r e n c e 6 , shows proved s u c c e s s f u l i n s t a b i l i z i n g Mach a t y p i c a l v a r i a t i o n of l e f t nozzle t o t a l number, a l t i t u d e , a n g l e of a t t a c k , and p r e s s u r e (NPRL) with M , f o r a number of a n g l e of s i d e s l i p a t t h e d e s i r e d values.

t e s t p o i n t s . The d a t a r e p r e s e n t s e v e r a l Details of t h i s t e c h n i q u e are given i n b o a t t a i l a n g l e s .

r e f e r e n c e 1 1 .

The l e f t i n l e t cowl w a s s e t t o 0' and To a c h i e v e a v a r i a t i o n i n Reynolds t h e bypass door c l o s e d t o s i m u l a t e t h e number, nominal t e s t a l t i t u d e s of 6100, f a i r e d - i n l e t c o n f i g u r a t i o n of t h e model.

10,700, and 13,700 m w e r e flown. A s However, t h e r i g h t cowl and bypass door shown i n f i g u r e 16, t h e f l i g h t test were i n t h e a u t o m a t i c c o n t r o l mode, sub- Reynolds number range f o r t h e d a t a j e c t i n g t h e upper f u s e l a g e flow f i e l d t o p r e s e n t e d h e r e i n extended from less s l i g h t geometric asymmetry.

t h a n 80 m i l l i o n t o 302.5 m i l l i o n . Com- p a r i n g t h e Reynolds number test ranges I n f l y i n g t h e test p o i n t s , t h e sta- f o r model and f l i g h t r e v e a l s t h e l a r g e b i l a t o r and r u d d e r s w e r e v a r i e d as r a n g e i n Reynolds number t h a t e x i s t e d required. The average s t a b i l a t o r posi- f o r t h e comparison.

t i o n varied from 2.1 O l e a d i n g edge up t o 3.2O l e a d i n g edge down, and t h e average The b o a t t a i l a n g l e of t h e instrumen- r u d d e r t r a i l i n g edge p o s i t i o n v a r i e d from t e d l e f t nozzle w a s c a l c u l a t e d i n real 2.5O a i r p l a n e nose r i g h t t o 1.4O a i r p l a n e t i m e i n t h e g r o u n d - s t a t i o n computer using nose l e f t .

t h e downlinked measurement of t h e nozzle t h r o a t area. T h i s v a l u e w a s t h e n up- Water Tunnel T e s t s l i n k e d t o a p i l o t d i s p l a y . The r e l a t i o n between b o a t t a i l a n g l e and t h r o a t area The flow v i s u a l i z a t i o n model w a s w a s o b t a i n e d from a ground c a l i b r a t i o n t e s t e d i n t h e w a t e r t u n n e l d e s c r i b e d i n p r i o r t o t h e f l i g h t tests. For each test r e f e r e n c e 10. The purpose of t h e tests p o i n t , t h e p i l o t set t h e l e f t nozzle w a s t o i d e n t i f y s i m i l a r i t i e s and d i f f e r - b o a t t a i l a n g l e by a d j u s t i n g t h e t h r o t t l e e n c e s i n t h e flow p a t t e r n s on t h e upper and o b s e r v i n g h i s d i s p l a y . Figure 17 f u s e l a g e caused by t h e d i f f e r e n c e i n shows t y p i c a l v a r i a t i o n s of t e s t v a l u e s i n l e t c o n f i g u r a t i o n s . Colored dye w a s of t h e l e f t n o z z l e b o a t t a i l a n g l e as a e j e c t e d from o r i f i c e s on t h e forebody, f u n c t i o n of Mach number. The four boat- n a c e l l e , and wing l e a d i n g edges t o create t a i l a n g l e s are i d e n t i f i e d by t h e h o r i - f l o w p a t t e r n s t h a t were photographed z o n t a l l i n e s d e s i g n a t e d a s 18.4O, 14.6', under s t e a d y test c o n d i t i o n s . The a n g l e 9 . 5 O , and 7.7O. From a free-stream Mach of a t t a c k extended from O o t o I O o , and number of 0.6 t o 1.2, w i t h a boat- t h e a n g l e of s i d e s l i p w a s O o . The t e s t t a i l chord a n g l e BTL of 18.4O, 14.6', Reynolds number w a s 40,000, based on t h e and 7.701 t h e d a t a show t h e p i l o t w a s l e n g t h of t h e model.

a b l e t o set t h e b o a t t a i l a n g l e t o w i t h i n k O . 5 O of t h e t a r g e t v a l u e s t h e m a j o r i t y D A T A REDUCTION of t h e t i m e .

Axial Force C o e f f i c i e n t I n most cases t h e r i g h t engine power s e t t i n g s , and hence t h e nozzle b o a t t a i l A l l d a t a p r e s e n t e d i n t h i s r e p o r t a n g l e , c l o s e l y matched t h o s e f o r t h e l e f t were o b t a i n e d d u r i n g s t a b i l i z e d test engine. Although t h e nozzle p r e s s u r e c o n d i t i o n s . S u r f a c e p r e s s u r e s w e r e r a t i o was s y s t e m a t i c a l l y v a r i e d i n t h e reduced t o p r e s s u r e c o e f f i c i e n t s Cp.

To determine n o z z l e a x i a l f o r c e coef- of p i t o t p r e s s u r e r a t i o s i n c e t h e i n s t r u - f i c i e n t , each of t h e 4 2 o r i f i c e s w a s mentation w a s i n s u f f i c i e n t t o c a l c u l a t e a s s i g n e d t h e a p p r o p r i a t e a f t - f a c i n g v e l o c i t y r a t i o p r o f i l e s .

p r o j e c t e d area predetermined from a n o z z l e c a l i b r a t i o n . T h i s area v a r i e d Nozzle P r e s s u r e R a t i o w i t h t h e nozzle b o a t t a i l angle. The a x i a l f o r c e c o e f f i c i e n t w a s obtained For t h e model, n o z z l e p r e s s u r e r a t i o w a s o b t a i n e d by d i v i d i n g t h e measured by summing t h e 4 2 p r o d u c t s of t h e n o z z l e e n t r a n c e pressure by free-stream l o c a l area and t h e p r e s s u r e c o e f f i - c i e n t and then d i v i d i n g t h a t sum by s t a t i c p r e s s u r e P-. For t h e a i r p l a n e , t h e winq area.

t h e appropriate F100-PW-100 engine s t a t u s deck w a s used t o o b t a i n n o z z l e Boundary Layer P r o f i l e s e n t r a n c e p r e s s u r e as w e l l as o t h e r e n g i n e parameters. These were o b t a i n e d P r o f i l e s of p i t o t pressure r a t i o were as a f u n c t i o n of Mach number, pressure o b t a i n e d from t h e boundary l a y e r rakes.

a l t i t u d e , and power s e t t i n g f o r a given I n d i v i d u a l p i t o t p r e s s u r e s w e r e d i v i d e d test p o i n t .

by t h e r e s p e c t i v e free-stream t o t a l pres- s u r e p t from model o r f l i g h t test, and A d d i t i o n a l d i s c u s s i o n of t h e d a t a oo r e d u c t i o n procedures i s qiven i n r e f e r - were p l o t t e d as a f u n c t i o n of t h e r a t i o s ences 3, 5, and 6.

p i t o t t u b e p o s i t i o n . To compare model and f l i g h t d a t a , t h e p i t o t tube p o s i t i o n s UNCERTAINTY f o r the model rake w e r e m u l t i p l i e d by 1 2 because t h e model w a s 1 /12-scale s i z e .

Mode 1 Both sets of d a t a w e r e compared w i t h a r e f e r e n c e boundary l a y e r p r o f i l e . T h i s Accuracies of s e v e r a l parameters r e f e r e n c e p r o f i l e w a s c a l c u l a t e d usinq o b t a i n e d i n the wind t u n n e l tests were t h e boundary l a y e r t h i c k n e s s obtained determined by t h e root-sum-square method from t h e f l i g h t t e s t c o n d i t i o n and an f o r combining errors from independent assumed t u r b u l e n t v e l o c i t y d i s t r i b u t i o n .

sources. The e r r o r s are given i n t h e The d e r i v a t i o n of t h e e x p r e s s i o n f o r t h i s table below.

p r o f i l e is given i n t h e appendix.

Airplane Boundary l a y e r t h i c k n e s s e s f o r t h e f l i g h t d a t a were o b t a i n e d from v e l o c i t y Reference 6 p r e s e n t s a d e t a i l e d d i s - r a t i o p r o f i l e s as d i s c u s s e d i n r e f e r - c u s s i o n and a n a l y s i s of t h e e s t i m a t e d ence 6. For t h e model, t h e t h i c k n e s s e s e r r o r s f o r t h e measured and c a l c u l a t e d w e r e o b t a i n e d by i n s p e c t i n g t h e p r o f i l e s Ma ~ 0.60 0.80 0.90

I I I 1.2

Parameter f0.0016 ?r 0.001 6 f 0.001 7 f 0.0019 4 0

* 0.00006 f 0.00005 * 0.00009

* 0.00004

Ca

* 0.006 f 0.006 f 0.009

f 0.004 CP f 0.0006 f 0.0005 f 0.001 P,, N / c m 2 0.0003 0.001 4 *0.0014 f 0.0014 qm, N / c m 2 0.001 6 q u a n t i t i e s i n f l i g h t . The r e s u l t s of The test c o n d i t i o n s f o r t h e model t h i s a n a l y s i s are summarized i n f i g - and f l i g h t d a t a are p r e s e n t e d i n table 3 u r e 19, which i s a p l o t of t h e l e f t f o r s u r f a c e p r e s s u r e s and i n t a b l e 4 f o r n o z z l e Cp and l e f t nozzle a x i a l f o r c e boundary l a y e r p r o f i l e s . Table 5 lists t h e p i t o t p r e s s u r e r a t i o s from t h e c o e f f i c i e n t (C,) errors as f u n c t i o n s boundary l a y e r tests. Table 6 lists o f free-stream dynamic p r e s s u r e G.

t h e s e v e r a l boundary l a y e r t h i c k n e s s e s Both errors show a s t r o n g dependency f o r the f l i q h t boundary l a y e r p r o f i l e s

on s. The e r r o r s corresponding to

presented. Table 7 lists t h e d a t a f o u r nominal test p o i n t s are shown on f i g u r e s , c o n t e n t of t h e f i g u r e s , and t h e each error curve. The h i g h e s t errors t e s t c o n d i t i o n s s e l e c t e d from t a b l e s 3 and 4 used i n each of t h e f i g u r e s .

occur a t l o w q , , corresponding to l o w Table 7 a i d s i n r e l a t i n g t h e p l o t t e d and M , a t t h e h i g h e r a l t i t u d e s ( f o r example, t a b u l a t e d d a t a .

M , = 0.6 a t 10,700 m). Conversely, t h e l o w e s t errors occur a t h i g h Q, corres- P r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s ponding t o f l i g h t a t h i g h M , a t t h e on t h e e x t e r n a l s u r f a c e s of t h e a f t e r - body and nozzle are p l o t t e d a g a i n s t X/L.

lower a l t i t u d e s ( f o r example, M , = 0.9 Except f o r model test p o i n t s 21 t o 25 a t 6,000 m ) . For t h e s e latter con- of t a b l e 3, t h e s e d a t a were o b t a i n e d d i t i o n s , t h e e r r o r i n Cp is less than w i t h the boundry l a y e r r a k e s removed.

0.015, w h i l e t h e e r r o r i n C a remains A s i n d i c a t e d i n r e f e r e n c e 7, t h e a d d i t i o n less t h a n 0.0001. These errors approach of t h e boundary l a y e r rakes g e n e r a l l y d i d t h e model e r r o r s . n o t change t h e s u r f a c e p r e s s u r e d i s t r i b u - t i o n s on t h e l e f t a f t e r b o d y and nozzle.

RESULTS AND D I S C U S S I O N Comparison of Afterbody P r e s s u r e P r e s e n t a t i o n of R e s u l t s C o e f f i c i e n t D i s t r i b u t i o n s The r e s u l t s i n t h i s paper are pre- Afterbody p r e s s u r e c o e f f i c i e n t s e n t e d as follows: comparisons of d i s t r i b u t i o n s from model and f l i g h t were a f t e r b o d y p r e s s u r e c o e f f i c i e n t d i s t r i - analyzed and i n t e r p r e t e d . The local flow b u t i o n s , boundary l a y e r p r o f i l e s and f i e l d s w e r e then c l a s s i f i e d as being t h i c k n e s s e s , and flow v i s u a l i z a t i o n . e i t h e r compression o r expansion f i e l d s .

These are followed by comparisons of The compression o r expansion w a s caused afterbody and nozzle p r e s s u r e coef- by i n t e r f e r e n c e flows from nearby s u r - f i c i e n t d i s t r i b u t i o n s , p r e s s u r e coef- f a c e s as w e l l as the l o c a l s u r f a c e curva- f i c i e n t d i s t r i b u t i o n s on t h e nozzle t u r e . In f i g u r e 20, t h e flow f i e l d clas- e x t e r n a l s u r f a c e , recompression on t h e s i f i c a t i o n s are shown, along w i t h t h e n o z z l e , and nozzle a x i a l force. a d j a c e n t v e h i c l e components i n f l u e n c i n g them. On t h e upper f u s e l a g e , t h e up- To analyze t h e e f f e c t s of the stream expansion f i e l d ( f i g . 2 0 ( a ) ) is s e v e r a l test v a r i a b l e s on t h e com- b e l i e v e d t o be i n f l u e n c e d by t h e forward p a r i s o n s , t h e d a t a were organized and p o r t i o n of t h e wing. F a r t h e r downstream, p l o t t e d so t h a t o n l y one test v a r i a b l e t h e observed compression f i e l d , as i n t e r - changed while t h e o t h e r test v a r i a b l e s p r e t e d , i l l u s t r a t e s the e f f e c t s of t h e were r e l a t i v e l y c o n s t a n t . S i m i l a r i t i e s n a c e l l e c u r v a t u r e and wing t r a i l i n g edge and d i f f e r e n c e s between t h e model and shock waves. I n t h e a f t e r b o d y and nozzle f l i g h t d a t a , t h e r e f o r e , became f u n c t i o n s r e g i o n , e i t h e r an expansion-recompression of t h e p a r t i c u l a r test v a r i a b l e t h a t o r an expansion f i e l d is i n d i c a t e d .

changed. The test v a r i a b l e s analyzed i n I n t e r f e r e n c e e f f e c t s from t h e t a i l s u r - t h i s way were Mach number, a n g l e of f a c e s , tailboom, nozzle i n t e r f a i r i n g s , a t t a c k , l e f t nozzle b o a t t a i l a n g l e and a f t e r b o d y b o a t t a i l i n g , and v a r i a b l e - p r e s s u r e r a t i o , and Reynolds number.

geometry nozzles a l l i n f l u e n c e t h e com- I n f i q u r e 21 ( b ) , a t M , = 0.8, t h e p l e x flow e f f e c t s observed. Fiqure 2 0 ( b ) o v e r a l l p r e s s u r e d i s t r i b u t i o n t r e n d s f o r shows a similar i n t e r p r e t a t i o n f o r t h e both sets of d a t a resemble those a t lower fuselage.

M , m 0.6. On a l l t h r e e upper s u r f a c e s , t h e r e d u c t i o n s i n Cp f o r s t a t i o n s down- E f f e c t of Mach Number stream of X / L = 0.800 are probably due t o c o m p r e s s i b i l i t y e f f e c t s caused by t h e The e f f e c t of Mach number on t h e increase i n free-stream Mach number from comparison of a f t e r b o d y p r e s s u r e coef- M , 0.6 t o M , 0.8.

f i c i e n t d i s t r i b u t i o n s i s shown i n f i g - u r e 21. Data are presented a t a n q l e of a t t a c k a m 1' f o r BTL 18.4' a t subsonic On t h e lower s u r f a c e s , t h e model speeds and f o r l3TL = 7.7' a t I $ . , , m 1.2 f o r d a t a e x h i b i t minimal change from t h e d a t a a t M , = 0.6. The f l i q h t d a t a e x h i b i t a n NPFU test range of 2.0 t o 5.0.

s l i g h t l y more n e g a t i v e values of Cp t h a n I n f i g u r e 21 ( a ) t h e a f t e r b o d y pres- f o r M , m 0.6 a t a l l values of X / L . On sure c o e f f i c i e n t d i s t r i b u t i o n s f o r t h e lower n a c e l l e , flow l o s s e s caused M , = 0.6 and a 1 O may be used t o i l l u s - by t h e f u e l dump v e n t f o r both sets of t r a t e t h e i n t e r p r e t a t i o n s d i s c u s s e d i n d a t a are i n c r e a s e d compared t o those a t f i g u r e s 2 0 ( a ) and 2 0 ( b ) . Thus f o r t h e M , 0.6.

upper f u s e l a g e i n f i g u r e 21 ( a ) , both sets

of data are s i m i l a r and show t h e expan-

I n f i g u r e 21 ( c ) , as Mach number

s i o n and compression r e g i o n s between increases t o M, 0.9, c o m p r e s s i b i l i t y X/L = 0.456 and X/L = 0.684. These are e f f e c t s on t h e p r e s s u r e c o e f f i c i e n t d i s - followed by t h e expansion and subsequent t r i b u t i o n s are more pronounced. Local recompression t o near ambient p r e s s u r e Mach numbers become supersonic over t h e downstream of X / L = 0.900.

upper n a c e l l e and tailboom because t h e local flow expansions reduce t h e Cp The upper nacelle d a t a are similar t o t h e upper f u s e l a g e d a t a downstream of l e v e l s below t h e p r e s s u r e c o e f f i c i e n t X/L = 0.530 w i t h t h e n e g a t i v e g r a d i e n t s a t s o n i c speed (Cp*). The model d a t a steeper for f l i g h t . The g r e a t e r steep- do n o t become as n e g a t i v e as t h e f l i g h t n e s s i s b e l i e v e d t o be due to t h e closer d a t a . The s h a r p rises i n Cp a t x / L = proximity, and t h e r e f o r e s t r o n g e r i n f l u - 0.870 on t h e upper tailboom are a t t r i b - ence, of t h e expansion f i e l d of t h e u t e d to s t a n d i n g shock waves recompress- l e a d i n g edge of t h e v e r t i c a l t a i l . The i n g t h e flow; t h i s may be p a r t i a l l y flow f i e l d f o r t h e upper l e f t tailboom caused by t h e compression f i e l d of t h e shows a steep recompression t o near ambi- v e r t i c a l tail. The rise f o r t h e f l i g h t e n t p r e s s u r e a t t h e l a s t o r i f i c e . T h i s d a t a exceeds t h e rise f o r t h e model d a t a , i s b e l i e v e d t o be i n f l u e n c e d by t h e com- s u g g e s t i n g a h i g h e r local Mach number p r e s s i o n f i e l d of t h e t r a i l i n g edge of immediately preceding t h e shock wave f o r t h e vertical tail.

t h e a i r p l a n e than f o r t h e model.

The lower f u s e l a g e d i s t r i b u t i o n s f o r On t h e lower f u s e l a g e a t M , m 0.9, model and f l i g h t i l l u s t r a t e t h e expan- t h e f l i g h t d a t a e x h i b i t a loss of recom- sion-recompression f i e l d downstream of p r e s s i o n beyond X/L 0.800. Other x/L = 0.684. For t h e lower n a c e l l e , both changes f o r both sets of d a t a are minimal sets of d a t a a l s o show a compression e x c e p t f o r t h e flow l o s s e s caused by t h e followed by a n expansion. The f u e l dump f u e l dump v e n t , which i n c r e a s e from t h o s e v e n t a t X/L = 0.800 causes o n l y a s l i g h t a t M , . , , = 0.8.

e f f e c t on t h e p r e s s u r e d i s t r i b u t i o n s .

E f f e c t of L e f t Nozzle Boattail Angle A t M , IJ 1.2 ( f i g . 21 ( d ) ) both sets o f upper f u s e l a g e and upper n a c e l l e data F i g u r e 23 p r e s e n t s t h e e f f e c t s of d i s p l a y steeper g r a d i e n t s i n Cp r e l a t i v e a n i n c r e a s e i n t h e l e f t nozzle b o a t t a i l t o M , LJ 0.9. The g r a d i e n t s induced by a n g l e on a f t e r b o d y p r e s s u r e c o e f f i c i e n t f l o w expansion and compression are d i s t r i b u t i o n s f o r an angle-of-attack i n c r e a s e d ; t h o s e caused by compression range of l o t o 2 O and an NPFU range of are probably i n t e n s i f i e d by s t a n d i n g 2.5 t o 5 . 0 . I n f i g u r e s 2 3 ( a ) and local shock waves. On t h e upper tail- 2 3 ( b ) , d a t a are shown f o r M , IJ 0.6, boom t h e s t a n d i n g shock wave e v i d e n t a t i l l u s t r a t i n g BTL v a l u e s of 15.1° and M , LJ 0.9 has disappeared.

1 8.4O, r e s p e c t i v e l y . Changes i n Cp l e v e l are minimal a t l o c a t i o n s i m m e - On t h e lower s u r f a c e s , both sets of d i a t e l y upstream of t h e nozzle. In d a t a a l s o demonstrate steep g r a d i e n t s f i g u r e s 2 3 ( c ) and 2 3 ( d ) , BTL values of similar t o those on t h e upper s u r f a c e s .

15.1 O and 18.4O, r e s p e c t i v e l y , are On t h e lower f u s e l a g e , t h e flow acceler- p r e s e n t e d f o r M, LJ 0.9. The e f f e c t s ates downstream of X/L = 0.684, causing o f t h e b o a t t a i l a n g l e change are mini- l a r g e r e d u c t i o n s i n Cp. The expansion m a l a t l o c a t i o n s immediately upstream f o r t h e f l i g h t d a t a exceeds t h a t f o r of t h e nozzle.

t h e model data.

Summary E f f e c t of Angle of Attack The e f f e c t s of t h e test v a r i a b l e s on The e f f e c t of a n g l e of a t t a c k on t h e comparison of p r e s s u r e c o e f f i c i e n t t h e comparison of a f t e r b o d y p r e s s u r e d i s t r i b u t i o n on t h e a f t e r b o d y s u r f a c e s c o e f f i c i e n t d i s t r i b u t i o n s f o r M , LJ 0.8 is are summarized as follows. The Cp pat- shown i n f i g u r e 22. D a t a are p r e s e n t e d t e r n s f o r model and f l i g h t d a t a are s i m i - f o r a n g l e s of a t t a c k of l o , 3O, S o , and l a r over t h e Mach number test range. On 7O ( f i g s . 2 2 ( a ) to 2 2 ( d ) , r e s p e c t i v e l y ) , t h e upper s u r f a c e s and t h e lower n a c e l l e a t BTL IJ 18.4O, and w i t h NPFU ranging a t subsonic speeds, t h e f l i g h t values of from 2.5 t o 3.7. The i n c r e a s e i n a n g l e Cp are g e n e r a l l y less n e g a t i v e than t h o s e o f a t t a c k from 1 to 7O produced t h e o f t h e model. The e f f e c t s of t h e v a r i a - f o l l o w i n g r e s u l t s . On t h e upper f u s e l a g e t i o n s i n a n g l e of a t t a c k and b o a t t a i l and n a c e l l e t h e shapes of t h e p r e s s u r e a n g l e on t h e l e v e l and t r e n d of a f t e r b o d y c o e f f i c i e n t d i s t r i b u t i o n s for t h e model p r e s s u r e coefficient d i s t r i b u t i o n s are and f l i g h t data do n o t change s i g n i f i - g e n e r a l l y small.

c a n t l y . Upstream of X/L = 0.800, t h e l e v e l of Cp d e c r e a s e s w i t h i n c r e a s i n g Comparison of Boundary Layer P r o f i l e s a n g l e of a t t a c k ; downstream of X/L = and Thicknesses 0.800, a minimal change is noted. On t h e upper tailboom, t h e e f f e c t of t h e The e f f e c t s of Mach number and a n g l e angle-of-attack change is small.

o f attack on t h e comparison of p i t o t p r e s s u r e ratios, boundary l a y e r profiles, For t h e lower n a c e l l e , t h e l e v e l and boundary l a y e r t h i c k n e s s e s were d e t e r -

o f cp i n c r e a s e s s l i g h t l y upstream of

mined and are shown i n f i g u r e s 24 and 25.

X / L = 0.700 f o r both sets of d a t a . For D a t a are p r e s e n t e d f o r s e v e r a l b o a t t a i l t h e f l i g h t d a t a on t h e lower f u s e l a g e , a n g l e s i n m i l i t a r y and partial a f t e r - i n c r e a s i n g t h e a n g l e of a t t a c k s l i g h t l y burning p o w e r and an NPRL range of 2.0 t o i n c r e a s e s t h e l e v e l of p r e s s u r e c o e f f i - 5.0. For t h e model d a t a , t h e measured c i e n t upstream of X/L = 0.800. Simulta- h e i g h t s of t h e p i t o t tubes w e r e multi- neously, t h e a f t m o s t Cp l e v e l is reduced p l i e d by 12 t o compare them with t h e from p o s i t i v e v a l u e s t o n e g a t i v e values.

1 1 shows t h a t i n t e r f e r e n c e e f f e c t s from t h e f l i g h t data. A l s o shown i n t h e s e f i g u r e s i s t h e r e f e r e n c e t u r b u l e n t boundary l a y e r f u s e l a g e f l o w f i e l d are n e g l i g i b l e a t p r o f i l e ; i t s h e i g h t i s t h e boundary l a y e r M, 0.6 and i n c r e a s e s l i g h t l y as M , t h i c k n e s s measured i n f l i g h t .

i n c r e a s e s from 0.8 t o 1.2. The f l i g h t p i t o t p r e s s u r e p r o f i l e s have a t u r b u l e n t Effect of Mach Number boundary l a y e r shape and d i s p l a y a n i n c r e a s i n g loss r e l a t i v e t o t h e r e f e r e n c e The e f f e c t of Mach number on boundary t u r b u l e n t p r o f i l e s w i t h an i n c r e a s e i n l a y e r p r o f i l e s and t h i c k n e s s e s f o r an Mach number. The t h i c k n e s s e s f o r t h e a n g l e of a t t a c k of approximately 3 O i s model d a t a do n o t change as Mach number shown i n f i g u r e 24. A t t h e forward rake is i n c r e a s e d . The f l i g h t d a t a d i s p l a y l o c a t i o n ( f i g . 2 4 ( a ) ) and a t M , fl 0.6, a small d e c r e a s e i n t h i c k n e s s as Mach number is i n c r e a s e d . The model d a t a t h e model, f l i g h t , and r e f e r e n c e turbu- t h i c k n e s s exceeds t h a t f o r f l i g h t a t l e n t p r o f i l e s a g r e e c l o s e l y . A s Mach number i n c r e a s e s , t h e r e is i n c r e a s i n g a l l Mach numbers.

disagreement between t h e t h r e e c u r v e s .

A t M , - 0.8 and 0.9, t h e model d a t a d i s -

E f f e c t of Angle of Attack p l a y h i g h e r p i t o t p r e s s u r e ratios t h a n e i t h e r t h e f l i g h t d a t a or t h e r e f e r e n c e F i g u r e 2 5 p r e s e n t s t h e e f f e c t o f profile, e x c e p t n e a r t h e f u s e l a g e sur- a n g l e of a t t a c k on boundary l a y e r pro- face. A t M , fl 1.2 t h e model p r o f i l e

f i l e s and t h i c k n e s s f o r M , - 0.8. I n

undergoes a dramatic loss i n p i t o t pres- f i g u r e 25(a), a t the forward rake loca- s u r e r a t i o r e l a t i v e t o the f l i g h t and t i o n , p r o f i l e s f o r model and f l i g h t show r e f e r e n c e profiles. The model boundary r e a s o n a b l e agreement w i t h t h e r e f e r e n c e l a y e r t h i c k n e s s is i n s e n s i t i v e t o Mach t u r b u l e n t p r o f i l e as a n g l e of a t t a c k number a t s u b s o n i c s p e e d s and i n c r e a s e s i n c r e a s e s from l o to Y o . A t a 7 O , a t M , - 1.2.

i n t e r f e r e n c e from t h e f u s e l a g e f l o w f i e l d reduces Pi/Pt, above a h e i g h t o f A s shown i n t h e f l i g h t d a t a i n 24 c m f o r t h e f l i g h t data (ref. 6).

f i g u r e 2 4 ( a ) , a n i n c r e a s e i n M , c a u s e s Boundary l a y e r t h i c k n e s s f o r t h e model i s a n i n c r e a s e d loss i n t h e p i t o t p r e s s u r e c o n s t a n t w i t h i n c r e a s i n g a n g l e of a t t a c k , r a t i o i n t h e boundary l a y e r P,/Pt, b u t t h e t h i c k n e s s f o r f l i g h t decreases

( * 1

w i t h i n c r e a s i n g a n g l e of a t t a c k above 3O.

relative t o a r e f e r e n c e t u r b u l e n t pro- A t a n a n g l e of a t t a c k of 7 O , t h e t h i c k - f i l e , p a r t i c u l a r l y n e a r t h e s u r f a c e .

n e s s f o r t h e f l i g h t i s less t h a n t h e T h i s r e s u l t s u g g e s t s flow losses n e a r model t h i c k n e s s .

t h e s u r f a c e i n e x c e s s of t h o s e a t t r i b u t - able t o a t u r b u l e n t boundary l a y e r . The

A s shown i n f i g u r e 25 ( b ) , b o t h sets

f l i g h t boundary t h i c k n e s s e s are rela- t i v e l y i n s e n s i t i v e t o Mach number. of data i n d i c a t e t h a t angle-of-attack e f f e c t s on t h e a f t r a k e exceed t h o s e on A t t h e a f t r a k e l o c a t i o n ( f i g . 2 4 ( b ) ) , t h e forward rake. A s a n g l e of a t t a c k i n c r e a s e s , t h e m o d e l profiles i n c r e a s - t h e m o d e l levels of Pi/Pt, e x h i b i t a i n g l y f a l l below t h e r e f e r e n c e t u r b u l e n t greater s e n s i t i v i t y t o Mach number t h a n and f l i g h t profiles. T h i s i n d i c a t e s f l o w w a s t h e case a t t h e forward rake. The losses n e a r t h e m o d e l s u r f a c e exceeding m o d e l p r o f i l e s have t h e lowest levels of t h o s e for a t u r b u l e n t boundary l a y e r .

Pi/Pt, relative t o t h e f l i g h t and turbu- With i n c r e a s i n g a n g l e of a t t a c k , bound- a r y l a y e r t h i c k n e s s i n c r e a s e s markedly l e n t p r o f i l e s f o r a l l test Mach numbers f o r t h e model. The f l i g h t p r o f i l e s have and show a n i n c r e a s i n g d e p a r t u r e from a a t u r b u l e n t shape and a g r e e w i t h t h e t u r b u l e n t shape as M , increases from 0.6 r e f e r e n c e p r o f i l e s throughout t h e angle- t o 1.2. The t r e n d of P i / P t above 24 c m Q) o f - a t t a c k range. I n t e r f e r e n c e e f f e c t s streams as t h e y flow downstream over t h e from t h e f u s e l a q e are noted a t a 3 O model s u r f a c e s .

and 7 O , and t h i c k n e s s decreases as a n g l e of attack i n c r e a s e s from 3 O t o 7 O . A s i n d i c a t e d on t h e f i g u r e s , t h e Reference 6 p r e s e n t s a d d i t i o n a l d a t a on dye ports are l o c a t e d on t h e upper cowl, t h e j u n c t u r e of the wing and t h e glove, boundary l a y e r t h i c k n e s s measured i n f l i q h t . t h e forward p o r t i o n of t h e i n l e t f a i r - i n g , and t h e forebody. A dye port is Summary l o c a t e d a t each of t h e boundary l a y e r r a k e l o c a t i o n s . The dye streams i s s u i n g Mach number and a n g l e of a t t a c k from t h e boundary l a y e r r a k e l o c a t i o n s are u s e f u l i n i n t e r p r e t i n g flow d i r e c - a f f e c t e d t h e model boundary l a y e r pro- f i l e s and t h i c k n e s s e s more than f o r t i o n a t t h e forward and a f t boundary l a y e r r a k e l o c a t i o n s .

t h e f l i g h t p r o f i l e s and t h i c k n e s s e s a t b o t h r a k e s t a t i o n s . For t h e model, an For t h e f l o w i n g - i n l e t c o n f i g u r a t i o n , i n c r e a s e i n Mach number reduced P i / P t OD which is similar t o t h a t of t h e a i r p l a n e near t h e f u s e l a g e s u r f a c e f o r both rakes.

( f i g s . 2 6 ( a ) t o 2 6 ( c ) 1, i n c r e a s i n g t h e An i n c r e a s e i n a n g l e of attack had a a n g l e of a t t a c k causes an i n c r e a s e i n similar e f f e c t on t h e a f t rake. The spanwise wing flow outboard near the wing boundary l a y e r t h i c k n e s s a t t h e a f t r a k e t r a i l i n g edge. T h i s i s e v i d e n t immedi- s t a t i o n exceeded t h a t f o r t h e forward a t e l y ahead of t h e flow channel c r e a t e d r a k e s t a t i o n and i n c r e a s e d markedly w i t h by t h e twin v e r t i c a l tails. An examina- a n g l e of a t t a c k .

t i o n of t h e s i d e views i n d i c a t e s t h a t t h e spanwise wing flow l i f t s upward from t h e For t h e f l i g h t d a t a , an i n c r e a s e i n wing s u r f a c e w i t h i n c r e a s i n g a n g l e of Mach number reduced t h e l e v e l of Pi/Pt, attack ( f i g s . 26(d) t o 2 6 ( f ) ) .

for both r a k e s , b u t t h e e f f e c t w a s less The dye stream i s s u i n g from t h e t h a n t h a t f o r t h e model d a t a . A n in- crease i n a n g l e of a t t a c k from 3 O t o 7 O l o c a t i o n of t h e forward rake i n i t i a l l y shows an outboard flow d i r e c t i o n reduced boundary l a y e r t h i c k n e s s f o r both r a k e s , a n e f f e c t o p p o s i t e of t h a t f o r t h e ( f i g s . 2 6 ( a ) to 2 6 ( c ) ) . F a r t h e r down- model. Boundary l a y e r t h i c k n e s s a t t h e stream between t h e two v e r t i c a l t a i l s a f t r a k e s t a t i o n w a s less than t h a t f o r t h e flow is d e f l e c t e d i n t o a streamwise t h e forward rake s t a t i o n and, r e l a t i v e t o d i r e c t i o n . F i n a l l y , near t h e nozzle, scale, w a s c o n s i d e r a b l y less than t h e t h e flow e x h i b i t s a s h a r p change i n flow d i r e c t i o n outboard and downward. T h i s boundary l a y e r t h i c k n e s s a t t h e a f t rake s t a t i o n on t h e model. t r e n d is d u p l i c a t e d a t t h e l o c a t i o n of t h e a f t rake. Another f a c t o r t h a t may Flow V i sua l i z a t i o n i n f l u e n c e t h e flow p a t t e r n between t h e v e r t i c a l tails is flow t h a t propagates An i n s i g h t i n t o t h e flow behavior downstream from t h e boundary l a y e r on t h e f u s e l a g e upper s u r f a c e can be d i v e r t e r , t h e gap l o c a t e d between t h e o b t a i n e d by examining the flow v i s u a l i - i n l e t and t h e forebody.

z a t i o n photographs o b t a i n e d from t h e I t i s b e l i e v e d t h a t t h e boundary w a t e r t u n n e l tests. F i g u r e s 26 and 27 show p l a n and s i d e views of t h e flowing- l a y e r a t t h e forward rake l o c a t i o n i s reduced i n t h i c k n e s s by t h e i n c r e a s i n g i n l e t and f a i r e d - i n l e t c o n f i g u r a t i o n s , r e s p e c t i v e l y , photographed i n t h e water spanwise wing flow w i t h i n c r e a s i n g a n g l e of a t t a c k . A t t h e a f t r a k e l o c a t i o n t h e t u n n e l a t a n g l e s of a t t a c k of l o , 3 O , a d d i t i o n a l r e d u c t i o n i n boundary l a y e r and 7 O . The flow p a t t e r n s are i l l u s - t h i c k n e s s may be caused by t h e l a r g e t r a t e d by t h e p a t h s t r a c e d by t h e dye i n c r e a s e i n outboard and downward flow as v o r t e x shed from t h e l e a d i n g edge of a n g l e of a t t a c k is increased. t h e i n l e t f a i r i n g may also c o n t r i b u t e t o t h e flow p a t t e r n between the t w o For t h e f a i r e d - i n l e t c o n f i g u r a t i o n v e r t i c a l tails.

( f i g s . 2 7 ( a ) to 2 7 ( c ) ) , which i s similar t o t h e p r o p u l s i o n model, t h e e f f e c t of Comparison of Afterbody and Nozzle a n increase i n a n g l e of a t t a c k on t h e P r e s s u r e C o e f f i c i e n t D i s t r i b u t i o n s spanwise wing flow d i f f e r s from t h e f l o w i n g - i n l e t c o n f i g u r a t i o n . Although A s shown i n f i g u r e 20, flow i n t e r - t h e flow p a t t e r n s are similar a t a l o a c t i o n s between n a c e l l e and nozzle can i n f l u e n c e t h e p r e s s u r e c o e f f i c i e n t ( f i g s . 2 6 ( a ) and 2 7 ( a ) ) , t h e y begin to d i f f e r a t a = 3 O ( f i g s . 26(b) and 2 7 ( b ) ) . d i s t r i b u t i o n s obtained from model and For t h e f a i r e d - i n l e t c o n f i g u r a t i o n , t h e f l i g h t tests. Figure 2 8 i l l u s t r a t e s flow i n t h e outboard d i r e c t i o n begins s e v e r a l p o s s i b l e v a r i a t i o n s i n p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n . Curve 1 shows t o break down as noted by t h e inboard t h a t t h e f l o w expanding over t h e a f t e r - s t r e a m l i n e s on each wing. A t a 7 O body reaches a minimum l e v e l upstream ( f i g . 2 7 ( c ) ) , t h e spanwise flow is of t h e a f t e r b o d y and nozzle juncture.

r e v e r s e d and flows inboard toward t h e I channel c r e a t e d by t h e v e r t i c a l tails. The subsequent p r e s s u r e rise downstream of t h e j u n c t u r e t o a p o s i t i v e value of Examining t h e s i d e views ( f i g s . 27(d) t o 27( f) 1 , t h e flow on top of t h e fuse- Cp i l l u s t r a t e s good recompression and l a g e a t a * 7 O ( f i g . 2 7 ( f ) ) appears t o a t t a c h e d flow.

be s e p a r a t i n g f r o m the s u r f a c e , which could r e s u l t i n a boundary l a y e r having Curve 2 is t h e same as curve 1 l o w energy.

e x c e p t f o r t h e downstream nozzle flow.

The d e c r e a s e i n slope and l a c k of recom- Corresponding changes i n flow d i r e c - p r e s s i o n s u g g e s t s flow s e p a r a t i o n . N o t e t i o n are noted a t t h e dye streams o r i g i - t h a t t h e loss i n p r e s s u r e i s due e n t i r e l y n a t i n g a t t h e l o c a t i o n s of t h e boundary t o flow over t h e nozzle and n o t t o any l a y e r rakes. A t a = l o ( f i g . 2 7 ( a ) ) , t h e change i n t h e upstream flow.

dye streams are flowing d i r e c t l y down- stream. A t a = 3 O ( f i g . 2 7 ( b ) ) , t h e flow Curve 3 i l l u s t r a t e s a lower minimum changes t o an inboard d i r e c t i o n , partic- p r e s s u r e than curves 1 and 2, followed by u l a r l y a t t h e a f t rake. F i n a l l y , a t a s h a r p p r e s s u r e jump. The s h a r p jump a = 7 O ( f i g . 2 7 ( c ) ) , t h e flow shows a i n d i c a t e s a s t a n d i n g compression shock s t r o n g inboard flow d i r e c t i o n f o r both wave. Following t h e shock wave, t h e flow dye streams.

s e p a r a t e s i n a manner similar to curve 2.

Note t h a t downstream of t h e s e p a r a t i o n For t h e f a i r e d - i n l e t c o n f i g u r a t i o n , p o i n t t h e l e v e l of Cp f o r curve 3 may l i e it is b e l i e v e d t h a t t h e t h i c k e n i n g of below t h a t f o r curve 2 because of t h e t h e boundary l a y e r a t t h e a f t r a k e with lower v a l u e of Cp immediately ahead of i n c r e a s i n g a n g l e of a t t a c k is influenced t h e s e p a r a t i o n p o i n t . Consequently, t h e by t h e wing flow. A s a n g l e of a t t a c k l a r g e loss i n Cp from curve 1 t o curve 3 i n c r e a s e s , t h e wing flow near t h e t r a i l - i s due t o an upstream i n f l u e n c e as w e l l i n g edge r e v e r s e s from a g e n e r a l l y out- as flow l o s s e s on t h e nozzle.

board d i r e c t i o n t o an inboard d i r e c t i o n .

The flow is channeled by t h e v e r t i c a l t a i l s toward t h e c e n t e r of t h e a i r p l a n e E f f e c t of Mach Number and t h e nozzles. The g e n e r a l flow d i r e c - t i o n between t h e vertical tails is i n - The e f f e c t of Mach number on t h e com- board, as opposed t o outboard f o r t h e p a r i s o n of a f t e r b o d y and nozzle p r e s s u r e f l o w i n g - i n l e t c o n f i g u r a t i o n . T h i s appar- c o e f f i c i e n t d i s t r i b u t i o n s a t a ~ l o i s e n t l y t h i c k e n s t h e boundary l a y e r . A shown i n f i g u r e 29. D a t a are presented X/L = 0.897 on nozzle row 1 8 2 O , t h e f o r a n NPRL test range of 2.0 t o 5.0 f l i g h t value of Cp exceeds t h a t f o r a t BTb 18.4O f o r subsonic speeds and BTb LI( 7.7' a t M, 1.2.

t h e model, although recompression is s i m i l a r f o r both sets of d a t a . On t h e upper nacelle t h e recompression f o r t h e F i g u r e 2 9 ( a ) p r e s e n t s d a t a f o r model is s l i g h t l y more e f f i c i e n t , y i e l d - M , FS 0.6, BTL FS 1 8 . 4 O , and NPRL 2.0 i n g a s l i g h t l y g r e a t e r value of Cp a t t o 3.0. Data are p r e s e n t e d f o r t h e upper X / b = 0.937 than f o r f l i g h t .

nacelle and nozzle row 350° and f o r t h e lower n a c e l l e and nozzle row 182O. Note For a Mach number of 0.9 ( f i g . 2 9 ( c ) ) , t h e break i n t h e X / b a x i s and t h e scale both sets of d a t a i n d i c a t e t h a t t h e d i f - i n c r e a s e f o r t h e nozzle p r e s s u r e d i s - f e r e n c e i n l e v e l s of Cp f o r both nozzle t r i b u t i o n . This a l l o w s a more d e t a i l e d rows exceeds t h a t a t M, LI( 0.8. On t h e examination of t h e v a r i a t i o n i n Cp on t h e upper n a c e l l e t h e f l i g h t d a t a e x h i b i t a nozzle. The d a t a a t X/L = 0.897 ( t h e s t a n d i n g shock wave immediately ahead f i r s t nozzle p r e s s u r e o r i f i c e ) appear on of t h e nozzle. Downstream of t h e shock both s i d e s of t h e p l o t s f o r t h e f l i g h t wave t h e r e i s a l a r g e loss i n recompres- d a t a and f o r nozzle row 1 8 2 O f o r t h e s i o n compared t o t h a t a t pJ 0.8; t h i s model d a t a .

l o s s is a t t r i b u t e d t o flow s e p a r a t i o n .

The pressure d i s t r i b u t i o n s shown f o r The amount of recompression does n o t t h e upper and lower s u r f a c e s i l l u s t r a t e d i f f e r g r e a t l y between model and f l i g h t ; t h e example of curve 1 i n f i g u r e 2 8 , t h a t i s , t h e d i s t r i b u t i o n s are roughly namely, good recompression t o p o s i t i v e p a r a l l e l . A s a r e s u l t , t h e d i f f e r e n c e l e v e l s of Cp on t h e downstream p o r t i o n of i n Cp e x i s t i n g a t upstream nozzle loca- t h e nozzle. As p r e v i o u s l y d i s c u s s e d , on t i o n s remains about t h e same everywhere t h e n a c e l l e s u r f a c e s t h e Cp l e v e l f o r t h e a long t h e nozzle.

f l i g h t d a t a is g r e a t e r than t h a t f o r t h e S i m i l a r r e s u l t s are e v i d e n t on t h e model, t h a t is, t h e v a l u e s of Cp are less lower n a c e l l e and nozzle. The recompres- n e g a t i v e . The d i f f e r e n c e i n t h e f u s e l a g e s i o n a t M , LI( 0.9 is about t h e same f o r v e n t c o n f i g u r a t i o n s between model and model and f l i g h t . The f l i g h t d a t a are f l i g h t may be c o n t r i b u t i n g t o t h e lower less n e g a t i v e t h a n t h e model d a t a a t p r e s s u r e c o e f f i c i e n t s immediately up- X/L = 0.897 and remain t h i s way as X/L stream of t h e nozzle f o r t h e model. A s i n c r e a s e s ; t h a t is, t h e d i s t r i b u t i o n s d i s c u s s e d i n r e f e r e n c e 6 , when t h e vent gap w a s sealed t o p r e v e n t outflow, the are roughly p a r a l l e l . Therefore, it would appear t h a t t h e d i f f e r e n c e s i n t h e pressure c o e f f i c i e n t s on t h e f i r s t and p r e s s u r e d i s t r i b u t i o n s are caused by second nozzle o r i f i c e s w e r e reduced.

upstream i n f l u e n c e s .

This nonflowing v e n t c o n f i g u r a t i o n w a s simulated on t h e model.

A t M , 1.2 ( f i g . 2 9 ( d ) ) , recompres- D e s p i t e t h e s l i g h t l y l a r g e r l e v e l s of s i o n d e t e r i o r a t e s on both nozzle s u r f a c e s Cp a t X/L g r e a t e r than 0.850 shown by t h e f o r t h e model and f l i g h t d a t a . The f l i g h t d a t a i n d i c a t e t h a t a s t a n d i n g shock wave f l i g h t d a t a on t h e nacelle s u r f a c e s , t h e i s l o c a t e d downstream of t h e f i r s t nozzle recompression p r o c e s s e s on both nozzle rows are very similar f o r both sets of o r i f i c e , t h a t i s , downstream of t h e junc- t u r e between t h e f i x e d and movable noz- d a t a . The f l i g h t d a t a e x h i b i t s l i g h t l y z l e s . This is i n d i c a t e d by t h e s h a r p rise h i g h e r values.

i n Cp. The model d a t a a l s o show t h i s A s Mach number increases t o 0.8 r e s u l t f o r nozzle row 1 8 2 O . The flow ( f i g . 2 9 ( b ) ) t h e t r e n d s are similar t o separates downstream of t h i s o r i f i c e f o r t h o s e shown i n f i g u r e 2 9 ( a ) . A t model and f l i g h t because of t h e shock wave. The l e v e l of Cp for f l i g h t is less on t h e upstream nozzle o r i f i c e s w e r e ~ n e g a t i v e than f o r t h e model downstream g e n e r a l l y less negative than t h o s e f o r t h e model because of t h e upstream n a c e l l e of t h e j u n c t u r e on both s u r f a c e s . The average l e v e l of Cp on t h e lower nozzle For both and data a t M , fl 0.6 and 0.8, good recompression s u r f a c e is less n e g a t i v e than t h e average e x i s t e d f o r nozzle rows 182' and 350O.

l e v e l of Cp on t h e upper nozzle s u r f a c e A t M, IJ 0.9 and 1 . 2 , recompression f o r both sets of d a t a .

d e t e r i o r a t e d because of flow s e p a r a t i o n and flow complexity. The upstream d i f - E f f e c t of Angle of Attack f e r e n c e s i n t h e Cp l e v e l s between model The e f f e c t of a n g l e of a t t a c k on t h e and f l i g h t were r e t a i n e d along both nozzle s u r f a c e s . These r e s u l t s demon- comparison of a f t e r b o d y nozzle p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s a t M , IJ 0.8 s t r a t e upstream i n f l u e n c e s on t h e nozzle p r e s s u r e s . A t M, fl 0.6 and 0.9, i n c r e a s - is presented i n f i g u r e 30 f o r a n g l e s of i n g b o a t t a i l a n g l e reduced t h e recompres- a t t a c k of l o , 3', 5O, and 6 O to 7O.

The o v e r a l l e f f e c t s of a n g l e of a t t a c k s i o n a t both nozzle rows, b u t t h e e f f e c t w a s l a r g e r a t nozzle row 350O. The on recompression on t h e upper and lower e f f e c t s of a n g l e of a t t a c k were small.

s u r f a c e s are small f o r both sets of d a t a .

Comparison of P r e s s u r e C o e f f i c i e n t D i s t r i b u t i o n s on t h e Nozzle E f f e c t of L e f t Nozzle B o a t t a i l Angle E x t e r n a l S u r f a c e The effect of l e f t nozzle b o a t t a i l The e f f e c t s of t h e test v a r i a b l e s on a n g l e on t h e comparison of a f t e r b o d y a l l e i g h t rows of t h e s u r f a c e p r e s s u r e n o z z l e p r e s s u r e c o e f f i c i e n t d i s t r i b u - o r i f i c e s of t h e nozzles are d i s c u s s e d i n t i o n s i s i l l u s t r a t e d i n f i g u r e 31 f o r an t h e following s e c t i o n s .

angle-of-attack range of l o t o 2 O and an NPRL range of 2.5 t o 5.0. F i g u r e s 31 ( a ) E f f e c t of Mach Number and 3 1 ( b ) p r e s e n t d a t a a t M , fl 0.6 f o r nominal b o a t t a i l a n g l e s of 15.1O and Figure 32 shows t h e e f f e c t of Mach 18.4O, r e s p e c t i v e l y . On nozzle r o w number on nozzle p r e s s u r e c o e f f i c i e n t 350°, f o r both sets of d a t a , a loss i n d i s t r i b u t i o n s f o r t h e e i g h t nozzle r o w s .

recovery ( t h a t is, a d e c r e a s e i n s l o p e ) D a t a are p r e s e n t e d a t c i fl l o f o r an NPRL is noted f o r X/L g r e a t e r than 0.920 as range of 2.0 t o 5.0, f o r E W L fl 18.4O and b o a t t a i l a n g l e is i n c r e a s e d .

BTL IJ 7.7O. For M , fl 0.6 ( f i g . 3 2 ( a ) ) , t h e p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s ( c ) and 31 ( d ) show c o r r e - F i g u r e s 31 f o r both sets of d a t a a g r e e reasonably sponding data a t M , fl 0.9. A decrease w e l l . Highest recompression is a t t h e i n slope is shown f o r both nozzle r o w s bottom rows ( f o r example, r o w s 230°, because of t h e b o a t t a i l a n g l e i n c r e a s e .

182O, and 134O 1 and t h e top rows (302O The loss i n recompression i s l a r g e r f o r and 350O). For t h e model, r o w 86O nozzle r o w 350° than it is f o r nozzle a p p e a r s t o be s e p a r a t e d . The v a l u e of r o w 182O.

Cp i s p o s i t i v e f o r a l l e i g h t r o w s a t t h e l a s t o r i f i c e . Where d a t a are a v a i l a b l e Summary a t X/L = 0.897, Cp f o r f l i g h t exceeds Cp The e f f e c t s of t h e test v a r i a b l e s f o r t h e model except for t h e s e p a r a t e d r o w . T h i s i s a t t r i b u t e d t o upstream on t h e comparisons of a f t e r b o d y nozzle p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s are e f f e c t s .

F l i g h t Cp l e v e l s summarized as follows.

E f f e c t of L e f t Nozzle Boattail Angle A t M , (J 0.8 ( f i g . 3 2 ( b ) 1 , model and f l i g h t d a t a e x h i b i t a s l i g h t loss of recompression compared with M , (J 0.6 on F i g u r e 34 d i s p l a y s t h e e f f e c t of l e f t nozzle b o a t t a i l a n g l e on t h e com- a l l rows; agreement i s good. A t a l l rows p a r i s o n of nozzle pressure c o e f f i c i e n t f o r both sets of d a t a , t h e value of Cp d i s t r i b u t i o n f o r an angle-of-attack range a t t h e l a s t o r i f i c e is smaller than a t of 1' t o 2 O and an NPRL range of 2.5 t o M , (J 0 . 6 . A t row 86O, both sets of d a t a 5.0. F i g u r e s 3 4 ( a ) and 3 4 ( b ) p r e s e n t d a t a a t Moo = 0.6 f o r nominal b o a t t a i l i n d i c a t e s e p a r a t i o n .

a n g l e s of 15.1 and 18.4O, r e s p e c t i v e l y .

The p r o f i l e s i n d i c a t e t h a t o v e r a l l recom- A t M , 0.9 ( f i g . 3 2 ( c ) ) , t h e pres- p r e s s i o n is reduced as t h e b o a t t a i l s u r e c o e f f i c i e n t d i s t r i b u t i o n s show a a n g l e is i n c r e a s e d a t a l l nozzle r o w s s i g n i f i c a n t loss of recompression f o r e x c e p t 182O and 230'. For t h e f l i g h t both sets of d a t a ( e x c e p t f o r r o w 182O 1, d a t a i n f i g u r e s 3 4 ( c ) and 3 4 ( d ) , f o r The f l a t p r o f i l e s f o r s e v e r a l of t h e M , 0.9, t h e r e s u l t s a l s o d i s p l a y a rows s u g g e s t an i n c r e a s i n g amount of s e p a r a t e d flow on t h e nozzle s u r f a c e . loss i n recompression as t h e b o a t t a i l The model and f l i g h t Cp p r o f i l e s appear a n g l e is i n c r e a s e d as w e l l as an i n c r e a s i n g tendency toward s e p a r a t i o n .

t o be almost parallel r a t h e r than coin- c i d e n t , which w a s t h e case a t M , = 0.8.

Summary R o w 182O shows a p o s i t i v e value of Cp a t X f i = 0.931; t h e v a l u e s of Cp f o r a l l The e f f e c t s of t h e test v a r i a b l e s o t h e r r o w s a t X/L = 0.931 are e i t h e r on t h e comparisons of pressure coef- z e r o or negative. f i c i e n t d i s t r i b u t i o n s on t h e nozzle e x t e r n a l s u r f a c e are summarized as A t M , (J 1.2 ( f i g . 3 2 ( d ) ) , t h e pres- follows. A t M , (J 0.6 and 0.8, t h e r e s u r e c o e f f i c i e n t d i s t r i b u t i o n s f o r both w a s reasonably good agreement between sets of data a t a l l r o w s e x h i b i t l a r g e r t h e model and f l i g h t i n l e v e l and trend.

n e g a t i v e Cp l e v e l s , as w e l l as i n c r e a s e d A t M , (J 0.9 and 1.2, both sets of d a t a flow complexity r e l a t i v e to t h e subsonic showed i n c r e a s i n g evidence of flow sep- d a t a ( f i g s . 3 2 ( a ) , ( b ) , and ( c ) ) . A t a l l a r a t i o n , flow complexity, and a loss i n rows, t h e p r o f i l e s f o r t h e f l i g h t d a t a recompression. The t r e n d s e x h i b i t e d by t h e model and f l i g h t d a t a were t h e same, are less n e g a t i v e than those f o r t h e model data. Row 1 8 2 O r e v e a l s a s t a n d i n g b u t t h e f l i g h t data e x h i b i t e d less neg- a t i v e l e v e l s of Cp than t h e model on a l l shock wave and shock-induced s e p a r a t i o n for both model and f l i g h t .

t h e nozzle r o w s . T h i s i s a t t r i b u t e d t o upstream e f f e c t s . I n c r e a s i n g l e f t nozzle E f f e c t of Angle of A t t a c k b o a t t a i l a n g l e g e n e r a l l y caused losses i n recompression f o r a l l rows, except The e f f e c t of a n g l e of a t t a c k on t h e r o w s 182O and 230°. The e f f e c t s of a n g l e of a t t a c k on t h e d a t a were minimal.

comparison of nozzle p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s is shown i n f i g u r e 33 f o r Comparison of Recompression on t h e a n g l e s of a t t a c k of l o , 3O, 5O, and 6 O Nozzle E x t e r n a l S u r f a c e t o 7'. The m o d e l and f l i g h t d a t a show r e a s o n a b l e agreement i n l e v e l and t r e n d The comparison of recompression on f o r a l l a n g l e s of a t t a c k . The angle-of- t h e nozzle e x t e r n a l s u r f a c e and t h e a t t a c k v a r i a t i o n has m i n i m a l e f f e c t on b o t h sets of d a t a .

e f f e c t s of Mach number and l e f t nozzle i n t e r f e r e n c e e f f e c t s of t h e t a i l s , t a i l - b o a t t a i l angle a r e d i s c u s s e d i n t h e boom, o r nozzle i n t e r f a i r i n g than are following s e c t i o n s . any of t h e o t h e r nozzle rows. Therefore, t h e h i g h e s t p r e s s u r e c o e f f i c i e n t s are E f f e c t of Mach Number achieved. The v e r t i c a l recompression l i n e s show t h a t recompression f o r f l i g h t The e f f e c t of Mach number on nozzle exceeds t h a t f o r t h e model.

recompression i s shown i n f i g u r e 35.

The p r e s s u r e c o e f f i c i e n t s f o r t h e f i r s t A t M , 0.8 with BTL IJ 18.4' and l a s t nozzle o r i f i c e s of each row a r e ( f i g . 3 5 ( b ) ) , t h e d a t a show s l i g h t l y p l o t t e d a g a i n s t c i r c u m f e r e n t i a l p o s i t i o n l e s s recompression than a t M , 0.6 f o r on t h e nozzle f o r s e v e r a l test Mach num- model and f l i g h t , a s i n d i c a t e d by t h e b e r s ( f i g s . 3 5 ( a ) t o 3 5 ( d ) ) and a g a i n s t s h o r t e r v e r t i c a l l i n e s . The shapes of Mach number ( f i g . 3 5 ( e ) ) .

t h e Cp p l o t s a g a i n s t (9 are n o t g r e a t l y changed from those a t M, a 0.6 f o r model A v a i l a b l e d a t a are presented f o r t h e and f l i g h t .

e i g h t rows a t a ranging from l o t o 2 O , BTL f( 18.4' and 7.7', and an NPRL range The d a t a i n f i g u r e 3 5 ( c ) are f o r of 2.0 t o 5.0. Shown t o t h e r i g h t of each d i s t r i b u t i o n (as a s o l i d symbol) is M , 0.9 with BTL 18.4O and show an t h e average value f o r each of t h e cir- i n c r e a s i n g l o s s i n recompression com- c u m f e r e n t i a l d i s t r i b u t i o n s . The l e n g t h s pared w i t h M , 0.8 €or both sets of and r e l a t i v e h e i g h t s of t h e v e r t i c a l d a t a , as i n d i c a t e d by t h e s h o r t e r ver- l i n e s connecting t h e s o l i d symbols f o r t i c a l l i n e s . I n a d d i t i o n , t h e model t h e f i r s t and l a s t o r i f i c e s are i n t e r - Cp p r o f i l e f o r t h e f i r s t o r i f i c e i s p r e t e d as a measure of t h e recompression c o n s i d e r a b l y more n e g a t i v e than t h a t on t h e nozzle f o r model and f l i g h t .

f o r f l i g h t e x c e p t a t row 86', t h e row Therefore, comparing t h e v e r t i c a l l i n e s w i t h s e p a r a t e d flow. A t t h e l a s t nozzle y i e l d s a comparison of recompression on o r i f i c e t h e f l i g h t Cp p r o f i l e is con- t h e nozzle.

s i s t e n t l y less n e g a t i v e than t h e model A t M, LI1 0.6 w i t h BTL 18.4O Cp p r o f i l e .

( f i g . 3 5 ( a ) ) , t h e Cp a t t h e f i r s t nozzle A t M , 1 . 2 and BTL 7.7O o r i f i c e s f o r f l i g h t are less negative t h a n t h o s e f o r t h e model, e x c e p t a t ( f i g . 3 5 ( d ) ) , t h e l e v e l s of recompres- s i o n i n d i c a t e d by t h e v e r t i c a l l i n e s do @ = 86'. Figure 3 2 ( a ) shows t h a t t h e model flow i s s e p a r a t e d a t row 86'. I n n o t d i f f e r g r e a t l y , f o r e i t h e r s e t of d a t a , from t h e r e s p e c t i v e v a l u e s a t f i g u r e 3 5 ( a ) , it is b e l i e v e d t h a t t h e p o s i t i v e l e v e l of Cp n e a r t h e nozzle M , f( 0.9. Therefore, t h e more n e g a t i v e e x i t a t row 86' f o r t h e model propagate Cp values a t t h e l a s t nozzle o r i f i c e upstream i n t h e subsonic f i e l d through r e l a t i v e t o M, 0.9 ( f i g . 3 5 ( c ) ) are t h e s e p a r a t e d r e g i o n , r a i s i n g t h e l e v e l caused by t h e more n e g a t i v e Cp o c c u r r i n g of Cp a t t h e f i r s t nozzle o r i f i c e .

a t t h e f i r s t nozzle o r i f i c e . This is Otherwise, t h e l e v e l would be more t r u e f o r both sets of d a t a .

n e g a t i v e than t h a t f o r f l i g h t .

Figure 3 5 ( e ) summarizes t h e e f f e c t The Cp l e v e l s f o r t h e l a s t nozzle of Mach number on nozzle recompression f o r an a range of l o t o 2 O , an NPRL range o r i f i c e a g r e e reasonably w e l l € o r model and f l i g h t . Highest Cp l e v e l s a r e noted of 2.0 t o 5.0, and BTL 18.4O and 7 . 7 ' .

Data are shown f o r t h e t e s t c o n d i t i o n s f o r o r i f i c e row 182' a t t h e bottom of t h e of f i g u r e s 3 5 ( a ) t o ( d ) and t h e addi- n a c e l l e . This row i s f a r t h e s t from and NPRL ranged from 2.5 t o 5.0. A t t i o n a l test c o n d i t i o n s shown i n t a b l e 7.

The d a t a were averaged f o r t h e s e v e r a l M , 0.6, t h e average Cp l e v e l a t t h e t e s t c o n d i t i o n s of t h e e i g h t nozzle o r i - f i r s t o r i f i c e is approximately t h e same f i c e rows y i e l d i n g s i n g l e values (Cpavg f o r both b o a t t a i l a n g l e s f o r t h e model

)

and f l i g h t d a t a . A s l i g h t loss i n recom- € o r t h e p r e s s u r e c o e f f i c i e n t s a t t h e p r e s s i o n r e s u l t s from t h e i n c r e a s e i n f i r s t and l a s t o r i f i c e s a t each Mach b o a t t a i l angle and i s about t h e same f o r number. A s defined i n t h e upper p a r t t h e model and f l i g h t d a t a . A t M , @ 0.9, of f i g u r e 3 5 ( e ) , kP i s a measure of t h e loss i n recompression caused by t h e t h e average recompression on t h e nozzle.

b o a t t a i l a n g l e i n c r e a s e exceeds t h a t a t M, 0.6. This is p a r t i c u l a r l y t r u e f o r For subsonic speeds a t t h e f i r s t t h e f l i g h t d a t a .

nozzle o r i f ice , ‘pavg f o r f l i g h t exceeds

t h a t f o r t h e model and i n c r e a s e s s l i g h t - Summary l y with Mach number. A t & @ 1 . 2 t h e increment of Cpavg f o r f l i g h t over The e f f e c t s of t h e t e s t v a r i a b l e s o n t h e comparison of nozzle recompression t h a t f o r t h e model i s g r e a t l y i n c r e a s e d .

are summarized as follows. A n i n c r e a s e The increments f o r subsonic speeds and i n subsonic Mach number caused a s t e a d y M , @ 1 . 2 are a t t r i b u t e d t o upstream l o s s i n nozzle recompression f o r t h e e f f e c t s .

model and f l i g h t d a t a . A t M, @ 0.84 t o 0.85, t h e average Cp a t t h e l a s t nozzle For subsonic speeds a t t h e l a s t o r i - o r i f i c e went from p o s i t i v e t o negative.

f i c e , cpavg f o r f l i g h t s l i g h t l y exceeds A t M , @ 0.6 and 0.9, an i n c r e a s e i n boat- f o r t h e model. Both values of ‘Pavg t a i l a n g l e from 15.1O t o 1 8 . 4 O caused a e x h i b i t a s t e a d y d e c r e a s e with 4, loss i n recompression.

‘Pavg and a t & @ 0.84 t o 0.85 t h e v a l u e s Comparison of Nozzle Axial become negative. A t & @ 1.2, t h e Force C o e f f i c i e n t p o i n t s are more n e g a t i v e than ‘Pavg The e f f e c t s of t h e test v a r i a b l e s t h o s e a t subsonic speeds p a r t i c u l a r l y on nozzle a x i a l f o r c e c o e f f i c i e n t a r e f o r t h e model.

d i s c u s s e d i n t h e f o l l o w i n g s e c t i o n s .

The recompression c u r v e s , t h a t is, E f f e c t of Mach Number and L e f t t h e A c p curves, show a s t e a d y l o s s i n Nozzle Boattail Angle recompression over t h e subsonic Mach number range. Recompression f o r t h e F i g u r e 37 shows t h e e f f e c t s of Mach model exceeds t h a t f o r f l i g h t . A t number and l e f t nozzle b o a t t a i l a n g l e 4 @ 1.2, recompression f o r model and on t h e comparison of nozzle a x i a l f o r c e f l i g h t a g r e e and are comparable t o t h e c o e f f i c i e n t Ca. I n f i g u r e 3 7 ( a ) , d a t a f l i g h t v a l u e f o r M, 0.9.

are shown f o r a l o t o 2 O and f o r BTL 18.4O and 7.7O over an NPRL range E f f e c t of L e f t Nozzle Boattail Angle of 2.5 t o 5.0. The f a i r i n g f o r t h e f l i g h t d a t a g i v e s more weight t o t h e 6100-m d a t a p o i n t s because t h e d a t a a t The e f f e c t of l e f t nozzle b o a t t a i l 10,700 m are s u b j e c t t o g r e a t e r e r r o r a n g l e on t h e comparison of nozzle recom- p r e s s i o n i s shown i n f i g u r e 36. Data are ( f i g . 1 9 ) . For BTL 18.4’ and M, p r e s e n t e d f o r l e f t nozzle b o a t t a i l a n g l e s ranging from 0.6 t o 0.8, t h e f a i r e d of 15.1O and 1 8 . 4 O and & 0.6 and 0.9.

v a l u e s of t h e model and f l i g h t d a t a a g r e e and i n c r e a s e s l i g h t l y w i t h Mach Angles of a t t a c k ranged from l o t o 2 O , number. The i n c r e a s e i n C a is caused M , (J 0.6, no e f f e c t of a n g l e of attack by t h e i n c r e a s i n g l y n e g a t i v e Cp l e v e l s i s e v i d e n t ; a t M , (J 0.8, C a i n c r e a s e s w i t h Mach number i n c r e a s e shown i n s l i q h t l y a s a n g l e of a t t a c k i n c r e a s e s f i g u r e s 3 2 ( a ) and 3 2 ( b ) and t h e l o s s i n from 1' to 7'. A t M , fl 0.9, t h e d a t a recompression shown i n f i g u r e 3 5 ( e ) .

scatter and small v a r i a t i o n i n a n g l e of a t t a c k p r e c l u d e any i n t e r p r e t a t i o n .

AS M , i n c r e a s e s t o 0.9, Ca f o r b o t h sets of d a t a i n c r e a s e s , b u t Ca E f f e c t of Nozzle P r e s s u r e R a t i o f o r t h e m o d e l d a t a i n c r e a s e s more than The e f f e c t of nozzle pressure r a t i o f o r t h e f l i g h t d a t a . As i n d i c a t e d i n (NPRL) on t h e comparison of n o z z l e a x i a l f i g u r e 3 2 ( c ) , l e v e l s of Cp f o r t h e model f o r c e c o e f f i c i e n t i s shown i n f i g u r e 39.

d a t a are c o n s i d e r a b l y lower than t h o s e T e s t d a t a are f o r a (J 1' t o 2' and En% IJ f o r t h e f l i g h t d a t a . T h i s i s a t t r i b u t e d 18.4' and 7.7' f o r s e v e r a l test Mach num- t o upstream e f f e c t s . A s shown i n f i g - b e r s . The d e s i g n p r e s s u r e r a t i o shown u r e 3 5 ( e ) , both sets of d a t a e x h i b i t a f o r each n o z z l e is t h e s o l i d v e r t i c a l loss i n recompression as M , i n c r e a s e s l i n e . I n f i g u r e s 3 9 ( a ) t o 3 9 ( c ) , f o r t h e from 0.8 t o 0.9.

s u b s o n i c test p o i n t s , t h e model d a t a were o b t a i n e d over an NPRL range e x t e n d i n g I n f i g u r e 3 7 ( a ) , a t M , (J 1.2 w i t h from j e t o f f , through t h e overexpanded BTL fl 7.7O, C a v a l u e s f o r both sets of r e g i o n where NPRL i s g r e a t e r than approx- i m a t e l y 1.9 and less than the d e s i g n d a t a exceed t h e r e s p e c t i v e subsonic l e v e l s . C , f o r t h e model i s g r e a t e r N P F U , i n t o t h e underexpanded r e g i o n where NPRL i s g r e a t e r than t h e d e s i g n value.

than t h e C a l e v e l f o r t h e f l i g h t d a t a .

T h i s r e s u l t is a l s o e x p l a i n e d by f i g - With t h e j e t o p e r a t i n g , peak l e v e l s u r e s 3 2 ( d ) and 3 5 ( e ) .

of C a occur near t h e d e s i g n v a l u e of NPRL = 3.4. A s e x p l a i n e d i n r e f e r - F i g u r e 3 7 ( b ) shows t h e e f f e c t of l e f t e n c e s 1 2 , 13, and 14, j e t i n t e r f e r e n c e n o z z l e boattail a n g l e on nozzle a x i a l e f f e c t s f o r v a l u e s of NPRL less than t h a t f o r c e c o e f f i c i e n t f o r subsonic Mach num- a t peak C a are i n f l u e n c e d p r i m a r i l y by b e r s . The l i n e s shown are the f a i r i n g s e n t r a i n m e n t of t h e n o z z l e e x t e r n a l flow o f f i g u r e 3 7 ( a ) f o r BTL fl 18.4'; t h e d a t a caused by v i s c o u s s h e a r and mixing.

p o i n t s are f o r t h e lower boattail angles.

Entrainment t e n d s t o reduce t h e p r e s s u r e A t M , LJ 0.6, C a f o r t h e l o w e r b o a t t a i l on t h e n o z z l e e x t e r n a l s u r f a c e . A s NPRL a n g l e s i s o n l y s l i g h t l y less than C a f o r i n c r e a s e s , j e t i n t e r f e r e n c e e f f e c t s are BTL (J 18.4', w i t h good agreement noted i n c r e a s i n g l y i n f l u e n c e d by t h e p h y s i c a l between t h e model and f l i g h t p o i n t s . A t presence of t h e j e t plume a c t i n g as a M , fl 0.9, t h e decrement i n Ca is a b o u t s o l i d body and less by entrainment. The 3 x 10-4 u n i t s because of t h e r e d u c t i o n j e t plume i n c r e a s e s i n s i z e a f t e r l e a v i n g i n b o a t t a i l angle. The decrement i s t h e t h e n o z z l e and t e n d s t o raise t h e pres- same f o r t h e model and f l i g h t d a t a .

s u r e on t h e n o z z l e e x t e r n a l s u r f a c e , t h e r e b y reducing Cam These j e t flow E f f e c t of Angle of Attack phenomena appear t o e x p l a i n t h e v a r i a - t i o n i n Ca e x h i b i t e d by t h e model d a t a .

The e f f e c t of a n g l e of a t t a c k on t h e comparison of n o z z l e a x i a l f o r c e coef- F l i g h t test d a t a are shown i n f i g - f i c i e n t i s shown i n f i g u r e 38 f o r t h r e e u r e s 3 9 ( a ) t o 3 9 ( c ) f o r two a l t i t u d e s a t s u b s o n i c Mach numbers w i t h BTL 18.4O v a l u e s of NPRL c l o s e to t h e d e s i g n value.

and an NPRL range of 2.5 t o 5.0. A t The f a i r i n g f o r t h e f o u r test p o i n t s i n p r e s s u r e , t h e e f f e c t i v e nozzle p r e s s u r e f i g u r e 3 9 ( c ) f o r M , 0.9 f a v o r s t h e d a t a r a t i o probably exceeds t h e c a l c u l a t e d values. T h i s c a u s e s t h e j e t to be under- o b t a i n e d a t 6100 m. Figure 39 also shows expanded and e x p l a i n s t h e s h a r p r e d u c t i o n t h e j e t temperature and mass f l u x ratio i n Ca.

f o r a model and a f l i g h t test p o i n t f o r each of t h e f o u r c o r r e l a t i o n s ( t h a t is, t h e e i g h t shaded p o i n t s ) . A s d i s c u s s e d A t an NPRL v a l u e of 5.0 ( f i g . 3 9 ( d ) ) , i n r e f e r e n c e s 13 and 14, t h e product of t h e f l i g h t l e v e l of C a is about 5 x 10'4 t h e j e t gas c o n s t a n t and t h e j e t temper- u n i t s less than t h e model d a t a . T h i s a t u r e and t h e r a t i o of t h e j e t - t o - f r e e - disagreement is t h e same as t h a t noted stream mass f l u x are t w o parameters t h a t a t M , f( 0.9. Examining t h e j e t parame- can be used t o analyze t h e e f f e c t s of j e t

ters from M , - 0.9 t o 1.2, t h e d a t a show

temperature on nozzle a x i a l f o r c e d u r i n g t t h e r e is e s s e n t i a l l y no change f o r t h e overexpanded flow. I n t h i s study, o n l y m o d e l parameters, b u t t h e r e is an t h e j e t temperature w a s used because t h e g a s c o n s t a n t f o r t h e m o d e l and t h e f l i g h t i n c r e a s e i n j e t temperature and a reduc- data are v i r t u a l l y i d e n t i c a l ( t a b l e 3 ) . t i o n i n mass f l u x r a t i o f o r t h e f l i g h t test d a t a . B e c a u s e t h e j e t i s o p e r a t i n g A t Ma, = 0.6 and 0.8 ( f i g s . 3 9 ( a ) and i n an underexpanded c o n d i t i o n , any change i n entrainment caused by t h e change i n 3 9 ( b ) , r e s p e c t i v e l y ) , e x c e l l e n t agreement j e t mass f l u x r a t i o f o r t h e f l i g h t d a t a i n Ca e x i s t s between t h e f l i g h t d a t a a t i s n o t considered to be s i g n i f i c a n t . I n 6100 m and t h e model d a t a . A t M , = 0.9 a d d i t i o n , s i n c e t h e r e is no change i n ( f i g . 3 9 ( c ) ) , t h e m o d e l d a t a exceed t h e t h e d i f f e r e n c e i n C a between model and f a i r e d f l i g h t d a t a by a b o u t 5 x f l i g h t , and d e s p i t e t h e increase i n j e t u n i t s of Cam These r e s u l t s s u g g e s t t h e r e temperature going from M , = 0.9 t o 1.2, is no a p p a r e n t e f f e c t of j e t temperature t h e i n c r e a s e i n j e t temperature is l i k e - on t h e agreement of C a between model and w i s e n o t considered to be s i g n i f i c a n t .

f l i g h t . There is n e g l i g i b l e change i n E f f e c t of Reynolds Number t h e j e t parameters going from M , = 0.8 t o M , = 0.9, y e t t h e agreement is excel- F i g u r e 40 shows t h e e f f e c t of Reyn- l e n t a t M , = 0.8 and r e l a t i v e l y poor o l d s number on nozzle a x i a l force coef- a t M , 0.9.

f i c i e n t . D a t a are presented f o r s e v e r a l Mach numbers a t a = l o t o 2'. The NPRL range is f r o m 2.0 to 3 . 6 at BTL 1 8 . 4 O F i g u r e 3 9 ( d ) shows Ca as a func- ( f i g s . 4 0 ( a ) , 4 0 ( b ) , and 4 0 ( c ) ) ; t h e NPRL t i o n of NPRL o b t a i n e d a t M, = 1.2. The

i s 5.0 f o r BTL - 7.7O ( f i g . 4 0 ( d ) ) . I n

NPRL t e s t range f o r t h e m o d e l d a t a f i g u r e s 4 0 ( a ) and 4 0 ( b ) , f o r Ma 0.6 extended f r o m 5.0 t o 7.0, which i n c l u d e s and 0.8, r e s p e c t i v e l y , t h e r e appears t o t h e d e s i g n NPRL of 6.1. Over this range, be a n e g l i g i b l e e f f e c t of Reynolds number t h e model C a d e c r e a s e s from 20 x I on t h e d a t a . T h i s r e s u l t follows from t h e good agreement of model and f l i g h t t o 15 x 10-4 u n i t s , i n d i c a t i n g t h a t j e t v a l u e s of C a shown i n f i g u r e 39 a t t h e s e i n t e r f e r e n c e is dominated by t h e jet plume shape r a t h e r than by entrainment.

Mach numbers. A t M , 0.9 ( f i g . 4 0 ( c ) ) T h i s is explained as follows. As shown

and a t M , - 1 . 2 ( f i g . 4 0 ( d ) ) t h e d a t a

i n f i g u r e 2 1 ( d ) , m o d e l and f l i g h t fuse- i n d i c a t e a d e c r e a s e i n C a of approxi- l a g e p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s close t o t h e j e t e x h a u s t are a t l a r g e mately 4 x 10-4 t o 5 x 10-4 u n i t s over n e g a t i v e l e v e l s . S i n c e the j e t is t h e test range. The d e c r e a s e s i n C a e x h a u s t i n g i n t o a r e g i o n where p r e s s u r e shown are a t t r i b u t e d t o d i f f e r e n c e s i n is c o n s i d e r a b l y less than atmospheric 2 1 t h e upstream flows between t h e m o d e l and t u n n e l and i n f l i g h t on a t w i n - j e t t h e f l i g h t data.

f i g h t e r a i r c r a f t over a Mach number range o f 0.6 t o 1 2, a Reynolds number range E f f e c t of Mach Number on the o f 17.5 m i l l i o n t o 302.5 m i l l i o n , and an S e n s i t i v i t y of Axial Force angle-of-attack range of l o to 7 O . These C o e f f i c i e n t t o Reynolds Number tests provided d a t a bases from which d a t a were s e l e c t e d and compared to y i e l d t h e following conclusions.

The e f f e c t of Mach number on t h e s e n s i t i v i t y of nozzle a x i a l f o r c e coef- 1 . For Mach 0.6 and 0.8, a t an f i c i e n t t o Reynolds number (bCa/ARe) i s a n g l e of a t t a c k of l o t o 2O, m o d e l values shown i n f i g u r e 41. The parameter of nozzle a x i a l f o r c e c o e f f i c i e n t agreed ACa/ARe w a s o b t a i n e d by t a k i n g t h e c l o s e l y with f l i g h t v a l u e s o b t a i n e d a t an slopes of t h e f a i r i n g s i n f i g u r e 40.

a l t i t u d e of 6100 m. Above Mach 0.8, A s i n d i c a t e d i n f i g u r e 41, t h e rate of nozzle a x i a l f o r c e c o e f f i c i e n t s i n c r e a s e d change of C a w i t h R e is z e r o up t o f o r both sets of d a t a , b u t t h e m o d e l d a t a M , = 0.8, d e c r e a s e s t o a minimum a t were 5 x 10-4 u n i t s h i g h e r than t h e f l i g h t M , = 0.9, and i n c r e a s e s s l i g h t l y with d a t a a t Mach 0.9 and 1 2. Reynolds num- b e r e f f e c t s on nozzle a x i a l f o r c e c o e f f i - f u r t h e r i n c r e a s e i n M , t o 1.2.

c i e n t were noted o n l y a t Mach 0.9 and 1.2.

V a r i a t i o n s i n j e t temperature and mass Summary f l u x r a t i o d i d n o t a f f e c t t h e comparisons o f nozzle a x i a l f o r c e c o e f f i c i e n t .

The e f f e c t s of t h e test v a r i a b l e s on nozzle a x i a l f o r c e c o e f f i c i e n t are sum- 2. For Mach 0.6 and 0.8, p r e s s u r e marized as follows. A s M , i n c r e a s e d from c o e f f i c i e n t d i s t r i b u t i o n s on t h e e x t e r n a l 0.6 t o 0.8, a t a l o t o 2O, model and nozzle s u r f a c e of t h e l e f t nozzles of t h e f l i g h t C a o b t a i n e d a t 6100 m agreed model and t h e a i r p l a n e agreed i n l e v e l c l o s e l y and i n c r e a s e d s l i g h t l y . Above and trend. Recompression w a s good. A t M , m 0.8, C a i n c r e a s e d f o r both sets of Mach 0.9 and 1.2, both sets of d a t a showed flow s e p a r a t i o n , i n c r e a s i n g flow d a t a ; t h e model d a t a were 4 t o 5 x 10-4 complexity, and a loss i n recompression.

u n i t s h i g h e r than t h e f l i g h t d a t a a t Although p r e s s u r e d i s t r i b u t i o n s agreed M , * 0.9 and 1.2. A t M, = 0.6 and 0.9, i n t r e n d , d i f f e r e n c e s i n t h e l e v e l of reducing t h e BTL reduced C,. A t sub- p r e s s u r e c o e f f i c i e n t caused by e f f e c t s s o n i c speeds, peak v a l u e s of C a oc- upstream of t h e nozzle propagated a l o n g t h e nozzle s u r f a c e . T h i s caused c u r r e d near t h e d e s i g n p r e s s u r e ratio t h e f l i g h t d a t a t o e x h i b i t less n e g a t i v e

f o r t h e model d a t a . A t M , = 1 . 2, t h e

l e v e l s of p r e s s u r e c o e f f i c i e n t than t h e j e t appeared t o be expanding i n t o a m o d e l d a t a on a l l t h e nozzle r o w s .

r e g i o n w e l l below atmospheric p r e s s u r e , which i n c r e a s e d t h e e f f e c t i v e nozzle 3. P r e s s u r e c o e f f i c i e n t d i s t r i b u - pressure ratio. Reynolds number e f f e c t s t i o n s on t h e a f t f u s e l a g e s u r f a c e s f o r were noted o n l y for M , * 0.9 and 1.2, and model and f l i g h t were similar over t h e j e t temperature, mass f l u x ratio, and Mach number and angle-of-attack test a n g l e of a t t a c k had n e g l i g i b l e e f f e c t s on ranges. A t subsonic speeds, f l i g h t n o z z l e a x i a l f o r c e .

l e v e l s of Cp w e r e g e n e r a l l y less nega- t i v e than those of t h e m o d e l on t h e upper CONCLUS IONS s u r f a c e s and t h e lower n a c e l l e .

A f t f u s e l a g e , boundary l a y e r , and 4. Mach number and a n g l e of a t t a c k nozzle p r e s s u r e s were measured on a a f f e c t e d t h e model boundary l a y e r pro- 1/12-scale p r o p u l s i o n model i n a wind 5. Except f o r t h e e f f e c t of a n g l e of f i l e s and t h i c k n e s s e s a t t h e forward and attack on t h e boundary l a y e r parameters, a f t r a k e s t a t i o n s more than f o r f l i g h t .

i t s o v e r a l l e f f e c t s on t h e comparisons o f the p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s For t h e m o d e l , boundary l a y e r t h i c k - and nozzle a x i a l f o r c e were small.

n e s s a t t h e a f t rake s t a t i o n exceeded t h a t a t t h e forward rake s t a t i o n and i n c r e a s e d markedly w i t h a n g l e of a t t a c k .

A m e s Research C e n t e r I n f l i g h t , t h e boundary l a y e r t h i c k n e s s Dryden F l i g h t Research F a c i l i t y a t t h e a f t rake s t a t i o n w a s less t h a n National A e r o n a u t i c s and Space t h a t a t t h e forward rake s t a t i o n and Admini s t r a t i on decreased w i t h a n g l e of attack.

Edwards, C a l i f o r n i a , A p r i l 11, 1984 APPENDIX -DERIVATION OF PITOT PRESSURE

RATIO I N THE BOUNDARY LAYER F O R AN (5 + M2)l12 - - (5 + Ma2)l12

ASSUMED VELOCITY RATIO DISTRIBUTION Y Ma flight where The Mach number i n t h e boundary l a y e r is 2 2 5 + M = C M first d e r i v e d as follows. Assume a v e l o c i t y d i s t r i b u t i o n w i t h index n: 2 2

M - CM = -5

($n = u

M 2 ( 1 - C) = -5

" a J 2 5 M = - Also assume t h a t t h e t o t a l temperature c - 1 and s t a t i c p r e s s u r e i n t h e boundarv l a y e r are equal t o t h e free-stream values: M w a s e v a l u a t e d i n t h e boundary layer by e q u a t i o n ( 2 1 , u s i n g v a l u e s of H from 0 t o 6 where M , and 6 w e r e o b t a i n e d from M i - t h e f l i g h t test p o i n t . The v a l u e of n selected w a s 9.

To c a l c u l a t e t h e p i t o t p r e s s u r e r a t i o with M subsonic,

- -

P i = Pti

Q(l + O.2M2)'l2

Rearranqinq g i v e s 2 4 2 3 5 P i = (1 + 0.2M ) P,

- P t i = ( 56:112)115 ( 7M26- , Y m 5

= (, + 0.2Mw2)3-5pm

P i (5 + M2>3*5P, w i t h M supersonic, REFERENCES and Nozzle P r e s s u r e D i s t r i b u t i o n s of a 1/12TScale F-15 P r o p u l s i o n

1 . Webb, Lannie D. ; and Nugent, Jack: Model a t Transonic Speeds -

S e l e c t e d R e s u l t s of t h e F-15 Pro- E f f e c t o f Fuselage Modifications p u l s i o n Interactions Program. and Nozzle Variables. NASA A I A A Paper 82-1 041, June 1982. TP-2333, 1984.

2. Webb, Lannie D. ; Varda, Dominic A. ; 8. Peddrew, Kathryn H . : A User's Guide and Whitmore, Stephen A. : F l i g h t t o t h e Langley 16-Foot Transonic and Wind Tunnel Comparisons of t h e Tunnel. NASA TM-83186, 1981 .

I n l e t / A i r f r a m e I n t e r a c t i o n of t h e F-15 Airplane. NASA TP-2374, 1984. 9. M a r t i n s , Richard E. : F-15 Nozzle Afterbody I n t e g r a t i o n . A I A A 3. P e n d e r g r a f t , Odis C., Jr.: Fuselage Paper 74-1100, O c t . 1974.

and Nozzle P r e s s u r e D i s t r i b u t i o n s on a 1/12-Scale F-15 P r o p u l s i o n 10. L o r i n c z , D a l e J.: A Water Tunnel Model a t Transonic Speeds. NASA Flow V i s u a l i z a t i o n Study o f t h e TP-1521, 1979. F-15. NASA CR-144878, 1978.

11. Swann, M.R.; Duke, E.L.; 4. Nugent, Jack; T a i l l o n , Norman V.; and Enevoldson, E.K.; and Wolf, T.D.: P e n d e r g r a f t , Odis C., Jr.: S t a t u s of a Nozzle Airframe Study of a Experience With F l i g h t T e s t Tra- Highly Maneuverable F i g h t e r . A I A A j e c t o r y Guidance. A I A A Paper 78-990, J u l y 1978. Paper 81-2504, Nov. 1981.

, P l a n t , Thomas J . ; Nugent, Jack; and 12. Bergman, D . : Effects of Engine

Exhaust Flow on B o a t t a i l D r a g .

Davis, Robert A. : F l i g h t Measured E f f e c t s of Boattail Angle and Mach J. A i r c r a f t , v01. 8, no. 6, pp. 434-439, June 1971.

Number on t h e Nozzle Afterbody Flow o f a Twin-Jet F i g h t e r . A I A A Paper 80-0110, Jan. 1980. 13. P e t e r s , W. L.: A Comparison of J e t Temperature E f f e c t s on Afterbody Drag With Those From J e t Molecular 6. Nugellt, Jack; P l a n t , Thomas J . ; Davis, R o b e r t A.; and T a i l l o n , Weight and Nozzle A r e a R a t i o V a r i - Norman V.: P r e s s u r e s Measured i n a t i o n s . A I A A Paper 80-1 161, June 30-July 2, 1980.

F l i g h t on the A f t Fuselage and E x t e r n a l Nozzle of a Twin-Jet F i g h t e r . NASA TP-2017, 1983. 14. Compton, W i l l i a m B. 111: E f f e c t s of J e t Exhaust G a s P r o p e r t i e s on Exhaust S i m u l a t i o n and Afterbody 7. P e n d e r g r a f t , Odis C., Jr.; and Carson, George T., Jr.: Fuselage Drag. NASA TR R-444.

TABLE 1 . - LOCATIONS OF SURFACE PRESSURE

O R I F I C E S ON FUSELAGE ( d ) Lower l e f t n a c e l l e c e n t e r l i n e ( a ) Upper f u s e l a g e c e n t e r l i n e Airplane

Mode 1 I Mode 1 i Airplane 1

X F y/L x/L y/L -0.039 0.534 -0.038 0.531 0.079 0 0.077 0 -0.038 0 . 6 1 1 -0.038 0 . 6 1 1 0.155 0 0.154 0

--- -0.038

0.684 -0.038 0.684 0 . 2 1 7 0 a

---

0.297 0 a -0.036 0.808 -0.036 0.836c

---

0.382 0 a ' -0.036 0.845 -0.037 0.849 0.456 0 0.455 0 ~ -0.035 -0.035 0.870

c 0.873

0.534 0 0.533 0 0 . 6 1 1 0 0.597 0 0.684 0 0.684 0.004 0.808 0 0.810 0.004 0.843 0 0.840 0.003 Model Airplane 0.913 0 0.929 0

X/L I Y/L X/L I Y/L

0.684 0.038 0.684 0.037 (b) Lower f u s e l a g e c e n t e r l i n e

---

--- 0.867b 0.037

I Model I Airplane I

(f) Lower r i g h t n a c e l l e c e n t e r l i n e 0.684 Model Airplane 0.807 0.843 0.844 0 . 9 1 3 0 0.930 ( c ) Upper l e f t n a c e l l e c e n t e r l i n e ( 9 ) L e f t s i d e

Mode 1 I Airplane I

Model Airplane X/L -0.038 0.534 0.533 -0.038 0 . 6 1 1 0.609 -0.038 -0.038 0.684 0.677 -0.038 -0.038

- -- ---

-0.036 0 . 7 8 1 0.808 -0.036 -0.036 0.808 ( h ) R i g h t s i d e

--- ---

-0.036 0 . 82gb -0.036 -0.036 0.845 0.845

-ode1 1 Airplane I

--- ---

-0.036 0 . 857b

Y F x/L Y F

X F 0.872 -0.035 -0.035 0.873

---

-0.035 0.684 0.134 0.684 0.134 0 . 882b 3Canopy area n o t instrumented on t h e a i r p l a n e .

h n s t a l l e d f o r i n l e t i n t e r a c t i o n s program ( r e f . 2).

CRelocated t o clear a f t e r b u r n e r vent.

TABLE 1. - Concluded

( i ) Upper l e f t tailboom Mode 1 Airplane X/L Y/L X/L Y/L 0.843 -0.076 0.839 -0.077 0.873 -0.076 0.871 -0.076 0.913 -0.076 0.919 -0.081

TABLE 2. - LOCATIONS OF SURFACE

PRESSURE ORIFICES ON NOZZLE ( a ) BTL = 18.4O and BTL = 15.1' ( 1 ) R o w s 35OoI 62OI 182' x/L ~ Model Airplane Model Airplane 0.912 0.91 1 0.897a 0 . 898a 0.91 9 0.91 9 0.904 0.904 0.925 0.924 0.91 1 0.91 2 0.932 0.931 0.91 9 0.91 9 0.924 0.924 0.931 0.931 ( 4 ) R o w 302' I

I x/L

A i r p l a n e Model 0. 898a 0.897a Model Airplane 0.904 0.904 0.91 2 0.91 1 0.897a 0. 898a 0.91 9 0.91 9 0.91 1 0.912 0.924 0.924 0.925 0.924 0.928 0.928 0.931 0.932 0.931 0.931 0.934 0.934 a Located on t h e f i x e d p o r t i o n of t h e nozzle; a l l o t h e r loca- t i o n s are on t h e v a r i a b l e portion of t h e nozzle.

I 28

TABLE 2. -Concluded I ( 1 ) R o w s 350°, 62O, 1 8 2 O Model A i r p l a n e Model Airplane 0.91 2 0.91 1 0. 898a 0. 897a 0.91 9 0.919 0.904 0.904 0.925 0.924 0.91 2 0.91 1 0.932 0.931 0.91 9 0.91 9 0.925 0,924 0.932 0.931 ( 4 ) R o w 302O Model Airplane 0.897” 0. 898a Model A i r p l a n e 0.904 0.904 0.91 1 0.912 0.897a 0. 898a 0.91 9 0.91 9 0.91 2 0.91 1 0.925 0.924 0.924 0.925 0.928 0.928 0.932 0.931 0.932 0.931 0.935 0.934 aLocated on t h e f i x e d p o r t i o n o f t h e n o z z l e ; a l l o t h e r loca- t i o n s are on t h e v a r i a b l e p o r t i o n of t h e nozzle.

TABLE 3 . -TEST CONDITIONS USED FOR THE COMPARISONS ( a ) Model test c o n d i t i o n s f o r comparisons of s u r f a c e pressuresa T e s t NPRL p o i n t 1 0.60 0.99 1 8 . 4 1 8 . 4 1 . 0 0 17.7 8 2 0 . 6 0 1.00 1 8 . 4 18.4 1 . 5 0 1 7 . 8 3 3 0.99 18.4 1 8 . 4 2 . 0 0 1 7 . 8 5 0.60 4 0 . 6 0 0 . 9 9 18.4 1 8 . 4 3 . 0 0 1 7 . 9 6 5 0 . 6 0 0 . 9 7 18.4 18.4 3 . 9 0 1 7 . 9 5 6 0 . 6 0 2.98 18.4 1 8 . 4 3 . 0 0 1 7 . 9 6 7 0.60 4.98 1 8 . 4 1 8 . 4 3 . 0 0 1 7 . 8 6 8 0 . 8 0 1 . 0 0 1 8 . 4 1 8 . 4 1.00 2 0 . 3 13 9 0 . 8 0 1.00 1 8 . 4 18.4 2.50 2 0 . 3 8 10 0.80 1 . 0 1 18.4 18.4 3.50 20.3 8 1 1 0 . 8 0 1.00 1 8 . 4 1 8 . 4 4.60 20.4 7 12 0.80 3 . 0 0 1 8 . 4 18.4 2 . 5 0 2 0 . 3 9 0 . 8 0 4.99 18.4 18.4 2.50 2 0 . 3 9 14 0 . 8 0 7 . 0 1 1 8 . 4 1 8 . 4 3 . 5 0 2 0 . 3 10 21.8 15 0.90 1.00 1 8 . 4 18.4 0.99 17 16 0 . 9 0 0 . 9 9 18.4 1 8 . 4 2.50 21 . 4 15 17 0.90 0.99 1 8 . 4 1 8 . 4 3 . 6 0 21.3 16 18 0 . 9 0 1 . 0 1 18.4 1 8 . 4 4.50 21.2 15 19 0.90 1 . 0 1 18.4 1 8 . 4 5 . 0 3 2 1 . 1 14 .2 0 0.90 2.98 18.4 1 8 . 4 2 . 5 1 21.4 1 .oo 1 5 . 1 1 8 . 4 3 . 9 0 1 7 . 1 4 0.60 2 lb 1 . 9 7 1 8 . 4 5 . 0 0 20.2 12 0.90 1 5 . 1 2 2b 7 . 7 2 4 . 9 9 21.0 20 1 . 2 0 1.00 7.72 23b 7.72 6.53 20.8 16 1.20 1.00 7 . 7 2 2 qb 7 . 7 2 7 . 0 5 20.7 15 1.20 1 .oo 7.72 2 Sb a s t a b i l a t o r p o s i t i o n was O o for these tests.

bRakes were i n s t a l l e d .

TABLE 3 . - C o n t i n u e d (b) A i r p l a n e test c o n d i t i o n s for comparison of s u r f a c e p r e s s u r e s T e s t M, NPRL p o i n t ~ 0.63 1.09 6,266 18.2 19.5 2.65 143.8 6 2 0.63 1.23 6,251 18.2 19.5 2.65 143.1 5 0.64 1.00 10,169 18.1 19.7 2.84 100.0 8 4 0.60 3.11 6,081 17.8 19.7 2.76 148.1 5 0.61 3.22 6,120 18.2 19.6 2.42 139.4 7 6 0.60 3.00 10,279 18.1 19.5 2.86 92.5 4 7 0.60 5.05 6,303 18.2 19.6 2.57 136.9 8 8 0.60 5.19 6,155 17.8 19.6 2.79 146.9 8 9 0.61 5.38 6,331 18.2 19.6 2.58 137.5 6 10 0.61 5.79 10,520 2.74 18.1 19.6 91.3 6 11 0.81 1.11 6,075 18.8 19.5 2.95 194.4 8 12 0.81 1.12 6,082 2.96 18.8 19.6 193.8 8 13 0.79 1.55 6,291 18.8 19.6 2.59 178.8 8 14 0.82 3.09 6,132 2.54 18.0 16.0 189.4 9 15 0.80 3.12 6,216 17.8 19.6 3.31 194.4 9 16 0.82 3.17 6,151 18.0 16.2 2.55 188.8 9 17 0.80 3.01 10,637 18.2 19.6 2.94 115.6 9 1 8 0.79 3.02 13,180 17.8 19.3 2.92 82.5 9 19 0.81 5.01 10,660 18.1 18.6 3.29 118.8 10 20 0.81 5.02 10,655 18.1 18.6 3.29 119.4 10 21 0.80 5.00 13,639 17.8 19.2 4.06 75.0 9 22 0.80 5.01 13,713 17.7 19.2 9 3.74 78.1 0.80 23 6.23 13,635 17.8 18.5 3.14 78.8 9 24 0.88 0.89 18.9 19.7

6,397 3.05 201 . 3 10

25 0.87 0.98 6,402 18.9 19.7 3.05 200.0 9 26 0.91 1.05 6,385 18.1 15.4 2.87 203.8 11 27 0.90 1.20 10,699 18.1 19.3 3.63 131.9 14 28 0.91 3.05 5,912 17.9 19.3 4.19 228.1 15

29 0.90 3.10 6,146 17.9 19.2 4 . 32 221 . 9 14

30 0.89 3.08 10,663 18.1 19.2 3.61 131.3 13 31 0.89 3.09 10,669 18.1 19.4 131.3 12 3.62 32 0.89 3.09 13,764 17.9 18.5 3.53 86.9 12 33 0.90 3.42 13,565 17.8 19.3 4.20 85.6 13 34 O m 61 1.00 10,894 14.1 19.6 4.01 87.5 5 35 0.91 2.20 10,624 14.7 19.5 5.14 130.7 8 36 1.19 0.84 6,008 295.0 15 7.3 8.0 5.00 37 1.22 0.84 6,272 302.5 14 7.2 7.9 4.97 TABLE 3. - Concluded ( c ) Thermodynamic properties of j e t f o r selected test c o n d i t i o n s of f i g u r e 3 9 ( 1 1 Model tests T e s t G a s c o n s t a n t , R a t i o of T,, ME’RM p o i n t K J/kg K specific h e a t s 4 307 287.1 1.4 2.39 31 7 287.1 1.4 1.44 1 6 2 82 287.1 1.4 1 . 3 0 23 298 287.1 1.4 1.44 ( 2 ) F l i g h t tests T e s t Ta I G a s c o n s t a n t , R a t i o of M F R A p o i n t K J/kg *K s p e c i f i c heats 3 54 1 287.2 1 . 3 8 1.25 11 5 80 287.2 1.37 1.04 2 4 6 2 0 287.2 1.37 1.07 3 7 1824 287.5 1.27 0.54 3 2 TABLE 4. -TEST CONDITIONS FOR COMPARISON OF BOUNDARY LAYER PROFILES ( a ) Model tests PtW' T t W T e s t B T R , a, BTL, R e m X deg NPRL p o i n t Ea, deg de9 N / c m 2 0.60 2.99 18.4 18.4 2.00 17.5 10.23 2 6a 31 0 0.80 1 - 0 0 18.4 18.4 3.51 20.0 10.23 2 7a 320 2.98 18.4 18. 4 3. 50 20.0 10.24 0.80 2 8a 320 0.80 6.99 18.4 18. 4 3.52 20.0 10.24 2 9a 320 2.99 18.4 18. 4 0.90 3.60 20.9 10.24 3 O a 322 2.90 7.72 1.20 7.72 5.01 20.9 10.20 31b 336 a s t a b i l a t o r p o s i t i o n is - 2 O .

k t a b i l a t o r p o s i t i o n is 2 O .

( b ) F l i g h t tests PtW' T t W ' T e s t BTL' a, HP mR' NPRL R e a X 10'6 p o i n t Mw deg K m deg deg N / c m 2 0.60 3.50 3.1 81.3 2.38 10,684 38 18.0 19.6 251 19.6 2.8 203.8 7.26 281 5,910 39 0.82 1 . 2 8 18.1 40 0.80 3.14 18.0 19.5 2.0 125.0 2. 42 241 10,581 41 0.80 6.97 17.9 19.2 4.4 109.4 2.38 264 10,686 42 0.91 3.10 17.7 19.6 3.7 224.4 7.95 288 5,949 9.7 3.4 108.9 3.50 285 13,884 4 3 1.21 2.50 12.3

TABLE 5. - PITOT PRESSURE RATIOS OBTAINED

F R O M BOUNDARY LAYER RAKeS ( a ) Model ( 1 Forward rake 26 27 2 8 29 30 31 0 0 . 7822b 0.6667 0.6577 0.6435 0.5995 0.4736 1 82 0.8982 0.8491 0.8363 0.8035 0.8097 0.6778 11.58 0.9958 0.9929 0.9926 0.9837 0.9879 0.8430 17.98 0.9969 0.9931 0.9937 0.9927 0.9891 0.8624 0.9987 0.9958 0.9965 23.46 0.9957 0.9906 0.8901 29.56 0.9946 0.9896 0.9903 0.9898 0.9827 0.891 0 0.6681 0.661 3 0.6446 0.61 17 0.4826 41 045 0 . 776BC a s c a l e d t o f l i g h t .

b s u r f a c e s t a t i c p r e s s u r e d i v i d e d by Pt .

W C S t a t i c p r e s s u r e d i v i d e d by P t .

m ( 2 ) A f t rake cm 2 8 29 30 31

--- --- --- --- ---

3.05 0.8947 0.8480 0.8207 0.8086 0.7753 0.6063 12.19 0.9640 0.9691 0.9224 0.8800 0.8904 0.7499 18.29 0.9864 0.9905 0.9684 0.9055 0.9523 0.8261 24.38 0.9946 0.9899 0.9880 0.9427 0.9776 0.9003 30.48 0.9877 0.9569 0.9354 0.9388 0.931 2 0.8490 0.6507 0.641 4 0.631 8 0.61 76 41 - 4 5 0.4114 C S t a t i c p r e s s u r e d i v i d e d by Pt .

m TABLE 5 . -Concluded (b) F l i g h t (1 1 Forward rake H, c m 40 4 1 38 39 42 43 0.672 0.664 0 0.795b 0.666 0 . 6 1 1 0 . 4 9 1 0.826 0 . 7 6 0.903 0 . 8 1 4 0.768 0.647 0.835 0 . 8 3 1 0.856 1 . 4 0 0 . 9 1 1 0.823 0 . 7 8 1 0.662 0.879 0.914 3.99 0.942 0.864 0.824 0 0 738 0.928 0.962 7 . 8 7 0.968 0 . 9 1 3 0 0 869 0.824 00975 0.988 13.46 0.986 0.974 0.937 0.930 0.995 0.992 19018 0.997 0.995 0.978 0.980 0.999 0 . 9 9 1 24.64 1 006 1 . 001 0.997 0.984 1 . 0 0 1 0.983 2 9 . 4 1 1 005 1 003 1 . 001 0.985 0.999 0.975 33.93 1.007 1. 001 0.987 0.999 0.991 0.962 38.43 0.999 0.997 0.985 0.998 k u r f a c e static pressure divided by P t .

(2) A f t rake T e s t p o i n t H, c m 40 4 1 42 43 0.590 0.432 0 0.7541: 0.622 0.626 0.624 0 . 8 1 5 0 . 8 1 8 0.829 0.773 0.622 0.76 0.887 1 . 4 0 0 . 9 1 4 0.845 0.847 0.868 0.804 0.660 3 . 9 4 0.950 0.930 0.929 0.958 0.894 0 . 8 1 9 7.77 0.981 0.973 0.970 0.983 0 . 9 5 1 0.923 0.990 0.997 0.990 0.978 0.990 0.955 13.41 0.995 0.956 19.18 0.996 0.997 0.992 0.965 24.89 0.997 0.998 0.996 0 . 9 5 1 0.991 0.958 0 0 987 00955 2 9 . 1 1 00995 0.999 00 987 0 0 940 0.998 0.977 0.935 0.984 0.956 33. 83 0.991 0.987 0.954 38.43 0.987 1 . 0 0 1 0.969 0.929 burface static pressure divided by Pt .

TABLE 6. - BOUNDARY LAYER THICKNESSES MEASURED I N FLIGHT

Displacement thickness, c m Momentum thickness, c m i, c m Aft rake Forward rake Aft rake Forward rake A f t rake 1 . 8 1 . 0 1 . 4 0 . 7 17.5 38 2 3 . 1 12.3 2 . 3 1 . 0 1 . 5 0 . 7 39 22.3 1 6 . 1 2 . 1 0 . 9 1 . 4 0 . 7 40 2 2 . 0 1 . 1 0 . 5 0 . 8 0 . 4 7 . 0 4 1 13.3 3 . 0 1 .o 1 . 9 0 . 7 12.1 42 25.5 2 . 9 1.3 1 . 6 0 . 7 1 2 . 4 43 2 0 . 5

TABLE 7. - FIGURE CONTENT AND TEST CONDITIONS OF

FIGURES IN RESULTS AND DISCUSSION SECTION Test points ~~ Figure Figure content Model Flight 21, 29, 32 Comparisons of surface pressure 4, 9, 16, 23 coefficient distributions as Mach number is varied 22, 30, 33 Comparisons of surface pressure 9, 12 to 14 1 1 , 14, 22, 23 coefficient distributions as angle of attack is varied 23, 31, 34 Comparisons of surface pressure 4, 16, 21, 22 2, 26, 34, 35 coefficient distributions, nozzle recompression, and nozzle axial force as left boattail angle is varied Comparison of boundary layer 26, 28, 30, 31 38, 40, 42, 43 profiles and thicknesses as Mach number is varied Comparison of boundary layer 27 to 29 to 41 profiles and thicknesses as angle of attack is varied Comparison of recompression on 4, 9, 16, 23 the nozzle circumference as Mach number is varied at a (J l o to 2O Comparison of recompression on 3 to 5, 9 to 1 1 , 1 to 3, 1 1 to the nozzle circumference as 16 to 19, 23 13, 24 to 27, Mach number is varied at 36, 37 a (J l o to 2O 36 Comparison of recompression on 4, 16, 21, 22 2, 26, 34, 35 the nozzle as left boattail angle is varied at a (J l o to 2O 37 Effect of Mach number and left 3 to 5, 9 to 1 1 , 1 to 3, 1 1 to boattail angle on the compar- 16 to 19, 13, 24 to 27, ison of nozzle axial force 21 to 23 34 to 37 coefficient at a (J l o to 2O, NPRL 2 . 5 to 5 . 0 Effect of angle of attack on the 38 3 to 7, 9 to 14, 1 to 33 comparison of nozzle axial 16 to 20 force coefficient at NPRL 2 . 5 to 5 . 0 39 Effect of nozzle pressure ratio 1 to 5, 8 to 1 1 , 1 to 3, 1 1 to on the comparison of nozzle 15 to 19, 13, 24 to 27, axial force coefficient at 23 to 25 36, 37 a (J l o 3, 4, 9, 10, 16, 40 Effect of Reynolds number on 1 to 3 , 1 1 to nozzle axial force coefficient 17, 23 13, 24 to 27, 36, 37 41 Effect of Mach number on the rate 3, 4, 9, 10, 16, 1 to 3, 1 1 to 17, 23 of change of nozzle axial force 13, 24 to 27, coefficient with Reynolds number 36, 37 L-81-10,410 Figure 1. 1 / 1 2 - s c a l e F-15 p r o p u l s i o n model i n t h e NASA Langley t r a n - s o n i c wind t u n n e l .

F.S. 157.134 i FAG& 1s F.S.

ORSGFiAL 101-91

r F.S. 183.302 . QU A L I T ~ \

OF POOR

16.299 14.816 \ ~ ~ i l 35.878

f

L22.543 F.S.

6.481 Metric break Figure 2 . D e t a i l s of F-15 p r o p u l s i o n model geometry (dimensions i n c e n t i m e t e r s ) .

L-81-10.408 V i e w of F-15 p r o p u l s i o n model showing a f t e r b o d y d e t a i l s and Figure 3 .

m i l i t a r y power test n o z z l e ; BTL = 1 8 . 4 ' .

/ F.S. 165.423

*8.120 d

I diameter -

4 4.387 - 5.202

diameter !

( a ) M i l i t a r y power s e t t i n g , BTL = 1 8 . 4 ' .

F.S. 165.423 \ L - 8.2448-4

5.476 diameter

- - - -

I

7.026 f--. 4.384 + 6.030 1

diameter diameter ( b ) P a r t i a l afterburning power s e t t i n g , BTL = 1 5 . 1 ' .

Figure 4 . Drawings of t h e model n o z z l e s used i n t h e comparison (dimensions i n c e n t i m e t e r s ) .

0.399 8.722 6.081 diameter - -.4.341+

7-488 1

diameter V 7 Y (c) P a r t i a l a f t e r b u r n i n g power s e t t i n g ,

BTL = 7 . 7 ' . 0nGINAI: PAGE Is

OF POOR .QTJ !\T ,V'Y

F i g u r e 4 . Concluded.

ECN 9325 Figure 5 . Photograph of t h e F-15 a i r p l a n e i n f l i g h t .

- 3.43-

I U

19.43 b F i g u r e 60 Three-view drawing of t h e F-15 airplane (dimensions i n meters).

4 1 ECN 4041 5-001 A i r p l a n e a f t e r b o d y showing n o z z l e s i n t h e m i l i t a r y power n o z z l e F i g u r e 7 .

s e t t i n g .

Figure 8 . D e t a i l s of a i r p l a n e n o z z l e g e o m e t r y .

ORIGINAL PAGE I S

O F 1 POOR -QUALITY L-81-10,749 ( a ) Model i n l e t r e p l a c e d by f a i r i n g .

E 29406 ( b ) A i r p l a n e i n l e t .

Figure 9 . Model and a i r p l a n e i n l e t s .

Inlet reference Variable- Bypass door,

/- angle cowl /- open or closed

line 7

\ / /

Cowl rotation

/ point

( c ) I n l e t variable geometry.

Figure 9 . Concluded .

E 35115 Rear v i e w of the engine compartment a i r v e n t of F-15 a i r p l a n e .

( a ) Figure 1 0 . Engine compartment a i r v e n t .

L-81-10,454 Simulation of air vent on model w i t h a rearward f a c i n g s t e p .

(b) Figure 1 0 . Concluded .

-

( a ) A i r p l a n e . (b) Model.

Figure 1 1 . External n o z z l e s u r f a c e s .

4 5 ECN 24134 (a) F a i r e d - i n l e t c o n f i g u r a t i o n .

ECN 24136 ( b ) Flowing-inlet c o n f i g u r a t i o n .

Figure 1 2 . 1 / 4 8 - s c a l e f l o w v i s u a l i z a t i o n model.

ORIGINAL PAGE I S

OF POOR -QUALITY

Static pressure orifice Boundary layer rake Airplane only Model only Region over which com- parisons are made in this report

-* r

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 XIL ( a ) Upper s u r f a c e .

o Static pressure orifice Region over which com- parisons are made in this report Fuel dump vents-

--* r

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 XIL i (b) Lower s u r f a c e .

Figure 1 3 . Fuselage surface p r e s s u r e orifices.

XIL = 0.693 XIL = 0.707 0.635

-

3.556 4.13' Y I Parallel to nacelle centerline Left nacelle, XIL = 0.684 XIL = 0.871 1 . 5 2 4 - - + XIL = 0.881 Right nacelle, XIL = 0.867 (a) Drawings of model rakes.

Boundary layer rakes (dimensions in Figure 14.

centimeters).

L-81-10,754 ( b ) Rakes i n s t a l l e d on the model.

Figure 1 4 . Continued.

to nacelle f

LAirplane centerline surface Drawing of forward a i r p l a n e rake.

(c) Figure 1 4 . Continued.

ORIGINAL FASE !S

OF POOR QUALITY

( d ) Forward a i r p l a n e rake on t h e l e f t n a c e l l e .

Figure 1 4 . Concluded.

cp = o o

350 O, 6 orifices, I , 4 orifices

1820, 6 orif ices Ji

Figure 1 5 . Angular l o c a t i o n s of t h e 4 2 ori- fices on t h e l e f t n o z z l e s u r f a c e s of t h e model and a i r p l a n e ( l o o k i n g f o r w a r d ) .

5 1 - 0 Flight, Hp = 6,100 m 0 Flight, Hp = 10,700 m

-

0 Flight, Hp = 13,700 m

A Model Fairing - - Reynolds

-/

number .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 Mach number V a r i a t i o n o f test Reynolds number Figure 1 6 .

w i t h Mach number.

20 I - Test values of BTL 0 MIL 0 14.6O

0 9.50

n 7.70

-

Target values BTL,

deg 12 t

I I 1 I 1 I

.4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 M a Figure 1 7 . T y p i c a l v a l u e s o f l e f t n o z z l e b o a t t a i l a n g l e from f l i g h t tests (from r e f . 6).

BTL, deg 0 18.1 (MIL) 0 14.6 -

0 9.5

0 4

a 7.7

.4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 * a J Figure 1 8 . V a r i a t i o n of n o z z l e p r e s s u r e r a t i o w i t h f l i g h t test Mach number ( r e f . 6).

-

2.05

- l r M , = 0.6, H P = 10,700 m

2 .04 2.03 Error in Cp 2.02 - 2 .01 2 5

r

/M, = 0.6, Hp = 10,700 m

a

2 4

- 1 /M,= 0.6, HP = 6,100 m

2 3 Error in C, = 0.9, H = 10,700 m P

-

* 2

- OK 6 , l I m

2 1

I 1 I I I

0 1 2 3 4 5 q , , Nlcm2 Figure 1 9 . V a r i a t i o n of errors i n Ca and Cp w i t h f r e e - s t r e a m dynamic p r e s s u r e ( a s obtained i n ref. 6 ) .

O r

.1 I -

Type of

I Vehicle component A

flow field influencing flow

.* r

-

.4 Wing

-

XIL .5

-

.6

-

Nacelle, wing .7

-

.8 Vertical tail.

-

.9

-

1.0 ( a ) upper f u s e l a g e .

Figure 2 0 . I n f l u e n c e o f vehicle components on f u s e l a g e f l o w f i e l d s .

.1 .2 .3 Type of

Vehicle component I 1

flow field influencing flow

1 1

.4 f - XIL .5

I Compression

Nacelle, wing .6 .7

I

.8

- Fuselage and

Expansion- nozzle interfairing recompression boattailing, nozzle .9 - geometry

I Or expansion

1 .o

(b) Lower f u s e l a g e .

Figure 2 0 . Concluded.

Vertical tail Vertical tail trailing edge leading edge

T

' I I

'1

Lower

r fuselage

cP

- .2

- .4

m 2 r Upper nacelle

cP - .2 om

t

- .4 u

.4 .6 .8 1.0 X I L -.:p, 0 Flight 0 Model cP

- Fairing

- .4

---

Faired in accord- ante with flight I data trend

- .6

.4 .6 .a 1.0 XIL ( a ) MoD (LI 0.6; BTL * 1 8 . 4 ' ; BTR 1 8 . 4 ' ; and NPRL = 3 . 0 .

Figure 2 1 . Comparison of a f t e r b o d y p r e s - s u r e c o e f f i c i e n t d i s t r i b u t i o n s a s Mach number is varied a t a 1 O .

Vertical tail Vertical tail

leading edge7 7 trailing edge

' I 'I

' I

Lower fuselage fuselage

m 2 r upper

r

cP - .2 O m

1-

- .4 u

-2 r Upper nacelle Lower nacelle

r

cP

--:E

- .4

.4 .6 .8 1.0 X I L Upper 0 Flight tailboom 0 Model

- Fairing

cP - --c .4

---

i cp*

Faired in accord- ante with flight

- .6 L u data trend

.4 .6 .8 1.0 XI L (b) MaD = 0.8; BTL 1 8 . 4 ' ; BTR = 1 8 . 4 ' t o 1 9 . 5 ' ; and NPRL = 2.0 t o 3 . 0 .

Figure 2 1 . Continued.

Vertical tail Vertical tail

leading edge, \ trailing edge

vent

I I

Lower fuselage .' r Upper fuselage

r

cP

- .2

p

- .4 UIllrrl

a 2 r Upper nacelle r Lower nacelle

cP - .2 om

cP*

1 9

-

- . 4 1 , , f , .4 .6 .8 1.0

X I L

- .6

.Cp' cP 0 Flight Upper 0 Model tailboom

-

Fairing

- .6

--- Faired in accord.

ante with flight I data trend

- .8

.4 .6 .8 1.0 XI L

( C ) - 0.9; BTL * 1 8 . 4 ' ; BTR 18.4O t o

1 9 . 7 ' ; and NPRL = 2.0 t o 3.0.

Figure 2 1 . Continued.

Vertical tail Vertical tail leading edge- trailing edge

' I ' I ' I

r

cp -.2

- .4

t

r

O cP

- .2

- .4 L

.4 .6 .8 1.0 X I L

- .2

0 Flight Model cP

- Fairing

- .4

--- Faired in accord-

ante with flight data trend

- .6

.4 .6 .8 1.0 X I L Figure 21 . Concluded.

Vertical tail Vertical tail leading edge7 trailing edge Lower fuselage cP

- .2

L

- .4 u

Lower nacelle

.* r Upper nacelle

r

cP

- .2

- .4

.4 .6 .8 1.0 X I L cP - . 2 l T , J , 0 Flight t l Model

- Fairing

- .4

--- Faired in accord-

ante with flight data trend

- .6

.4 .6 .a 1.0 X I L ( a ) a 1'; BTL 18.4'; BTR 18.4' t o 19.5'; and NPRL 3 . 0 - Comparison o f a f t e r b o d y p r e s - F i g u r e 2 2 .

s u r e c o e f f i c i e n t d i s t r i b u t i o n s as a n g l e o f a t t a c k i s v a r i e d a t I , = 0 . 8 .

Vertical tail Vertical tail

leading edge, \ trailing edge

I t

Lower

r fuselage

cP - .2 O W F

- .4 -u

s 2 r Upper nacelle

r Lower nacelle

cP

- .2

- .4

.4 .6 .8 1.0 X I L

-.2m 0 0 Flight Model

cP

- Fairing

- .4

--- Faired in accord- ance with flight

- .6 data trend

.4 .6 .8 1.0 X I L (b) a rn 3'; BTL rn 18.4O; BTR 18.4O to 1 6 . 0 ' ; and NPRL rn 2.5.

Figure 22. Continued.

Vertical tail Vertical tail

leading edge? \ trailing edge

vent

' I ' I 'I

r Lower fuselage

.2 r Upper

fuselage

a

cP - .2 - t *

I-

- .4 u

a 2 r Upper nacelle r Lower nacelle

cP

- .2

- .4 .4 .6 .8 1.0 X I L

c Upper 7

- .2 tailboorn

0 Flight 0 Model

- Fairing

- .6

.4 .6 .8 1.0 X I L

I

( c ) U s ' 5'; BTL 18.4O; BTR 18.4' t o 19.2'; and NPRL 2.5 t o 3 . 7 .

Figure 2 2 . Continued.

I Vertical tail \Vertical tail leading edge, trailing edge vent

' I ' I 'I

r Lower fuselage

Upper fuselage

.2 r

0 - cP -m2pfy ,

- .4

r Lower nacelle

m 2 r Upper nacelle

0 ' .4 .6 .8 1.0 X I L tail boom cP

- .4

0 Flight 0 Model

- Fairing

- .6

.4 .6 .8 1.0 X I L = 1 7 . 8 " t o 1 8 . 4 ' ; ( d ) a = 6" t o 7 O ; BTL BTR 1 8 . 4 ; a d NPRL 3.0 to 3.5.

Figure 2 2 . Concluded.

Vertical tail Vertical tail

leading e d g e 1 \ trailing edge

' I \I

.2 Lower fuselage cP

- .2

r

- .4

.2 r Upper nacelle Lower nacelle

r

-.2 f

t " " "

- .4 - -

.4 .6 .8 1.0 X I L tailboom

cP - -:L:" - .4 .6

0 Flight 0 Model

- Fairing

.4 .6 .8 1.0 XI L ( a ) M , 0.6; BTL a 1 5 . 1 " ; BTR 1 8 . 4 " t o 1 9 . 5 " ; and NPRL 4 . 0 .

Figure 2 3 . Comparison o f a f t e r b o d y p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s as l e f t n o z z l e b o a t t a i l a n g l e is v a r i e d a t a 1" t o 2 ' .

Vertical tail Vertical tail leading edge7 trailing edge vent Lower

r fuselage

cP

- .2

- .4 L

r Lower nacelle

cP - .2 om

t

- .4 u

.4 .6 .8 1.0 XIL

-.:F

0 Flight 0 Model cP

-

Fairing

- .4

---

Faired in accord- ante with flight data trend

- .6

.4 .6 .8 1.0 XIL

(b) MaD - 0.6; BTL = 1 8 . 4 ' ; BTR - 1 8 . 4 ' t o

1 9 . 5 ' ; and NPRL 3 . 0 .

Figure 2 3 . Continued.

Vertical tail leading e d g e A F i : : e Lower fuselage

r

cP - .2 O W - p K

I I I

-.4 I

s2 r Upper nacelle

cP

- .2

- .4

U .4 .6 .8 1.0 X I L cP

- .4

0 Flight 0 Model

-

Fairing

- .6

.4 .6 .8 1.0 X I L (c) M- * 0.9; BTL 1 5 . 1 ' ; BTR * 1 8 . 4 O to 1 9 . 5 ' ; a d NPRL * 5 . 0 .

Figure 2 3 . Continued.

Vertical tail Vertical tail leading edge, trailing edge vent

'I 'I ' I

Upper fuselage Lower fuselage

r

cP

- .2

- .4 -

.2 r Upper nacelle Lower nacelle

r

cP

- .2

I

L

- .4

.4 .6 .8 1.0 X I L

- .2 t a Y E m

0 Flight 0 Model cP

- Fairing

- .4

---

Faired in accord- ante with flight data trend

- .6

.4 .6 .8 1.0 X I L

( d ) MaD 0.9; BTL - 1 8 . 4 ' ; BTR - 18.4' to

1 5 . 4 ' ; and NPRL = 2.5 to 2 . 9 .

Figure 23. Concluded.

0 Flight 0 Model, scaled

--- 1/9-power profile, flight

- Boundary layer thickness, flight Approximate thickness, model, scaled

-

Fairing M , = 0.9

- M , = 0.8

- M , = 0.6

P

- 8

I

i I

H, cm

-

I

, k g I I

/

.6 .8 1 .o

.8 1 .o .6 .8 1 .o .6 .8 1 .Q

.6 P IP P IP 'ilPt, i t , i 1 , ( a ) Forward rake, X / L = 0 . 6 8 4 .

2 4 . Comparison o f boundary l a y e r p r o f i l e s and t h i c k n e s s e s a s Mach number Figure is v a r i e d . a = 3 ' ; BTL m 12.3" to 18.4"; BTR 9.7' t o 29.6'; and NPRL 2.0 t o 5 . 0 .

0 Flight 0 Model, scaled

--- 119-power profile, flight

- Boundary layer thickness, flight E Approximate thickness, model, scaled Fairing

- -

M , = 0.6 M , = 1.2

P

- -

E z

H, cm I I

.8 1 .o .6 .8 1 .o

.6 .8 1 .o .8 1 .o

P IP P IP i t , PiIPt , ‘il‘t, I i t , ( b ) Aft rake, X/L = 0.867.

Figure 2 4 . Concluded.

0 Flight Model, scaled

---

119-power profile, flight I Boundary layer thickness, flight Approximate thickness,

E

model, scaled Fairing H, cm

.6 .8 1 .o

.6 .8 1 .o

P I P P I P P I P i t , i t , i t , ( a ) Forward rake, X/L = 0 . 6 8 4 .

F i g u r e 2 5 . Comparison o f boundary l a y e r p r o f i l e s and thick- nesses a s angle of a t t a c k i s v a r i e d . N , = 0.8': BTL = 18.4' t o 1 7 . 9 ' ; BTR 1 8 . 4 ' to 1 9 . 6 ' ; and NPRL 2 . 0 t o 4 . 4 .

0 Flight 0 Model, scaled

--- 119-power profile, flight

- Boundary layer thickness, flight E Approximate thickness, model, scaled

- Fairing

- a = 7 O - I H, cm

.6 .8 1 .o .6 .8 1 .o

.6 .8 1 .o

'il't, ( b ) A f t rake, X / L = 0 . 8 6 7 .

Figure 2 5 . Concluded.

ECN 25684 ( a ) Plan v i e w a t a = 1 '.

F i g u r e 2 6 . E f f e c t o f angle o f a t t a c k on f l o w p a t t e r n s o f f 1 owing-in1 et c o n f i g u r a t i o n .

ECN 25686 (b) Plan v i e w a t a = 3 O .

Figure 2 6 . Continued.

7 4 ECN 25687 ( c ) Plan v i e w a t a = 7".

Figure 2 6 . Continued.

ECN 25693 ( d ) S i d e view a t a = 1 *.

Figure 2 6 . Continued.

ORIGINAL I’Y\SE 2 OF POOR ‘QUALITY

ORIGINAL PAGE 7 s

POOR QUALl’Ty E C N 25694 ( e ) Side v i e w a t a = 3 O .

Figure 2 6 . Continued.

ECN 25696 ff) S i d e v i e w a t a = 7 O .

Figure 2 6 . Concluded.

DRIGINAL PAGE E J

D F POOR QUALITY

ECN 25688 (a) Plan view at a = 1 O .

Figure 27. Effect of angle of attack on flow patterns of faired-inlet configuration.

ECN 25689 ( b ) Plan v i e w a t a = 3 O .

Figure 2 7 . Continued.

ECN 25683 (c) Plan v i e w a t a = 7 " .

Figure 2 7 . Continued.

ORIGINAL PACT TT

DE POOR QUALITY

ECN 25697 ( d ) S i d e v i e w a t a = 1 O .

Figure 2 7 . Continued.

ECN 25698 ( e ) S i d e v i e w a t a = 3 O .

Figure 2 7 . Continued.

ECN 25692 (f) S i d e view a t a = 7 ' .

Figure 2 7 . Concluded.

ORTGINAL PAGE IS OF POOR 'QUALITY 8 4

+

1 Expansion, recompres- sion, attached flow

tl

I 2 Expansion, partial re- compression, flow separation

- 1

I 3 Expansion, standing shock wave, flow separation cP

J

Af terbody Centerline - Figure 2 8 . V a r i a t i o n s i n p r e s s u r e c o e f f i c i e n t a s s o c i a t e d w i t h a f t e r b o d y and n o z z l e f l o w i n t e r a c t i o n s .

0 Flight 0 Model m/zz Instrumented area

-

Fairing

--- Faired in accordance

with flight data trend

-2 r Upper nacelle

cP - -2 O W

I I

-.4 I

Lower surface

-2 r Lower nacelle

cP - .2 O m

1. Start of variable nozzle 2. Nozzle exit

- .4 3. Fuel dump vent

.4 .6 .8 1.0 -88 -90 mg2 -94 4. Leading edge of vertical tail XIL XIL M , m 0.6; BTL = 1 8 . 4 ' ; BTR = 1 8 . 4 ' : and NPRL 2 . 0 t o 3 . 0 .

( a ) Comparison o f a f t e r b o d y and n o z z l e p r e s s u r e c o e f f i c i e n t Figure 2 9 .

d i s t r i b u t i o n s a s Mach number i s v a r i e d a t a m 1 ".

0 Flight 0 Model '//////. Instrumented area

- Fairing

---

Faired in accordance with flight data trend

-2 r Upper nacelle r Nozzle row 350'

Upper surface cP

- .2

I ''I , 21, I - .4 Lower surface

-2 r Lower nacelle

-

cP - .2 O m

1. Start of variable nozzle 2. Nozzle exit

- .4

3. Fuel dump vent .4 .6 .8 1.0 .92 .88 '94 4. Leading edge of vertical tail X I L XIL (b) Mm = 0 . 8 ; BTL m 1 8 . 4 ' ; BTR 18.4' t o 19.5O; and NPRL 2 . 0 t o 3 . 0 .

Figure 2 9 . Continued.

0 Flight Model Y////L Instrumented area Fairing

---

Faired in accordance with flight data trend Nozzle row 350°

-2 r Upper nacelle

Upper surface

-

cp -.2

- .4

- .6

cP 1. Start of variable nozzle 2. Nozzle exit 3. Fuel dump vent .4 .6 .8 1.0 -88 .92 4. Leading edge of vertical tail XIL XIL ( C ) Mo. 0.9; BTL m 1 8 . 4 ° ; BTR 1 8 . 4 ' t o 1 9 . 7 ' ; a d NPRL m 2.0 t o 3 . 0 .

Figure 2 9 . Continued.

0 Flight 0 Model WA Instrumented area - Fairing

--- Faired in accordance

with flight data trend Nozzle row 350' Upper surface

r

-

cP I

-.4 I

Lower surface

-2 r Lower nacelle r Nozzle row 182'

cP

- .2

1. Start of variable nozzle 2. Nozzle exit I r 2 1 ,

- .4

..

3. Fuel dump vent .4 .6 .8 1.0 .88 .90 .92 .94 4. Leading edge of vertical tail XIL XIL Ma = 1.2; BTL PJ 7 . 7 ' ; BTR PJ 7 . 7 ' ; and NPRL PJ 5 . 0 .

( d ) Figure 2 9 . Concluded.

Flight Model '///L Instrumented area Fairing

---

Faired in accordance with flight data trend

m2 r Upper nacelle

cP

- .2

I 11. 21

- .4

Lower surface

m2 r Lower nacelle r Nozzle row 182O

I

C P P

cP - .2 O m

1. Start of variable nozzle 2. Nozzle exit

- .4

3. Fuel dump vent .4 .6 .8 1.0

.* 30 .92 '94 4. Leading edge of vertical tail

X I L XIL ( a ) a l o ; BTL w 1 8 . 4 O ; BTR 1 8 . 4 O t o 1 9 . 5 O ; and NPRL m 3 . 0 .

Figure 30. Comparison of a f t e r b o d y and n o z z l e p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s as angle o f a t t a c k is v a r i e d a t M, = 0 . 8 .

0 Flight 0 Model m k Instrumented area

-

Fairing

---

Faired in accordance with flight data trend

-2 r Upper nacelle r Nozzle row 350°

Upper surface

- .6 LA3-l

a 2 r Lower nacelle

#-

cP

- .2

1. Start of variable nozzle 2. Nozzle exit

- -4 L2.L

_ .

3. Fuel dump vent .4 .6 .8 1.0 .88 -90 .92 -94 4. Leading edge of vertical tail XIL XI L (b) a = 3 O ; BTL 1 8 . 4 ' ; BTR a 18.4' t o 16.0°; and NPRL 2 . 5 .

Figure 30. Continued.

0 Flight 0 Model w m Instrumented area

- Fairing

m2 r Upper nacelle

r row 3500 Upper surface

cP -+ % ,

- .4

Lower surface

r Nozzle row 182O

-2 r Lower nacelle

I n cP

- .2

1. Start of variable nozzle 2. Nozzle exit 3. Fuel dump vent

- .4

.4 .6 .8 1.0 J8 ' 9 0 '92 '94 4. Leading edge of vertical tail XIL XIL

(c) a - 5 ' ; BTL - 18.4'; BTR 18.4' to 19.2'; and NPRL 2 . 5

t o 3 . 7 .

Figure 30. Continued.

0 Flight 0 Model wm Instrumented area - Fairing

.2 r Upper nacelle Nozzle row 350°

Upper surface

-

cP

- .2

I 11, , 21,

- .4

a 2 r Lower nacelle

cP

- .2

1. Start of variable nozzle It, ,

I 2. Nozzle exit

- .4 L

3. Fuel dump vent .4 .6 .8 1.0 .88 .go .92 -94 4. Leading edge of vertical tail XIL XIL ( a ) a = 6" to 7 " ; BTL 1 8 . 4 ' ; BTR 1 8 . 4 ' : and NPRL = 3.0 to 3 . 5 .

Figure 30. Concluded.

0 Flight 0 Model w m Instrumented area

-

Fairing

m2 r Upper nacelle

Nozzle row 350' Upper surface

-

cP

- .2

- .4

Lower surface

s 2 r Lower nacelle r Nozzle row 182O

1. Start of variable nozzle

2 I

I 2. Nozzle exit ' I

- .4

3. Fuel dump vent .4 .6 .8 1.0 '88 '92 '94 4 . Leading edge of vertical tail XIL XIL M , a 0.6; BTL a 15.1O; BTR a 18.4' t o 1 9 . 5 ' ; and NPRL a 4 . 0 .

( a ) Figure 3 1 . Comparison o f a f t e r b o d y and n o z z l e p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s as l e f t nozzle b o a t t a i l angle i s v a r i e d a t a a 1 ' to 2 ' .

Flight Model Instrumented area Fairing Faired in accordance with flight data trend

a 2 r Upper nacelle r Nozzle row 3500

Upper surface

I

cP - .2 O W

to-rv

I 11, , 21,

-.4 I I I I

Lower surface

-2 r Lower nacelle :le row

cP

- .2

1. Start of variable nozzle I 2. Nozzle exit

- .4 u I I I

' 3. Fuel dump vent

.4 .6 .8 1.0 .92 .94 4. Leading edge of vertical tail XIL XIL (b) M , 0.6; BTL 1 8 . 4 ' ; BTR 1 8 . 4 ' t o 1 9 . 5 ' ; and NPRL 3 . 0 .

F i g u r e 31 . C o n t i n u e d .

0 Flight 0 Model '//////. Instrumented area

- Fairing

Nozzle row 350'

-2 r Upper nacelle

Upper surface

-

cp -.2

- .4

- .6 LLL

Lower surface

-2 r Lower nacelle

cP

LdfY

- .2

1. Start of variable nozzle 2. Nozzle exit

I 31

-.4 I I

3. Fuel dump vent .4 .6 .8 1.0 .94 4. Leading edge of vertical tail XIL XIL ( C ) M , m 0.9; BTL m 1 5 . 1 * ; BTR 18.4' to 19.5'; a d NPRL 510.

Figure 31 Continued.

0 Flight 0 Model mm Instrumented area

-

Fairing --- Faired in accordance with flight data trend

m2 r Upper nacelle Nozzle row 350°

Upper surface

-

cp -.2

- .4

- .6

02 r Lower nacelle

cP

- .2

1. Start of variable nozzle 2. Nozzle exit - A ..

3. Fuel dump vent -4 .6 .8 1.0 '88 '92 '94 4. Leading edge of vertical tail XIL XIL ( d ) M m = 0.9; BTL 1 8 . 4 ' : BTR 18.4' t o 1 5 . 4 ' ; and NPRL m 2.5 t o 2 . 9 .

Figure 3 1 . Concluded.

0 Flight Start of Nozzle 0 Model

e 7

I I I I

cP

- .2

cP

- .2

- .4

- .4

.88 .90 .92 .94 .88 -90 .92 .94 .88 .90 .92 .94 X I L XIL XIL M- m 0.6; BTL * 1 8 . 4 ' ; BTR 13 18.4O; a d ( a ) NPRL 3.0.

Figure 3 2 . Comparison of n o z z l e p r e s s u r e coeffi- c i e n t d i s t r i b u t i o n s as Mach number i s v a r i e d a t a m l o .

0 Flight Start of Nozzle 0 Model variable exit,

1 I

.;p, cP

- .2

- .4

p, ril?l

cP

- .2

- -4

.88 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 XI L XI L XIL Ma, a 0.8; BTL m 18.4'; BTR (I 18.4' t o 1 9 . 5 ' ; (b) and NPRL 2 . 0 t o 3 . 0 .

Figure 3 2 . Continued.

Start of Nozzle 0 Flight variable 0 Model

nozzle? ' ' 7

I I I I

-2 p = 3020 = 350° = 62'

cP

- .2

"[+, cP

- .2

- .4

-2 p = 230O

p = 1340

cP - .2 O W

-

I I I J

-.4 -88 . 9 0 .92 .94 .88 . 9 0 .92 .94 .88 . 9 0 .92 .94 X I L X I L XIL (c) Elo. ~ ~ 0 . 9 ; BTL 1 8 . 4 ' ; BTR 1 8 . 4 O t o 19.7.; and NPRL 2 . 0 t o 3.0.

Figure 3 2 . Continued.

Start of Nozzle 0 Flight variable 0 Model nozzle

exi7 7

I I

f r' I I

.2 = 302O ri = 350° = 62'

t - I

- .6 11111111

.ire, cP

- .2

- .4

.88 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 X I L X I L X I L

( d ) Mor, 1 . 2 ; BTL 7.7O; BTR - 7.7O; and

NPRL 5 . 0 .

F i g u r e 32. Concluded.

0 Flight Start of Nozzle 0 Model variable nozzle-/

I I

q = 350° cP

- .2

- .4 u u u

.Ip, cP

- .2

- .4

- .4

.88 .90 .92 .94 .88 -90 .92 .94 -88 .90 .92 .94 XI L XIL X I L ( a ) a = 1 " ; BTL ss 1 8 . 4 " ; BTR = 1 8 . 4 ' t o 1 9 . 5 ' ; and NPRL W 3 . 0 .

F i g u r e 3 3 . Comparison o f n o z z l e p r e s s u r e coeffi- c i e n t d i s t r i b u t i o n s a s angle o f a t t a c k i s v a r i e d a t M m = 0 . 8 .

o Flight

Start varmr;; of \Nozzle

0 Model exit I

I I I I I

.2

rq = 302O rq = 350°

rq = 62O

cP - .2 w w

- .4 u 11111111

cp = 860 cP

- .2

- .4 u k

cP

- .2

t-

- .4

.88 .90 .92 .94 .aa .go .92 .94 .aa .go .92 .94

X I L X I L X I L (b) a 3"; BTL 01 1 8 . 4 " ; BTR 1 8 . 4 " t o 16.0"; and NPRL I2.5.

Figure 33. Continued.

I

I 1 0 2 0 Flight Start of 0 Model variable r Nozzle

I I I I I nozz'7 ' I exit

rq = 350' rq = 62' m2 rq = 302'

cP - .2 'P w P

.Ir?, cP

- .2

- .4

- .4

.80 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 XI L X I L X I L (c) a II 5 " ; BTL II 1 8 . 4 O ; BTR Io 1 8 . 4 O to 1 9 . 2 ' ; and NPRL 2.5 to 3 . 7 .

Figure 33. Continued.

Start of o Flight

variable 0 Model n o z z l 7

I

I I I I

rq = 350° rq = 62O

m2 rq = 302°

-&

cP - .2 w

- .4 b

m2 T(p = 278O cp = 86O

cP - .2 O P

k

- .4 u

cp = 134O cp = 230° cP

- .2

P

- .4

.88 .90 .92 .94 .88 .90 .92 .94 . 8 8 .90 .92 .94 X I L X I L X I L (a) a IJ 6" t o 7 ' ; BTL 1 8 . 4 ' ; BTR 1 8 . 4 " ; and NPRL 3.0 to 3.5.

Figure 33. Concluded.

Start of 0 Flight variable 0 Model

i i i i

.2 cp = 3500 cP

- .2

- .4

- .4

.88 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 X I L X I L X I L Moo I ( 0.6; BTL 1 ( 1 5 . 1 ' ; BTR 1 8 . 4 t o 19.5'; ( a ) and NPRL 4 . 0 .

Figure 3 4 . Comparison of n o z z l e p r e s s u r e coeffi- c i e n t d i s t r i b u t i o n s a s t h e l e f t n o z z l e b o a t t a i l angle is v a r i e d a t a 1 ' to 2 ' .

o Flight Start of Nozzle 0 Model variable exitJ nozzle 7

I I

I I

.2 cP

- .2

- .4

*Ii-.; I

cP

- .2

- .4

.88 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 XIL X I L XIL (b) MaD 0.6; BTL ss 1 8 . 4 ; BTR 1 8 . 4 t o 1 9 . 5 ' ; a n d NPRL m 3 . 0 .

F i g u r e 3 4 . C o n t i n u e d .

I Start of 0 Flight variable Nozzle 0 Model nozzle7 exit

I I I I

.-

I q = 6 2 "

cP

- .2

- .4

cP

- .2

.;[?, [ J I

- .4

cP

- .2

.;fqTI [ ; ; ; ,

- .4

.88 .90 .92 .94 . 8 8 .90 .92 .94 .88 . 9 0 .92 .94 XIL X I L X I L 1 5 . 1 . ; BTR I ( 18.4' to 1 9 . 5 ' ; ( C ) M- m 0.9; BTL and NPRL = 5 . 0 .

Figure 3 4 . Continued.

Start of variable r Nozzle 0 Flight

nozzIeJ \, exit

0 Model

i i i i

m 2 rcp = 302O rcp = 350° rcp = 62O

cP

- .2

- .4 u LL

m2 rcp = 278O cp = 86O

cP

- .2

- .4 u k

.2

rcp = 230° cp = 134O

cP

- .2

- .4

.88 .90 .92 .94 .88 .90 .92 .94 .88 .90 .92 .94 X I L X I L X I L ( d ) Ma a 0.9; BTL 1 8 . 4 ' ; BTR 18.4 to 15.4O; and NPRL 2 . 5 to 2.9.

Figure 34. Concluded.

0 Flight 0 Model

- Fairing

--- Extrapolated Solid symbol indicates Last nozzle average value

cP I1

0 90 180 270 360 'p, deg M , = 0.6; BTL m 1 8 . 4 ' ; BTR = 18.4' to ( a ) 1 9 . 5 ' ; and NPRL = 2 . 7 to 3.0.

0 Flight 0 Model

- Fairing

--- Extrapolated

Solid symbol indicates average value r L a s t nozzle orifice

cP I1

First nozzle orifice in each row 1 1

- .4

0 90 180 270 360 'p, deg M , m 0 . 8 ; BTL = 1 8 . 4 ' ; BTR m 18.4' t o ( b ) 19.5'; and NPRL * 2 . 5 t o 3.0.

F i g u r e 3 5 . Comparison of r e c o m p r e s s i o n on the n o z z l e c i r c u m f e r e n c e a s Mach number is v a r i e d a t a = 1' t o 2 ' .

0 Flight 0 Model

-

Fairing --- Extrapolated Solid symbol indicates average value .2

r /Last nozzle orifice

in each row

- .2

I*

I I I I each row I

- .4

0 90 1 80 270 360 q , deg (c) M , 0.9; BTL ~ ~ 1 8 . 4 ' ; BTR m 1 8 . 4 ' t o 19.7'; and NPRL 2 . 5 t o 3 . 0 .

0 Flight 0 Model

-

Fairing

--- Extrapolated

Solid symbol indicates average value Last nozzle orifice 0 90 1 80 270 360 q, deg (a) Elm 1 . 2 ; BTL m 7.7'; BTR m 7.7'; and NPRL 5 . 0 ' .

Figure 3 5 . Continued.

at first orifice at last orifice AcP = (Cpavg

1 - ('Pavg

)7

C Pavg

t

cP orifice 0 Flight 0 Model

s4 r

0 -

L Last orifice

in each row7 0 Pavg First orifice -.4 +BTL = 18.4'+

t BTL = 7.7'

- .6 +

.6 .7 .8 .9 1.0 1.1 1.2 Mach number ( e ) Summary plot; BTL 18.4' and 7.7'; and NPRL 2.0 to 5 . 0 .

F i g u r e 3 5 . C o n c l u d e d .

1 1 1 0 Flight 0 Model Solid symbol indicates average value Last nozzle orifice

m 2 r A in each row -

cP

- .2

First nozzle orifice in each row

- .4

0 90 180 270 360 C P I deg ( a ) M , 0.6; BTL o ( 1 5 . 1 ' ; BTR 1 8 . 4 " t o 1 9 . 5 ' ; and NPRL * 4 . 0 .

0 Flight 0 Model Solid symbol indicates average value .2 orifice W cP

I1

- .2

- .4

0 90 180 270 360 C P I deg M , m 0.6; BTL m 1 8 . 4 ' ; BTR o ( 1 8 . 4 ' t o (b) 1 9 . 5 ' ; and NPRL 3 . 0 .

F i g u r e 36. Comparison o f recompression on the nozzle circumference a s the l e f t nozzle boat- t a i l angle is varied a t a = 1 ' t o 2 ' .

0 Flight 0 Model Solid symbol indicates average value .2 Last nozzle orifice cP

- .2

nozzle orifice in

leach row I I I

- .4

0 90 180 270 40, deg ( C ) Ma m 0.9; BTL 15.1'; BTR * 18.4O to 19.5'; and NPRL 5 . 0 .

0 Flight 0 Model Solid symbol indicates average value Last nozzle orifice in each row

m2 r 7

cP

- .2

I*

I I I each row I

-.4 L

0 90 180 270 360 'p, deg Ma m 0.9; BTL m 18.4O; BTR m 18.4' to ( d ) 15.4'; and NPRL 2.5 t o 2 . 9 .

Figure 3 6 . Concluded.

1 1 3 0 Flight, H = 6,100 m P d Flight, Hp = 10,700 m 0 Model

- Fairing

--- Reference 6

r

-*

--

OB--, *BTL = 7.7'- .6 .7 .8 .9 1.0 1.1 1.2 1.3 M m ( a ) Effect o f Mach number.

0 Flight, BTL 14.4' 0 Model, BTL = 15.1O

- Fairings from figure 37(a)

r

-Model

l6 t I

1 I I I 1 I I

.6 .7 .8 .9 1.0 1.1 1.2 1.3 M m Effect o f l e f t n o z z l e b o a t t a i l a n g l e .

( b ) F i g u r e 3 7 . Effect of Mach number and l e f t n o z z l e b o a t t a i l angle on t h e comparison o f n o z z l e a x i a l f o r c e c o e f f i c i e n t a t a = 1' t o 2 O and NPRL = 2 . 5 t o 5 . 0 .

0 Flight, Hp = 6,100 m

d Flight, Hp = 10,700 m

'0 Flight, Hp = 13,700 m

0 Model

- Fairing

r

p , = 0.9

0 2 4 6 8 a, deg Figure 38. Effect o f angle o f a t t a c k on the comparison of n o z z l e a x i a l force c o e f f i c i e n t a t BTL I18.4O and NPRL 2.5 to 5 . 0 .

0 Flight, H = 6,100 m P

d Flight, Hp = 10,700 m

0 Model

- Fairing

I Design pressure ratio

Solid symbols indicate test points for which jet parameters are given 12 x 10-4

8 c, #MFRA = 1.25

I MFRM = 2.39

~ 0 1 2 3 4 5 NPRL (a) M , 0.6; BTL 1 8 . 4 O .

0 Flight, H = 6,100 m P 0 Model

- Fairing

I Design pressure ratio

Solid symbols indicate test points for which jet parameters are given

r

1 2 1 T, = 580 O K , MFRA = i.o4\ 8 - 'a m 4 - MFRM = 1.44

I I

(b) M , 0.8; BTL 1 8 . 4 O .

Figure 39. E f f e c t o f n o z z l e p r e s - sure r a t i o on t h e comparison o f n o z z l e a x i a l f o r c e c o e f f i c i e n t a t a = 1 ' to 2 ' .

0 Flight, Hp = 6,100 m

d Flight, Hp = 10,700 m

IJ Model

- Fairing

I Design pressure ratio

Solid symbols indicate test points for which jet parameters are given MFRM = 1.30 0 1 2 3 4 5 NPRL (C) M , 0.9; BTL 1 8 . 4 ' .

0 Flight, Hp = 6,100 m 0 Model

- Fairing

I Design pressure ratio

Solid symbols indicate test points for which jet parameters are given 24 10-4

r

Tm = 298 O K , MFRM = 1 . 4 4 7 Ta = 1824 O K , MFRA = 0.54\ M , = 1.2; BTL = 7 . 7 ' .

( a ) Figure 39. Concluded.

0 Flight, H = 6,100 m P

d Flight, Hp = 10,700 m

0 Model

- Fairing

12 10-4

-

8 -

d

n

'a

0 8

4 -

I I I I I I

0 Flight, H = 6,100 m P 0 Model

- Fairing

0 . 8 ; BTL m 1 8 . 4 " ; and NPRL = 2.5 t o 3 . 5 .

(b) M , Figure 4 0 . E f f e c t of Reynolds number on n o z z l e a x i a l force c o e f f i c i e n t a t a m l o to 2".

0 Flight, H = 6,100 m P

d Flight, Hp = 10,700 m

0 Model

-

Fairing 20 x 10-4

r

'a 1 6 t

l2 t

" 0 50 100 150 200 250 300 x 106 Reynolds number 0.9; BTL = 1 8 . 4 O ; and NPRL = 2 . 5 to 3.6.

( c ) Moo 0 Flight, Hp = 6,100 m 0 Model

l2 t

8 1 I I I I I I

0 50 100 150 200 250 300 x 106 Reynolds number M - m 1 . 2 ; BTL = 7.7O; and NPRL = 5 . 0 .

( d ) Figure 4 0 . Concluded.

2 x 10-l2

r

- “ a

- 2 - 4 Figure 4 1 . Effect o f Mach number on t h e r a t e of change of n o z z l e a x i a l force coefficient w i t h Reynolds number.

2. Government Accession No. 3. Recipient's Catalog No.

1. Report No.

NASA TP-2588 5. Report Date 4. Title and Subtitle Comparison of Wind Tunnel and F l i s h t T e s t MARCH 1987 Afterbody and Nozzle P r e s s u r e s f o r a 6. Performing Organization Code Twin-Jet F i g h t e r A i r c r a f t a t T r a n s o n i c Speeds 8. Performing Organization Report No.

7. Author(s1 J a c k Nugent and Odis C. P e n d e r g r a f t , Jr. H-1214 10. Work Unit No.

9. Performing Organization Name and Address RTOP 533-02-21 NASA Ames Research C e n t e r 11. Contract or Grant No.

Dryden F l i g h t Research F a c i l i t y P.O. Box 273 Edwards, CA 93523-5000 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address T e c h n i c a l Paper N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n 14. Sponsoring Agency Code Washington, D.C. 20546

I

15. Supplementary Notes Odis C. P e n d e r g r a f t , Jr., i s a f f i l i a t e d w i t h NASA Langley Research C e n t e r , Hampton, V i r g i n i a .

16. Abstract A f t e r h o d y and n o z z l e p r e s s u r e s measured on a 1 / 1 2 - s c a l e model and i n f l i g h t on a t w i n - j e t f i g h t e r a i r c r a f t were cornpared as Mach number v a r i e d from 0.6 t o 1 . 2 , Reynolds number v a r i e d from 17.5 m i l l i o n t o 302.5 m i l l i o n , and a n g l e of a t t a c k v a r i e d from l o t o 7 O .

A t Mach 0.6 and 0.8, n o z z l e p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s and n o z z l e a x i a l f o r c e c o e f f i c i e n t s a g r e e d and showed good recompression.

A t Mach 0.9 and 1 . 2 , f l o w c o m p l e x i t y caused a loss i n recompression f o r b o t h f l i g h t and wind t u n n e l n o z z l e d a t a . The f l i g h t d a t a e x h i b i t e d less nega- t i v e v a l u e s of p r e s s u r e c o e f f i c i e n t and lower a x i a l f o r c e c o e f f i c i e n t s t h a n d i d t h e wind t u n n e l d a t a . Reynolds number e f f e c t s were noted o n l y a t t h e s e Mach numbers. Jet t e m p e r a t u r e and mass f l u x r a t i o d i d n o t a f f e c t t h e comparisons o f n o z z l e a x i a l f o r c e c o e f f i c i e n t .

A t s u b s o n i c speeds, t h e l e v e l s of p r e s s u r e c o e f f i c i e n t d i s t r i b u t i o n s on t h e upper f u s e l a g e and lower n a c e l l e s u r f a c e s f o r f l i g h t w e r e less n e g a t i v e t h a n t h o s e f o r t h e model.

The model boundary l a y e r t h i c k n e s s a t t h e a f t r a k e s t a t i o n exceeded t h a t f o r t h e forward r a k e s t a t i o n and i n c r e a s e d w i t h i n c r e a s i n g a n g l e of a t t a c k . The f l i g h t boundary l a y e r t h i c k n e s s a t t h e a f t r a k e s t a t i o n w a s less t h a n t h a t f o r t h e forward r a k e s t a t i o n and d e c r e a s e d w i t h i n c r e a s i n g a n g l e of a t t a c k .

7. Key Words (Suggested by Author(s)) 18. Distribution Statement Nozzle-afterbody flow i n t e r a c t i o n s U n c l a s s i f i e d - Unlimited Reynolds number e f f e c t s Wind t u n n e l and f l i g h t comparison 22. Price" 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO, of Pages I l n c l a s s i f i e d U n c l a s s i f i e d 124 A 0 6 *For sale by t h e National T e c h n i c a l I n f o r m a t i o n Service, Springfield, V i r g i n i a 22161.

NASA-Langley, 1987