Flight and wind-tunnel comparisons of the inlet-airframe interaction of the F-15 airplane

NASA

Technical

Paper

November 1984

Flight and Wind-Tunnel

Comparisons of the

Inlet)Airframe Interaction

of the F-15 Airplane

Lannie D. Webb, Dom Andriyich-Varda,

and Stephen A. Whitmore

NASA

NASA

Technical

Paper

Flight and Wind-Tunnel

Comparisons of the

Inlet/Airframe Interaction

of the F-15 Airplane

Lannie D. Webb, Dom Andriyich-Varda,

and Stephen A. Whitmore

Ames Research Center

Dryden Flight Research Facility

Edwards, Calfornia

National Aeronautics and Space Administration Scientific and Technical Information Branch SUMMARY The d e s i g n of i n l e t s and nozzles and t h e i r i n t e r a c t i o n s with t h e a i r p l a n e may account f o r a l a r g e percentage of t h e t o t a l d r a g of modern high-performance air- craft. This paper d e s c r i b e s t h e 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 program and t h e f l i g h t tests conducted a t t h e Dryden F l i g h t Research F a c i l i t y of t h e NASA A m e s Research Center. I n l e t - d r a g and l i f t d a t a from a 7.5-percent-scale wind-tunnel model are compared w i t h d a t a from a n F-15 a i r p l a n e with i n s t r u m e n t a t i o n t o match t h e model.

P r e s s u r e c o e f f i c i e n t v a r i a t i o n s with v a r i a b l e cowl a n g l e s , c a p t u r e ratios, and a n g l e s of attack are examples of flow i n t e r a c t i o n s p r e s e n t e d . Data are p r e s e n t e d f o r Mach numbers of 0.6, 0.9, 1.2, and 1.5.

INTRODUCTION A s t h e e f f o r t t o optimize a i r c r a f t performance over an i n c r e a s i n g l y l a r g e r f l i g h t r e g i o n c o n t i n u e s , more complicated f l i g h t c o n f i g u r a t i o n s evolve. Flow i n t e r - a c t i o n s of t h e i n l e t s , nozzles, and airframe can change t h e a i r c r a f t ' s s t a b i l i t y and t r i m drag, and i n c o n j u n c t i o n with o t h e r f a c t o r s , can r e s u l t i n i n c r e a s e d t o t a l a i r c r a f t drag. The need t o understand the e f f e c t s of these i n t e r a c t i o n s on high- performance a i r c r a f t having complex e x t e r n a l d e s i g n c h a r a c t e r i s t i c s has p r e c i p i t a t e d a wave of t e s t i n g and a n a l y s i s ( r e f . 1 ) . These s t u d i e s could do two things: p o i n t t o new d r a g d e t e r m i n a t i o n methods, and a i d i n t h e design of more e n e r g y - e f f i c i e n t a i r c r a f t .

To g a t h e r d a t a for u s e w i t h flow i n t e r a c t i v e s t u d i e s , a f l i g h t r e s e a r c h program cosponsored by NASA and t h e U.S. A i r Force w a s conducted a t t h e Dryden F l i g h t Research F a c i l i t y of t h e NASA A m e s Research Center ( A R C ) , u s i n g an F-15 a i r p l a n e ( f i g . 1 ) . The program w a s designed t o compare f l i g h t and wind-tunnel data for 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 and t o produce a data base f o r f u t u r e study. The wind- t u n n e l data w e r e o b t a i n e d from tests conducted i n t h e 14-Foot Transonic Wind Tunnel a t t h e NASA Ames Research Center ( r e f . 2). Follow-on tests were conducted by t h e A i r Force F l i g h t Dynamics Laboratory and t h e McDonnell Douglas Corporation i n t h e Arnold Engineering Development Center (AEDC) 16-Foot Transonic Wind Tunnel ( r e f . 3 . ) This report d i s c u s s e s i n d e t a i l t h e N A S A Ames Dryden i n l e t / a i r f r a m e i n t e r a c t i o n program, which w a s part of a l a r g e r program t o study nozzle-airframe i n t e r a c t i o n s (ref. 4 ) and t o perform a n e x t e n s i v e engine c a l i b r a t i o n test ( r e f . 5). The test c o n d i t i o n s and t h e i n s t r u m e n t a t i o n on both the model and t h e a i r c r a f t are described.

data, derived by p r e s s u r e i n t e g r a t i o n e q u a t i o n s , F l i g h t and wind-tunnel i n l e t d r a g are compared. These, i n t u r n , are compared w i t h t h e i n l e t d r a g measured by a f o r c e balance on t h e wind-tunnel model ( f i g . 2). The e f f e c t s of a n g l e of a t t a c k , v a r y i n g c a p t u r e r a t i o s , a movable c o w l , and o t h e r system components on p r e s s u r e flow f i e l d s a l o n g t h e a i r f r a m e are d i s c u s s e d .

NOMENCLATURE A i n l e t c a p t u r e as a f u n c t i o n of p and a, m2 (ft2) i n l e t c a p t u r e area a = 0 0 as a f u n c t i o n of p , m 2 ( f t 2 )

Ac

i n l e t c a p t u r e area a = p = 00, m 2 ( f t 2 ) Aco t o t a l i n l e t c a p t u r e d stream-tube area i n c l u d i n g d u c t and b l e e d flow, A 0 m 2 ( f t 2 ) mass flow r a t i o r e f e r e n c e d t o i n l e t c a p t u r e area as a f u n c t i o n of p and a AO/A mass f l o w ratio referenced t o i n l e t c a p t u r e area u = O o as a f u n c t i o n of p A o / A c mass flow ratio r e f e r e n c e d t o i n l e t c a p t u r e area a t a = p = Oo A o / A c o i n l e t t h r o a t c a p t u r e area, m 2 ( f t 2 ) A I

P m - Pw

pressure c o e f f i c i e n t , CP % a

cPlf - pw

lower f u s e l a g e p r e s s u r e c o e f f i c i e n t , CPlf q W

CPUf - P o 2

upper f u s e l a g e p r e s s u r e c o e f f i c i e n t , Cpuf q W C D I P i n l e t d r a g c o e f f i c i e n t CLIP i n l e t l i f t c o e f f i c i e n t a d d i t i v e drag, d e f i n e d i n f i g u r e I l ( d ) , N ( l b ) Dadd a x i a l f o r c e component a c t i n g on l o w e r c o w l , N ( l b ) FC1, normal force component a c t i n g on l o w e r cowl, N ( l b ) Fc ly a x i a l f o r c e component a c t i n g on upper cowl, N ( l b ) Fcux normal force component a c t i n g on upper cowl, N ( l b ) F cuY stream t h r u s t a t left-hand i n l e t p l a n e s e c t i o n , N ( l b ) FI a x i a l f o r c e component a c t i n g on inboard s i d e p l a t e , N ( l b ) F i spx normal f o r c e component a c t i n g on i n b o a r d s i d e p l a t e , N ( l b ) F i spy a x i a l f o r c e component a c t i n g on outboard s i d e p l a t e , N ( l b ) Fospx normal f o r c e component a c t i n g on outboard s i d e p l a t e , N ( l b ) FosPY free-stream stream t h r u s t , N ( l b ) F O Flr,F2r, f o r c e s a c t i n g on first, second, and t h i r d ramps, r e s p e c t i v e l y , normal t o ramp s u r f a c e , N ( l b ) F3r F.S. f u s e l a g e s t a t i o n , c m ( i n ) t r u e p r e s s u r e a l t i t u d e , m ( f t ) h W l e n g t h from t i p of nose t o end of tail boom ( L = 1890.3 c m (744.21 i n ) ) , L see f i g u r e 8 a d d i t i v e l i f t , d e f i n e d i n f i g u r e l l ( d ) , N ( l b ) Ladd f ree-stream Mach number MW

s u r f ace s t a t i c p r e s s u r e , N/m2 ( lb/f t2 )

P m

f ree-s tream t o t a l p r e s s u r e , N/m2 ( l b / f t2 )

PO

f ree-stream s t a t i c p r e s s u r e , N/m2 ( l b / f t2 )

Pce p u l s e -code modulation Pcm f ree-stream dynamic p r e s s u r e , N/cm2 ( l b / i n 2 ) s, R e Reynolds number based on l e n g t h of model o r a i r c r a f t l e n g t h r m s r o o t mean square free-stream v e l o c i t y , m/sec (ft/sec) VO X d i s t a n c e measured from t i p of a i r c r a f t nose, c m ( i n ) r a t i o of d i s t a n c e back from t i p of a i r c r a f t ' s nose t o L (see f i g . 8 ) X/L WAT2 corrected engine a i r f l o w , kg/sec (lb/sec)

a f ree-s tream a n g l e of a t t a c k , deg

free-stream a n g l e of s i d e s l i p , deg A I 1 , A g 2 , f i r s t , second, and t h i r d ramp a n g l e s r e l a t i v e t o w a t e r l i n e , deg AI3 P i n l e t r o t a t i o n a n g l e ( f i g . 3 ( b ) ) DESCRIPTION OF APPARATUS A i r p l a n e The F-15 a i r p l a n e ( f i g . 3 ( a ) ) i s a s i n g l e - s e a t , high-performance, a l l - w e a t h e r s u p e r i o r i t y f i g h t e r aircraft which h a s performance capabilities a t speeds i n excess of Mach 2. The a i r p l a n e h a s twin v e r t i c a l s t a b i l i z e r s , h o r i z o n t a l stabilators, a high-mounted swept-back wing, and t w i n F100-PW-100 a f t e r b u r n i n g t u r b o f a n engines.

The i n l e t system of t h e F-15 a i r p l a n e ( f i g . 3 ( b ) ) c o n s i s t s of t w o t w o - dimensional, e x t e r n a l compression, h o r i z o n t a l ramp i n l e t s . Each i n l e t h a s t h r e e ramps mounted i n an overhead arrangement that rotates a b o u t a t r a n s v e r s e h i n g e p o i n t a t t h e lower c o w l l i p . T h i s arrangement provides f o r a v a r i a b l e geometric c a p t u r e area. The second and t h i r d ramps, a l o n g w i t h t h e sideplates, are designed w i t h porous bleed h o l e s t o remove t h e lower energy boundary-layer flow. T h i s f l o w is exhausted through the louvered bleed e x i t s shown i n f i g u r e 3 ( c ) . I n a d d i t i o n t o t h e louvered b l e e d e x i t s , a variable bypass door p r o v i d e s f o r removal of l a r g e r of air. A t s u p e r s o n i c speeds, t h e bypass door i s modulated t o m a i n t a i n t h e amounts proper t h r o a t Mach number.

The i n l e t c o n t r o l system can be operated from one of t w o modes. I n one mode, t h e r o t a t i n g cowl, i n l e t ramps, and bypass door are a u t o m a t i c a l l y p o s i t i o n e d by t h e a i r i n l e t c o n t r o l l e r . The second mode used f o r t h e s e tests c o n s i s t e d of a manual i n l e t c o n t r o l system t h a t permitted t h e p i l o t t o set t h e i n l e t geometry t o any desired p o s i t i o n .

Wind-Tunnel Model The 7.5-percent wind-tunnel model ( r e f . 2 ) i s b o t h a f o r c e and p r e s s u r e model and is a s c a l e d v e r s i o n of t h e F-15 a i r p l a n e . The model has t w o b a l a n c e s : one t o measure f o r c e s on t h e l e f t i n l e t , and one t o measure f o r c e s on t h e e n t i r e a i r c r a f t ( f i g . 4 ) . The metric p o r t i o n of t h e i n l e t i n c l u d e s the upper and l o w e r cowls p l u s model w a s supported by t w o flow-through s t i n g t u b e s which b o t h s i d e p l a t e s . The s e r v e as i n l e t mass flow tubes.

The i n l e t system on t h e wind-tunnel model is shown i n d e t a i l i n f i g u r e 5. I n a d d i t i o n t o t h e bleed system, movable ramps, and c o w l , t h e i n l e t h a s a f i x e d - t h r o a t s l o t bleed/bypass e x i t .

Model-to-Airplane Comparison For t h e f l i g h t - t e s t program, t h e a i r p l a n e w a s i n s t r u m e n t e d w i t h p r e s s u r e o r i f i - ces on t h e l e f t s i d e of t h e f u s e l a g e , where much of t h e needed i n s t r u m e n t a t i o n w i r - i n g a l r e a d y e x i s t e d . With t h e e x c e p t i o n of a f e w a d d i t i o n a l p r e s s u r e orifices on t h e c e n t e r l i n e of t h e lower cowl l i p and i n t h e wing root area, t h e e x t e r i o r of t h e l e f t s i d e o f the a i r p l a n e w a s c o n f i g u r e d t o match t h e r i g h t side o f t h e wind-tunnel model ( f i g . 6 ) as c l o s e l y as possible. To d e f i n e i n l e t c o n d i t i o n s ( r e f . 61, three t o t a l p r e s s u r e probes were i n s t a l l e d i n t h e i n l e t and w e r e used i n a d d i t i o n t o t h e t o t a l p r e s s u r e probe f o r t h e a i r i n l e t c o n t r o l l e r . The a i r p l a n e ' s l e f t - h a n d secon- d a r y environmental c o n t r o l system (ECS) w a s covered w i t h a wedge t i p to match t h e model's boundary-layer d i v e r t e r . The F-15 a i r p l a n e h a s a noseboom P i t o t - s t a t i c probe system ( f i g . 3 ( a ) ) and a variable bypass door ( f i g . 3 ( b ) ) , both Of which t h e model lacked. A comparison between t h e a i r c r a f t ' s and t h e wind-tunnel modells bleed and bypass systems, which are l o c a t e d on t h e upper c o w l , is shown i n f i g u r e 7.

Figure 8 is p r e s e n t e d t o a i d i n t h e comparison of model and a i r p l a n e l o c a t i o n s .

relates major f u s e l a g e s t a t i o n s (used i n wind-tunnel r e p o r t s ) t o t h e X/L This f i g u r e r a t i o used i n t h i s r e p o r t .

INSTRUMENTATION To o b t a i n t h e same p r e s s u r e measurements on t h e F-15 a i r p l a n e as on t h e wind- t u n n e l model, over 150 s t a t i c p r e s s u r e p o r t s w e r e i n s t a l l e d f o r t h e i n l e t / a i r f r a m e i n t e r a c t i o n program.

P r e s s u r e measurement accuracy was increased by c o n t r o l l i n g t h e temperature of a l l p r e s s u r e t r a n s d u c e r s and by u s i n g a p r e s s u r e r e f e r e n c e system ( r e f s . 6 and 7 ) .

All s u r f a c e ports w e r e monitored by d i f f e r e n t i a l p r e s s u r e t r a n s d u c e r s having ranges of 21.4 N / c m 2 (22 l b / i n 2 ) , 22.8 N / c m 2 ( f 4 l b / i n 2 ) , and f4.1 N / c m 2 ( f 6 l b / i n 2 ) .

Wind-tunnel d a t a w e r e used p r i m a r i l y i n t h e s e l e c t i o n of t h e s e p r e s s u r e t r a n s d u c e r ranges with two o b j e c t i v e s i n mind: ( 1 ) t o minimize p r e s s u r e measurement e r r o r , and ( 2 ) t o maximize f l i g h t envelope coverage. The r e f e r e n c e s i d e of t h e s e t r a n s d u c e r s w a s connected to one of t h r e e tanks ( h i g h , medium, and l o w ) . The a b s o l u t e p r e s s u r e s i n t h e s e tanks were measured by p r e c i s i o n d i g i t a l - q u a r t z p r e s s u r e t r a n s d u c e r s . The l o w and medium r e f e r e n c e p r e s s u r e source i s a h e m i s p h e r i c a l probe ( r e f . 6 ) l o c a t e d on t h e r i g h t wing f a i r i n g . The l o w r e f e r e n c e p r e s s u r e tank is s u p p l i e d by f o u r manifolded s t a t i c o r i f i c e s l o c a t e d 90° from t h e s t a g n a t i o n port on t h e h e m i s p h e r i c a l s e n s o r , while t h e medium r e f e r e n c e tank is fed by t h e 60° port. The h i g h r e f e r e n c e p r e s s u r e source is a d u c t s t a t i c o r i f i c e near t h e l e f t engine compressor f a c e ( r e f . 6 ) . The d a t a were recorded d i g i t a l l y u s i n g a pulse-code modulation (pcm) sys- t e m , and w e r e both recorded on board a n d telemetered t o the ground.

PROCEDURE Wind Tunnel Tests The model w a s t e s t e d i n t h e AEDC 16-Foot Transonic Wind Tunnel t o o b t a i n i n l e t drag, l i f t , and i n t e r a c t i o n d a t a f o r primary Mach numbers of 0.6, 0.9, 1.2, and 1.5, o v e r a n g l e s of a t t a c k from - l o t o 1 7 O , and a t f i v e basic i n l e t c o n f i g u r a t i o n s .

E f f e c t s on t h e i n l e t / a i r f r a m e i n t e r a c t i o n from i n l e t r o t a t i o n a n g l e , second and t h i r d ramp a n g l e s , i n l e t b l e e d s , i n l e t mass flow r a t i o , Reynolds number, and h o r i - z o n t a l s t a b i l a t o r a n g l e w e r e determined ( r e f . 3 ) .

F l i g h t T e s t s From t h e AEDC wind-tunnel c o n d i t i o n s t e s t e d , c e r t a i n p o i n t s were s e l e c t e d t h a t w e r e considered primary f o r wind-tunnel-to-flight comparisons. The basic c o n d i t i o n s w e r e Mach numbers of 0.6, 0.9, 1.2, and 1.5, and a n g l e s of a t t a c k of Oo, 3 O , and 5O.

These c o n d i t i o n s were t e s t e d w i t h combinations of f i v e c o w l a n g l e s and t h r e e engine power s e t t i n g s t o o b t a i n v a r i o u s i n l e t capture ratios. A complete l i s t of d e s i r e d test c o n d i t i o n s is contained i n table 1 of r e f e r e n c e 6.

F l i g h t Conditions. - A m a j o r i t y of t h e primary f l i g h t p o i n t s w e r e flown a t a n a l t i t u d e of approximately 6.1 km (20,000 f t ) , a t Mach numbers of 0.6 and 0.9, and a n g l e s of a t t a c k of O o , 3 O , and 5O. A few f l i g h t p o i n t s w e r e flown a t an a n g l e of attack of 8 O and an a l t i t u d e of approximately 10.7 km (35,000 f t ) .

Two types of maneuvers w e r e performed i n o b t a i n i n g flight-to-wind-tunnel match p o i n t s . For a n g l e s of a t t a c k n e a r O o , a pullup-pushover maneuver w a s used, whereas f o r t h e higher a n g l e s of a t t a c k , a c o n s t a n t s u s t a i n e d t u r n w a s performed.

For m o s t test c o n d i t i o n s , t h r e e i n l e t c a p t u r e ratios (small, medium, and l a r g e ) w e r e d e s i r e d .

The numerator of t h e r a t i o w a s changed by v a r y i n g t h e engine mass f l o w ; t h e denominator w a s changed w i t h geometric c a p t u r e area. The p i l o t would run t h e engines asymmetrically w i t h t h e l e f t power l e v e r a n g l e a t e i t h e r i d l e , 80 per- c e n t , o r m i l i t a r y power, while t h e r i g h t engine w a s s e t t o maintain f l i g h t condi- t i o n s . Figure 9 shows engine a i r f l o w as a f u n c t i o n of Mach number f o r two engine power s e t t i n g s - i d l e and m i l i t a r y . Over t h e subsonic Mach number range, a v a r i e t y of engine a i r f l o w s are a v a i l a b l e .

A t t h e higher Mach numbers (1.2 and 1.51, which w e r e flown a t approximately 9.1 k m (30,000 f t ) , t h e r p m lockup l i m i t e d t h e a i r f l o w range and, hence, t h e range of c a p t u r e r a t i o s ( f i g . 9). An a t t e m p t w a s made a t M , = 1.2 t o expand t h e i n l e t c a p t u r e r a t i o by performing level d e c e l e r a t i o n s through M , = 1.2, a t reduced power s e t t i n g s . For t h e M , = 1.5 t e s t c o n d i t i o n , o n l y one a i r f l o w is p o s s i b l e , because of t h e rpm lockup which p r e v e n t s i n l e t buzz.

Uplink. - E s t a b l i s h i n g a f l i g h t p o i n t t o match a p a r t i c u l a r wind-tunnel condi- t i o n r e q u i r e d more t h a n s t a n d a r d i n s t r u m e n t a t i o n and f l i g h t techniques ( r e f s . 8 and 9). A s p e c i a l i n s t r u m e n t u s i n g two sets of n u l l i n g c r o s s p o i n t e r s w a s added t o permit t h e p i l o t to f l y a l t i t u d e and angle of a t t a c k simultaneously. T h i s i n s t r u - ment o p e r a t e s as follows: I n d i c a t e d a n g l e of a t t a c k and a l t i t u d e are t r a n s m i t t e d ( a l o n g w i t h o t h e r p a r a m e t e r s ) from t h e a i r p l a n e t o t h e c o n t r o l room, where t h e y are computer c o r r e c t e d and compared w i t h values i n p u t by t h e f l i g h t c o n t r o l l e r . The d i f f e r e n c e s between t h e d e s i r e d and a c t u a l values are t h e n t r a n s m i t t e d ( u p l i n k e d ) back t o t h e a i r c r a f t d i s p l a y d i a l . I n a d d i t i o n , f l i g h t and p r e s s u r e s e n s o r param- eters w e r e displayed i n t h e ground s t a t i o n on s t r i p c h a r t s and on cathode r a y tube (CRT) d i s p l a y s .

Tolerances. - Immediately following completion of a maneuver, t h e d e c i s i o n w a s

made whether t o r e p e a t t h e f l i g h t p o i n t based on a s e n s i t i v i t y a n a l y s i s of t h e i n l e t d r a g and l i f t equations. P e r t u r b a t i o n s of t h e s e e q u a t i o n s u s i n g i n c r e m e n t a l v a l u e s of f l i g h t parameters showed t h a t i n l e t d r a g and l i f t w e r e q u i t e s e n s i t i v e t o a n g l e of a t t a c k and, t o a lesser degree, t o Mach number and free-stream s t a t i c p r e s s u r e ( a l t i t u d e 1 . The study e s t a b l i s h e d t h e f o l l o w i n g as a c c e p t a b l e t o l e r a n c e s i n t h e f l i g h t parameters :

a, deg . . . . . 20.25

B, deg . . . . kO.25

M , e 20.01 Reynolds Number. - It w a s n o t p o s s i b l e t o match Reynolds numbers, s i n c e t h e wind-tunnel values were s i g n i f i c a n t l y lower than t h e f l i g h t v a l u e s ( f i g . 10). Most of t h e wind-tunnel runs w e r e a t t h e primary Reynolds number of 12 m i l l i o n . The f l i g h t Reynolds numbers ranged from 150 m i l l i o n t o 280 m i l l i o n .

DATA ANALYSIS Equations Used Wind-Tunnel A p p l i c a t i o n . - I n l e t d r a g and l i f t i n t h i s r e p o r t a r e plotted as a f u n c t i o n of t h e i n l e t v a r i a b l e c a p t u r e r a t i o , Ao/A. F i g u r e l l ( a ) shows t h e rela- t i o n s h i p of c a p t u r e d stream t u b e areas t o the geometric c a p t u r e a r e a f o r t h e F-15 i n l e t .

The d e f i n i n g e q u a t i o n s f o r i n l e t d r a g a r e shown i n f i g u r e l l ( b ) and a r e d i s - c u s s e d i n d e t a i l i n r e f e r e n c e 3 . I n general, t h e t o t a l i n l e t d r a g is t h e sum of t h e a x i a l p r e s s u r e f o r c e s a c t i n g over t h e e x t e r n a l i n l e t s u r f a c e s (upper and lower c o w l s , p l u s i n b o a r d and o u t b o a r d s i d e p l a t e s ) , p l u s t h e p r e s s u r e f o r c e a c t i n g on t h e unbound c a p t u r e d stream t u b e ( a d d i t i v e drag) between t h e f r e e - s t r e a m c o n d i t i o n s and t h e i n l e t - l i p c o n d i t i o n s (momentum change).

F i g u r e l l ( c ) i l l u s t r a t e s t h e r e l a t i o n s h i p of t h e c o n t r o l volume and i n l e t . The f i g u r e a l s o p r e s e n t s a d e t a i l e d breakdown of t h e b a s i c i n l e t d r a g and l i f t e q u a t i o n s and g i v e s t h e r e l a t i o n s h i p of the a d d i t i v e d r a g and l i f t t e r m s t o t h e s e e q u a t i o n s .

A l l t h e components of f o r c e are r e s o l v e d p a r a l l e l t o t h e l o c a l flow c o n d i t i o n s by t h e s i n a and cos a t e r m s .

The a d d i t i v e d r a g and l i f t t e r m s a r e composed of the normal f o r c e s on t h e t h r e e ramps, and t h e free-stream and i n l e t - p l a n e stream t h r u s t ( f i g . l l ( d ) ) . The f r e e - stream s t r e a m ' t h r u s t (Fo) t e r m has e i g h t p a r t s which account f o r a l l t h e p o s s i b l e a i r f l o w s . The parts of the a i r f l o w are engine, f i v e b l e e d s (second and t h i r d ramps, i n b o a r d and outboard, and t h r o a t s l o t ) , bypass door, and leakage.

F l i g h t A p p l i c a t i o n . - The i n l e t flow e q u a t i o n s i n v o l v e momentum t e r m s , free- stream flow q u a n t i t i e s , and pressures. To h a n d l e t h e s e terms, flight-measured p a r - ameters w e r e s u b s t i t u t e d i n t o t h e wind-tunnel e q u a t i o n s . F o r example, c a l i b r a t e d airflow ( r e f . 5 ) was s u b s t i t u t e d f o r t h e wind-tunnel c a l i b r a t e d n o z z l e air- e n g i n e flow. Corrected a i r - d a t a q u a n t i t i e s f r o m t h e c a l i b r a t e d P i t o t - s t a t i c probe on t h e noseboom w e r e used f o r free-stream Mach number ( M o o ) , f r e e - s t r e a m s t a t i c p r e s s u r e (pool, and f r e e - s t r e a m t o t a l p r e s s u r e (Po) i n t h e e q u a t i o n s . Corrected a n g l e of attack (a) and a n g l e of s i d e s l i p (6) were t a k e n f r o m t h e vanes on t h e noseboom.

Data S e l e c t i o n F l i g h t data a r e recorded onboard and i n t h e ground s t a t i o n and t h e n are proc- essed on a CDC CYBER 7 3 computer, a t which t i m e a l l c o r r e c t i o n s (such a s p r e s s u r e t r a n s d u c e r z e r o s and a i r - d a t a c o r r e c t i o n s from i n d i c a t e d t o t r u e ) a r e made. The d a t a are r u n through t h e wind-tunnel e q u a t i o n s t o c a l c u l a t e i n l e t d r a g and l i f t . A 10-sec i n t e r v a l of semisteady c o n d i t i o n s is t h e n s e l e c t e d t o match t h e wind-tunnel p o i n t f o r a n g l e s of attack and s i d e s l i p , Mach number, c o w l a n g l e , and t h i r d ramp a n g l e . For t h e f i n a l d a t a , a 1-sec t i m e p o i n t i s s e l e c t e d from t h e 10-sec i n t e r v a l based on s e n s i t i v i t y a n a l y s i s .

ACCURACY U n c e r t a i n t i e s i n t h e c a l c u l a t e d d a t a were a s s e s s e d p a r t i a l l y on t h e b a s i s of c a l i b r a t i o n s of t h e wind t u n n e l ( r e f . 3 ) and from r e s u l t s of t h e l a b o r a t o r y c a l i b r a - t i o n of the approximately 150 t r a n s d u c e r s used t o measure s u r f a c e pressures.

I n a d d i t i o n , the accuracy of t h e a i r - d a t a parameters , needed t o c a l c u l a t e i n l e t d r a g and l i f t , was a s s e s s e d based on (a) t h e l a b o r a t o r y c a l i b r a t i o n s of t h e a i r - d a t a t r a n s d u c e r s and ( b ) a n a l y s i s of t h e data from s e v e r a l d e d i c a t e d a i r s p e e d c a l i b r a t i o n f l i g h t s . From t h e s e tests and from a n error a n a l y s i s s t u d y p r e p a r e d by McDonnell A i r c r a f t Company f o r Ames Dryden (see appendix), f i g u r e 12 and table 1 w e r e prepared.

F i g u r e 1 2 ( a ) relates u n c e r t a i n t i e s i n t h e v a l u e s of Cp measured on t w o i m p o r t a n t areas used t o c a l c u l a t e i n l e t d r a g and l i f t ( t h a t is, upper and l o w e r cowls), over t h e f l i g h t range of dynamic p r e s s u r e s . The f i g u r e shows t h a t t h e Cp v a l u e s d e r i v e d from f l i g h t d a t a approach t h e 20.005 wind-tunnel accuracy ( r e f . 3 ) o n l y f o r s, > 4 N / c m 2 ( l b / i n 2 ) . I n f i g u r e 1 2 ( b ) , a 2a u n c e r t a i n t y i n i n l e t d r a g and l i f t from T h i s may be compared t o wind- f l i g h t d a t a (app.) is p l o t t e d as a f u n c t i o n of Ma.

t u n n e l u n c e r t a i n t i e s shown i n t a b l e 1 (ref. 3 ) and from d a t a from t h e appendix.

Surface-Pressure D i s t r i b u t i o n T h i s s e c t i o n of t h e r e p o r t i l l u s t r a t e s a few s e l e c t e d e f f e c t s of flow i n t e r a c - t i o n a l o n g t h e l e f t s i d e of t h e a i r c r a f t . Figure 13 shows p r e s s u r e changes a l o n g t h e upper and lower s u r f a c e s of t h e a i r c r a f t a t a f l i g h t c o n d i t i o n where Ma = 0.9, a = 3O, and p = O o . These changes are t h e r e s u l t of t h e varying amounts of spil- l a g e a i r t h a t r e s u l t from running t h e engine a t t h e t h r e e s e t t i n g s ( i d l e , 80 per- c e n t , and m i l i t a r y ) . Large v a r i a t i o n s i n C as a f u n c t i o n of X/L e x i s t over t h e P l f forward p o r t i o n of t h e lower cowl ( f i g . 1 3 ( a ) ) . P r e s s u r e c o e f f i c i e n t s on t h e l o w e r c o w l l i p ( x / L = 0.402) v a r y widely w i t h m a s s flow changes ( f i g . 1 3 ( a ) ) and a l s o w i t h o t h e r parameters ( r e f . 1 0 ) . These v a r i a t i o n s i n C d e c r e a s e r a p i d l y w i t h i n c r e a s - P l f i n g x/L i n both t h e wind-tunnel and f l i g h t d a t a . Over t h e upper s u r f a c e ( f i g . 1 3 ( b ) ) , less v a r i a t i o n with power s e t t i n g s is noted. The v a r i a t i o n i n C a t X/L = 0.476 Puf r e s u l t s from t h e f a c t t h a t t h e o r i f i c e on t h e a i r c r a f t is l o c a t e d j u s t a f t of t h e open bypass door b l e e d ( f i g . 7 ) . The l o c a t i o n of t h e metric break on t h e wind-tunnel model i s shown i n f i g u r e 1 3 f o r r e f e r e n c e . The t r e n d s and t h e l e v e l s of t h e d a t a are s i m i l a r t o values found a t o t h e r f l i g h t c o n d i t i o n s ( r e f s . 6 and 1 0 ) .

F i g u r e 14 i l l u s t r a t e s t h e v a r i a t i o n s i n CPlf and C w i t h X/L f o r two cowl Puf a n g l e s ( p = 4 O and 7') a t Moo = 0.9, a = 3 O , and a m i l i t a r y p o w e r e n g i n e s e t t i n g .

S t a r t i n g n e a r the l e a d i n g edge of t h e lower cowl l i p , v a l u e s of C are a f f e c t e d P l f by r o t a t i o n of t h e upper cowl ( f i g . 1 4 ( a ) ) . With i n c r e a s i n g v a l u e s of X/L, t h e I pressures r e a c h a minimum n e a r X/L = 0.418 and then r e t u r n t o n e a r free-stream con- d i t i o n s over most of t h e lower f u s e l a g e . Downstream of t h e lower cowl area, t h e e f f e c t s from cowl r o t a t i o n are m i n i m a l ( f i g . 1 4 ( a ) and r e f . 1 0 ) . Near t h e left-hand nozzle area, t h e flight-measured pressures became more n e g a t i v e a g a i n . On t h e upper s u r f a c e ( f i g . 1 4 ( b ) ) , e x c e p t f o r t h e leading-edge o r i f i c e (X/L = 0.303) and t h e o r i - f i c e a f t of t h e b y p a s s door b l e e d ( x / L = 0.476), cowl-angle e f f e c t s w e r e n o t l a r g e .

Downstream of the upper cowl t h e e f f e c t s of a v a r i a b l e cowl a n g l e w e r e n o t e v i d e n t , e x c e p t near t h e left-hand nozzle (X/L = 0.878). These e f f e c t s of cowl a n g l e on p r e s s u r e s i n the nozzle area w e r e found t o e x i s t almost t o t h e end of t h e nozzle ( r e f . 1 0 ) . A s i n t h e p r e v i o u s f i g u r e , wind-tunnel and f l i g h t d a t a show good agree- I ment i n most areas.

F i g u r e 1 5 i l l u s t r a t e s t h e v a r i a t i o n s i n C and C w i t h X/L f o r two a n g l e s of P l f Puf a t t a c k ( a = O o and 5 O ) a t M , = 0.9, p = 00, and a m i l i t a r y p o w e r engine s e t t i n g . In c o n t r a s t w i t h t h e p r e v i o u s f i g u r e where the e f f e c t s of p w e r e n o t sensed over much of t h e mid and a f t areas of t h e a i r c r a f t ( f i g . 1 4 ) , p r e s s u r e s v a r i e d w i t h a o v e r t h e e n t i r e a i r c r a f t ( f i g s . 15(a) and ( b ) ) . Large v a r i a t i o n s i n p r e s s u r e s w e r e found n e a r t h e l e a d i n g edges, while s m a l l ones were found t o e x i s t near t h e nozzle. Wind- t u n n e l and f l i g h t d a t a agreed w e l l over much of t h e range.

I n a d d i t i o n t o t h e i n t e r a c t i o n e f f e c t s from changes i n mass flow ( f i g . 1 3 ) , c o w l a n g l e ( f i g . 141, and a n g l e of a t t a c k ( f i g . 1 5 ) , many o t h e r v a r i a b l e s can a f f e c t p r e s s u r e s around t h e a i r c r a f t . For example, t h e e f f e c t s r e s u l t i n g from ori- f i c e l o c a t i o n o r a change i n s i d e s l i p were examined i n r e f e r e n c e 6, and r e f e r e n c e 10 c o n t a i n s forebody, boundary-layer d i v e r t e r , and nozzle i n t e r a c t i o n d a t a . Although o t h e r areas of t h e a i r c r a f t could be s t u d i e d , t h e d a t a i n t h e p r e v i o u s f i g u r e s , a l o n g w i t h r e f e r e n c e s 3, 6, and 10, p r e s e n t a s i z a b l e cross s e c t i o n of r e s u l t s .

I I n l e t Drag and L i f t

The preceding a n a l y s i s of f l o w - f i e l d i n t e r a c t i o n s around t h e i n l e t and a l o n g t h e f u s e l a g e r e p r e s e n t e d o n l y a p a r t of t h i s study. The o t h e r p o r t i o n c o n s i s t e d of t h e c o r r e l a t i o n of f l i g h t and wind-tunnel v a l u e s of i n l e t d r a g and l i f t o b t a i n e d by p r e s s u r e i n t e g r a t i o n s . From t h e s e comparisons, t h e f e a s i b i l i t y and accuracy of u s i n g t h e p r e s s u r e i n t e g r a t i o n technique t o determine i n l e t d r a g and l i f t from f l i g h t data may be a s s e s s e d . I n a d d i t i o n , t h e s e comparisons, along w i t h o t h e r f l i g h t and wind-tunnel s t u d i e s ( r e f s . 11 and 121, can g i v e i n s i g h t i n t o s p i l l a g e - d r a g e f f e c t s .

F i g u r e 16 shows t h e v a r i a t i o n of i n l e t d r a g (CDIP) and l i f t ( C L I P ) as a f u n c t i o n of Ao/A a t M , = 0.6, 0.9, 1.2, and 1.5 f o r p = Oo and A'3 = 11.2O. Figure 1 6 ( a ) compares t h e f l i g h t and wind-tunnel c a l c u l a t i o n of t o t a l i n l e t d r a g and l i f t a t M , = 0.6 and a = O O . The p r e s s u r e - i n t e g r a t e d d r a g s show reasonable agreement w i t h each o t h e r and w i t h t h e force-balance drag f o r l o w i n l e t - c a p t u r e r a t i o s . A t t h e h i g h e r i n l e t - c a p t u r e r a t i o , t h e p r e s s u r e - i n t e g r a t e d drags d i f f e r c o n s i d e r a b l y from t h e force-balance drag. This anomaly may be t h e r e s u l t of i n a b i l i t y of t h e p r e s s u r e i n t e g r a t i o n technique t o account f o r static p r e s s u r e o r flow a n g u l a r i t y g r a d i e n t s i n t h e i n l e t p l a n e , e s p e c i a l l y a t t h e lower Mach numbers ( r e f . 3 ) . The p r e s s u r e - i n t e g r a t e d d a t a w e r e f a i r e d f o r ease of comparison w i t h t h e force-balance d a t a and t h e two-dimensional p o t e n t i a l flow curve (ref. 3 ) . I n l e t l i f t shows good agreement among a l l t h r e e s o u r c e s . I n f i g u r e 1 6 ( b ) , f o r M , = 0.9, there is improved agreement among a l l t h r e e s o u r c e s of i n l e t drag. This is a t t r i b u t e d t o t h e reduced t u r n i n g of a t t h e i n c r e a s e d Mach number. I n l e t lift showed more scatter f o r t h i s con- t h e flow d i t i o n . Good agreement between f l i g h t and wind-tunnel d a t a w a s found a t t h e super- s o n i c Mach numbers (1.2 and 1.5) f o r both CDIP and CLIP. For t h e s e c o n d i t i o n s , o n l y one f l i g h t p o i n t w a s possible because of t h e engine l i m i t a t i o n s d i s c u s s e d i n t h e PROCEDURE s e c t i o n . I n all f o u r parts of this f i g u r e , the e f f e c t s of i n c r e a s e d s p i l - l a g e d r a g w i t h d e c r e a s i n g v a l u e s of Ao/A can be noted. A d e t e r m i n a t i o n of t h e abso- l u t e magnitude of s p i l l a g e d r a g i s n o t p o s s i b l e s i n c e v a l u e s of Ao/A never reached a " n o - s p i l l " c o n d i t i o n .

Figure 17 shows t h e v a r i a t i o n of C D I P and CLIP w i t h Moo f o r a = 0 0 and 50, res- p e c t i v e l y , and f o r p = O o . The f l i g h t and wind-tunnel data show reasonable agree- ment, and i l l u s t r a t e i n c r e a s i n g i n l e t d r a g and l i f t w i t h i n c r e a s i n g M , . Figure 18 p r e s e n t s t h e v a r i a t i o n of i n l e t d r a g ( f i g . 1 8 ( a ) ) and l i f t ( f i g . 1 8 ( b ) ) w i t h a f o r p = O o a t M , = 0.6 and 1.5, r e s p e c t i v e l y . Generally good agreement i n t h e d r a g t r e n d s and, t o a lesser degree i n t h e l i f t curves, are noted o v e r t h e a range. I n f i g u r e 19 t h e v a r i a t i o n of CDIP and C L I P w i t h p a t Moo = 0.6 and a = 00 i s p r e s e n t e d .

i n i n l e t d r a g is p r e s e n t i n a l l t h r e e methods a t A "dip" p = 4 O . The same t r e n d w a s found i n the wind-tunnel d a t a f o r o t h e r v a l u e s of M , and a ( r e f . 3 , p. 3 3 ) . I n l e t l i f t shows l i t t l e v a r i a t i o n w i t h p from - 4 O t o + 4 O . Some i n c r e a s e i n l i f t coef- f i c i e n t becomes e v i d e n t f o r p > 4 O .

I The preceding f o u r f i g u r e s have p r e s e n t e d a b r i e f overview of r e l a t i o n s h i p s of i n l e t d r a g and l i f t w i t h s e v e r a l o t h e r parameters: namely, Ao/A, M , , a, and p . I n g e n e r a l , u s i n g t h e p r e s s u r e i n t e g r a t i o n technique, t h e agreement between t h e f l i g h t and wind-tunnel v a l u e s of C D I P and C L I P w a s + l o counts and f 2 0 c o u n t s or better, ~ r e s p e c t i v e l y .

I CONCLUDING REMARKS The NASA A m e s Dryden i n l e t / a i r f r a m e i n t e r a c t i o n program h a s f u l f i l l e d i t s basic o b j e c t i v e s . Surface p r e s s u r e c o e f f i c i e n t s w e r e measured and matched, wind t u n n e l t o f l i g h t . I n t e r a c t i o n e f f e c t s determined from f l i g h t and wind-tunnel data have been compared and analyzed f o r s e v e r a l areas. Wind-tunnel p r e s s u r e i n t e g r a t i o n tech- n i q u e s f o r measuring and c a l c u l a t i n g i n l e t d r a g s and l i f t s w e r e s u c c e s s f u l l y adapted I t o f l i g h t t e s t i n g . Analysis of data taken d u r i n g t h e i n l e t / a i r f r a m e i n t e r a c t i o n program h a s r e s u l t e d i n the f o l l o w i n g conclusions: 1. I n general, m o s t of t h e f l i g h t and wind-tunnel s u r f a c e p r e s s u r e c o e f f i c i e n t s a g r e e d t o w i t h i n 20.01 t o f0.05.

2. P r e s s u r e i n t e g r a t i o n t e c h n i q u e s and e q u a t i o n s used f o r wind-tunnel t e s t i n g w e r e s u c c e s s f u l l y employed d u r i n g t h e f l i g h t - t e s t i n g program f o r i n l e t d r a g and l i f t determination.

3 .

P r e s s u r e - i n t e g r a t e d v a l u e s of i n l e t d r a g and l i f t were d e r i v e d from f l i g h t data and v e r i f i e d t h e wind-tunnel v a l u e s t o w i t h i n 210 counts or better f o r i n l e t drag and t o w i t h i n 220 counts o r better f o r i n l e t l i f t .

4 .

A s w i t h the wind-tunnel data, t h e f l i g h t i n l e t d r a g from p r e s s u r e i n t e g r a - t i o n d i d n o t a g r e e w i t h t h e wind-tunnel force-balance measurements of d r a g f o r c e r t a i n c o n d i t i o n s a t a Mach number of 0.6. The d i f f i c u l t y i n d e f i n i n g t h e flow c o n d i t i o n s a t t h e i n l e t e n t r a n c e p l a n e i s t h e p r o b a b l e cause.

Ames Research Center Dryden F l i g h t Research F a c i l i t y I Na t ioml Aerona u t ics a n3 Space Admi n i s t ra t i o n Edwards, C a l i f o r n i a , March 4 , 1983 APPENDIX - ERROR ANALYSIS An i n c r e a s e d understanding of t h e d r a g and l i f t d a t a is provided by i n c l u s i o n o f s e l e c t e d segments from t h e error a n a l y s i s s t u d y performed by McWnnell A i r c r a f t Company f o r NASA A m e s Dryden.

Nineteen f l i g h t and wind-tunnel combinations w e r e used as r e p r e s e n t a t i v e condi- t i o n s t o perform t h e e r r o r a n a l y s i s . u n c e r t a i n t i e s i n engine a i r f l o w , a n g l e of a t t a c k , i n l e t c o w l a n g l e , t o t a l and s t a t i c p r e s s u r e s , a l o n g w i t h t h e measurement u n c e r t a i n t i e s i n t h e p r e s s u r e t r a n s d u c e r s , were used i n t h e i n l e t i n t e g r a t i o n equa- t i o n s f o r t h e s t u d y . The mathematical model used f o r t h e e r r o r a n a l y s i s s t u d y c a n be d e s c r i b e d as a propagation-of-errors procedure (ref. 1 3 ) .

For t h e f l i g h t test r e s u l t s , t h e major c o n t r i b u t o r s t o t h e errors i n i n l e t d r a g and l i f t are as f o l l o w s ( i n g e n e r a l o r d e r of s i g n i f i c a n c e ) : free-stream s t a t i c p r e s s u r e , a n g l e of a t t a c k , i n l e t p l a n e t o t a l p r e s s u r e , free-stream t o t a l p r e s s u r e , and engine a i r f l o w .

For t h e r e s u l t s from t h e AEDC 16-Foot Transonic Wind Tunnel, t h e major c o n t r i b u - t o r s t o t h e errors i n i n l e t d r a g and l i f t a r e as f o l l o w s ( i n g e n e r a l o r d e r of s i g n i - f i c a n c e ) : f r e e - s tream s t a t i c p r e s s u r e , engine a i r f l o w , free-stream t o t a l p r e s s u r e , i n l e t p l a n e t o t a l p r e s s u r e , and a n g l e of a t t a c k .

M o s t of the t r e n d s i n t h e i n l e t d r a g and lift l e v e l s compare w e l l between f l i g h t t e s t and wind-tunnel r e s u l t s , even though the computed u n c e r t a i n t i e s f o r t h e f l i g h t t e s t p o i n t s are a b o u t an o r d e r of magnitude g r e a t e r t h a n f o r t h e small-scale model d a t a . A r e p r e s e n t a t i v e sample of the trends i n i n l e t d r a g and l i f t u n c e r t a i n t i e s is shown i n f i g u r e 20. The f i g u r e relates f 2 a u n c e r t a i n t i e s i n i n l e t d r a g and l i f t as a f u n c t i o n of Mach number f o r a = p = O o . In t h i s f i g u r e and f i g u r e 21, t h e wind- t u n n e l force-balance data s e r v e as a standard with which t h e p r e s s u r e - i n t e g r a t e d d a t a are compared. In f i g u r e 2 0 ( a ) t h e decreasing e f f e c t of free-stream s t a t i c p r e s s u r e u n c e r t a i n t y ( r e p r e s e n t e d by Mach number) on i n l e t d r a g i s i l l u s t r a t e d .

S t a r t i n g w i t h a l a r g e f2a u n c e r t a i n t y a t U , = 0.6, t h e force-balance and p r e s s u r e - i n t e g r a t e d f l i g h t data a g r e e c l o s e l y a t M , = 1.5. A similar t r e n d f o r i n l e t l i f t is noted i n f i g u r e 2 0 ( b ) .

Figure 2 1 i l l u s t r a t e s t h e v a r i a t i o n of C D I P and CLIP with a n g l e of a t t a c k f o r Also shown is a f 2 a e r r o r band a s s o c i a t e d with an angle-of- M , = 0.6 and p = Oo.

a t t a c k u n c e r t a i n t y of 20.5' (assumed f o r e r r o r a n a l y s i s ) . The f l i g h t and wind- t u n n e l p r e s s u r e - i n t e g r a t e d d r a g p o i n t s agree, b u t t h e force-balance-measured i n l e t d r a g f a l l s o u t s i d e of t h e f 2 0 u n c e r t a i n t y band f o r the t w o h i g h e r a n g l e s of a t t a c k o f 5 O and 8 O , r e s p e c t i v e l y . As w i t h CDIP, t h e p r e s s u r e - i n t e g r a t e d v a l u e s of f l i g h t and wind-tunnel C L I P d a t a f a l l w e l l i n s i d e t h e f 2 a band. However, j u s t as f o r C D I P , t h e force-balance d a t a f o r CLIP f e l l outside the u n c e r t a i n t y band a t a = 5 O and 8 O .

This g e n e r a l agreement i n d i c a t e s that reasonably a c c u r a t e i n l e t d r a g can be o b t a i n e d by p r e s s u r e i n t e g r a t i o n methods. It should be noted, however, t h a t c a r e f u l c o n s i d e r a t i o n must be g i v e n t o t h e pressure i n s t r u m e n t a t i o n l o c a t i o n s and t o t h e area assignments.

In summary, t h e e r r o r a n a l y s i s showed t h a t t h e u n c e r t a i n t y i n t h e F-15 p r e s s u r e - i n t e g r a t e d i n l e t d r a g d a t a f r o m f l i g h t - t e s t r e s u l t s ranges from f0.0003 t o fO.0010 ( 2 0 i n CDIP) compared to f0.0007 f o r small-scale wind-tunnel test r e s u l t s .

The u n c e r t a i n t y i n t h e F-15 p r e s s u r e - i n t e g r a t e d i n l e t l i f t from f l i g h t - t e s t r e s u l t s ranges from *0.0004 t o k0.0022 ( 2 0 i n CLIP) compared t o +O.oooi f o r wind- tunne 1 t e s t r e s u l t s .

The i n l e t d r a g and lift l e v e l s and t r e n d s o b t a i n e d from f l i g h t tests and wind- t u n n e l tests compared r e a s o n a b l y w e l l o v e r a w i d e range o f test c o n d i t i o n s .

I n l e t d r a g and lift u n c e r t a i n t i e s i n c r e a s e w i t h a l t i t u d e and a n g l e of a t t a c k ; t h e u n c e r t a i n t i e s d e c r e a s e w i t h i n c r e a s i n g Mach number and mass flow r a t i o .

The i n l e t drag u n c e r t a i n t y can be reduced i n f l i g h t tests by a c q u i r i n g data a t h i g h dynamic p r e s s u r e c o n d i t i o n s ( t h a t is, a t minimum allowable a l t i t u d e f o r a g i v e n .

f l i g h t c o n d i t i o n ) For t h e f l i g h t - t e s t r e s u l t s , t h e free-stream static p r e s s u r e and angle-of - a t t a c k measurements are t y p i c a l l y t h e major c o n t r i b u t o r s to t h e i n l e t d r a g u n c e r t a i n t y .

Because of t h e u s e of d i f f e r e n t i a l p r e s s u r e t r a n s d u c e r s , t h e r e f e r e n c e p r e s s u r e measurement can also become a s i g n i f i c a n t c o n t r i b u t o r t o t h e p r e s s u r e - i n t e g r a t e d i n l e t force u n c e r t a i n t y .

A d d i t i v e (spillage) and l o w e r cowl d r a g forces are t h e most s i g n i f i c a n t compo- n e n t s of the i n l e t d r a g . The u n c e r t a i n t y i n t h e a d d i t i v e d r a g components is the largest c o n t r i b u t o r t o t h e t o t a l d r a g u n c e r t a i n t y .

1 2 REFERENCES 1 . Smith, Ronald H.: Propulsion-Airframe I n t e r a c t i o n s P r e d i c t a b i l i t y .

Performance P r e d i c t i o n s Methods, AGARD-CP-242, May 1978, pp. 6-1 t o 6-20.

2. Rejeske, J. V.; and P o r t e r , J. L.: I n l e t / A i r c r a f t Drag I n v e s t i g a t i o n .

AFFDL-TR-74-34, A i r Force F l i g h t Dynamics Lab., Wright-Patterson AFB, Ohio, A p r . 1974.

3. Kamman, J. H.; and Wallace, H. W.: Assessment of I n s t a l l e d I n l e t Forces and 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 . AFFDL-TR-76-62, A i r Force F l i g h t Dynamics Lab., Wright-Patterson AFB, Ohio, J u l y 1976.

4. Nugent, Jack; T a i l l o n , Norman V.; and Pendergraft, O d i s C., Jr.: S t a t u s of a Nozzle-Airframe Study of a Highly Maneuverable F i g h t e r . A I A A Paper 78-990, J u l y 1978.

5. Kurtenbach, Frank J. : Comparison of C a l c u l a t e d and Altitude-Facility-Measured T h r u s t and A i r f l o w of Two P r o t o t y p e F100 Turbofan Engines. NASA TP-1373, 1978.

6. Webb, Lannie D.; Whitmore, Stephen A . ; and Janssen, R a n d a l l L.: P r e l i m i n a r y F l i g h t and Wind Tunnel Comparisons of t h e 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. A I A A Paper 79-0102, Jan. 1979.

7. Schweikhard, W i l l i a m G.: T e s t Techniques, I n s t r u m e n t a t i o n , and D a t a Processing.

D i s t o r t i o n Induced Engine I n s t a b i l i t y , AGARD-LS-72, Nov. 1974, pp. 6-1 t o 6-43.

8. Swann, M. R.; Duke, E. L.; Enevoldson, E. K.; and WOlfe, T. D.: Experience w i t h F l i g h t T e s t T r a j e c t o r y Guidance. A I A A Paper 81-2504, Nov. 1981.

W . , Jr.; Myers, Lawrence P.; Nugent, Jack; Lasagna, P a u l L.; and 9. Burcham, Frank Webb, Lannie D.: Recent P r o p u l s i o n System F l i g h t T e s t s a t the NASA Dryden F l i g h t Research Center. A I A A Paper 81-2438, Nov. 1981.

10. Webb, Lannie D.; and Nugent, Jack: S e l e c t e d R e s u l t s of the F-15 P r o p u l s i o n I n t e r a c t i o n s Program. A I A A Paper 82-1041 , June 1982.

11. Arnaiz, Henry H.; P e t e r s o n , John B., Jr.; and Daugherty, James C.: Wind- T u n n e l / F l i g h t C o r r e l a t i o n Study of Aerodynamic C h a r a c t e r i s t i c s of a Large C r u i s e Airplane (XB-70-1). NASA TP-1516, 1980.

F l e x i b l e Supersonic 12. Schoelen, F. J.; P e t e r s e n , M. W.; Kostin, L. C.; and G i l b e r t s o n , M.: I n l e t / Nozzle F l i g h t Performance Determination (Wind Tunnel T e s t Program).

Volume I - Program R e s u l t s . AFFDL-TR-78-173, A i r Force F l i g h t Dynamics Lab., Wright-Patterson AFB, Ohio, A p r . 1979.

13. Mandel, John: S t a t i s t i c a l Analysis of Experimental D a t a . John Wiley and Sons, New York, 1967.

Technique CDIP CLIP

' m a, deg CP

( source 1 0 . 6 to 1.5 0 0 to 80 f0.01 to f0.05" +3 x 10-4 to 24 x lo"+ to Pressure + i o x 10-4 +22 x 10-4 integration (flight 1

Pressure 0 . 6 to 1.5 O o to 5 O --- +0.7 x 10-4 f i x 10-4

integration (wind tunnel) Force balance 0.9 50 0.005 + 3 x 10-4 26 X l o e 4 (wind tunnel, ref. 3 ) ECN 9325 F i g u r e 1 .

F-15 a i r p l a n e i n f l i g h t .

F i g u r e 2 .

S m a l l - s c a l e (7.5-percent) inlet model i n AEDC 1 6 - F m t Transonic W i r d T u m l .

Y Y "

A

p--1943.10 (63.75) - j

( a ) Three-view d r a w i n g . D i m e n s i o n s are i n centimeters ( f e e t ) .

F i g u r e 3 . F - 1 5 a i r p l a n e an3 i n l e t .

Second and thl&

Rotating cowl / // ramp b l : exits

Bypass door - P Bypass door bleed exit

Second ramp’ 1 / LThroat slot bypass

Third ramp Cowl rotatlon plvot Note: Inboard and outboard sideplate bleeds are not shown ( b ) Ramps, bleed exits, and b y p a s s door e x i t s , E 34013 Second and t h i r d ramp b l e e d ( c ) e x i t s , F i g u r e 3 . C o n t i n u e d .

E 35054 ( d ) Bypass door b l e e d e x i t s .

Figure 3 . C o n e l uded.

,-Complete aircraft metrlc (force balance) mass flow tubes Remote airflow measurements Pressure-instrumented right inlet Force balance (independent inlet drag) -to-sting tube metric seal Figure 4 . I n l e t model t w i n f o r c e - b a l a n c e sting s y s t e m .

hird ramp bleed flow exit Throat slot bleed/ Second ramp bleed chamber Cowl rotatlon point Figure 5 . Inlet b l e e d and b y p a s s s y s t e m for wind- tunnel model.

l o p vlew (upper cowl) Second and thlrd Throat slot bleed/ ramp bleeds bypass exit O D

Rotatlng --+- Flxed

Slde vbw Throat slot bleed/

f bvDass exit

Strtlc pressure Total pressure probe Port (typical) Ramps (bottom vlrw) Lower cowl ( a ) Inlet p r e s s u r e s .

Figure 6 . S t a t i c pressure ports on the 7.5-percent wind-tunnel model B : : % - ! Y g F - \ Forward fuselage, right side Aircraft

waterline- < - - -

Wing fairing A J n g fairing Boundry-layer diverter

c Boundry-layerdiverter

Wing fairing and boundry-layerdiverter, right side Right upper fuselage / ; . .

I . .-. , / Aircraft Q Static pressure port (typical)7 Aircraft Q Right lower fuselage

r Static pressure port (typical)

--- Aircraft Q -/--- Right upper aft fuselage Right lower aft fuselage ( b ) Fuselage, wing f a i r i n g , and boundary-layer d i v e r t e r p r e s s u r e s .

Figure 6 . C o x 1 uded.

Line of Metric rotation bryak I I I Second Third ramp ramp bleed _I_ bleed Wind tunnel model I ! ! bleed/ / I I ' . I I _ 1 .

I

. I*

.- \ I , I .

bleed I L I I I \-VariabletFtss door Comparison of a i r p l a n e and Figure 7 .

wind-tunnel model b l e e d and bypass s y s t e m s o n the upper c o w l .

Bypass 2185.7 crn (860.5 In) (XIL = 1) I (780.5 In) F.S.= 1052.8 cm (414.5 in) (XIL = 0.892) (XIL = 0.400) Lower cowl lip (side view) Left side of aircraft (top view) Figure 8 .

Major reference p o i n t s f o r F-15 a i r c r a f t and Wind- tunnel model.

mo _f_ Flight EZZZjlG-Foot Transonic Wind Tunnel

- Military /

1.6 1.4 Total 60 engine 120 Total 1.2 airflow, engine airflow, kglsec 40 80 lblsec M, 1.0 .8 . 6

-

.4

'

n

. . 40 80 120 160 200 240 280 320X1O6

.5 1 .o 1.5 2.0 Reynolds number based on length of model or aircraft Mcn F i g u r e 9 . T o t a l engine a i r f l o w F i g u r e 10. F l i g h t and w i r r i - f o r t h e F-100-PW-100 a f t e r b u r n i n g tunnel Reynolds numbers.

t u r b o f a n engine a t an a l t i t u d e of 9 . 1 km (30,000 f t ) .

Bleed airflows Total inlet flow-

1 1

Engine a i r f l o w 4

' Aspilled

A . = 1 (Bleed airflows) + Engine airflow

( a ) Diagram showing sum of i n l e t captured stream t u b e a r e a ( A o ) and geometric c a p t u r e area ( A ) .

F i g u r e 1 1 . General teras for p r e s s u r e - i n t e g r a t e d i n l e t d r a g s a n i i n l e t c a p t u r e r a t i o s .

Upper cowl drag& \

-----_____

Lower cowl drag

Inlet drag = additive drag + upper cowl drag + lower cowl drag + sideplate drag

where additive drag = drag on three ramps + stream thrust in inlet throat - free-stream stream thrust R e l a t i o n s h i p s used i n t h e c a l c u l a t i o n of p r e s s u r e - ( b ) i n t e g r a t e d inlet d r a g .

Aircraft waterline 7

-_-------_

L-- FcI Y Details of inlet drag and l i f t e q u a t i o n s .

( c ) Figure 1 1 . Continued.

LLadd

Dadd = Flr sin (I +A'1) + F2r sin (I + A > ) + FQr sin a + A '

( ( ( 3)

+ FI COS (a +At3) - Fo

Ladd = Fir COS (I +A';) + Fpr COS u +A>) + Fgr COS u + 6'3

( ( 0

- FI sin (I + A '

( 3)

Additive drag ard lift relationships for (d) the control volume shown in figure ll(c).

Figure 11. Concluded.

q , ib/in2 0 2 4 6 8 u = P = Oo, h, = 6.1 km (20,000 11) to 9.1 km (30,000 ft) I I I 1

=-05k \ 2 1 0 x 1 0 - ~

M , = 0.6, h , = 6.1 km (20,OOO ft) f .04

r

CDIP I ~ - ..

and 2 5 CLIP f .03 = 6.1 km (20,OOO 11) cP n - .6 .8 1.0 1.2 1.4 1.6 f .02 M, = 1.5, h, Mm = 9.1 km ~30,oOo 1 1 ) (b) Uncertainties in flight f .01

I Wind tunnel?

CDIP and CLIP.

-

0 1 2 3 4 5 q , , Nlcm2 (a) Uncertainties in Cp.

Figure 12. Uncertainties in Cp, CDIP, ard CLIP f o r various Mach numbers and altitudes.

XIL 0 .2 .4 .6 .8 1 .o I I I I I 1 .2 - .2 Engine power

- .4

0 setting

i

C 0 Military- Plf 0 80percent

- .6

0 Idle Open symbols denote

- .8

wind-tunnel data: closed symbols denote -1.0 flight data.

I ( a ) Lower s u r f a c e .

XIL .2 .4 .6 .8 1.0 I I 1 I 1 I 0 Milltaw C put 0 Idle - .8 break flight data.

-1.0 .9 .2 . 3 .4 .S .6 .7 .8 XIL ( b ) Upper s u r f a c e .

Figure 1 3 . Effects of mass f l o w on upper and lower s u r f a c e p r e s s u r e s a t M , = 0.98 a = 3'8 and p = 0 ' .

X I L 0 .2 .4

A

.4 deg

3 "

P l f -2 0 -4 0 7

Q'

C - .4 -:I

Open symbols denote wind-tunnel data; - Metric closed symbols denote -.6 break flight data.

I I ( a ) Lower s u r f a c e .

- -.2 p ,

- .4

C deg 0 - 4 0 7 Open symbols denote

- .8

wind-tunnel data; closed symbols denote puf -1.0 --.! 0 flight data.

( b ) Upper s u r f a c e .

Figure 1 4 . E f f e c t of variable cowl on p r e s - s u r e s over t h e upper and l o w e r s u r f a c e s of the a i r c r a f t a t M, = 0 . 9 , a = 3', and m i l i - t a r y power.

.2

- .2

Open symbols denote wind-tunnel data; Metric closed symbols denote

-'6 - .8 1 I y ( break, flight I data. I

I -1.0 .9 .3 .4 .5 .6 .7 .8 XIL ( a ) Lower s u r f a c e .

.2 - .2 - .4 C Puf 0 0 - .6 0 5 Open symbols denote wind-tunnel data; closed symbols denote

-1.0 -t 0

flight data.

I I I I

( b ) U p p e r s u r f a c e .

Figure 1 5 . V a r i a t i o n of s u r f a c e pressure c o e f f i c i e n t s w i t h X / L f o r t w o angle of a t t a c k a t M , , , = 0 . 9 , p = O o , ard m i l i t a r y power.

0 Pressure 0 Pressure integrated Wind integrated Wind 0 Force tunnel tunnel 0 Force

balance 1

balance 0 Pressure integrated, 0 Pressure integrated, flight flight

- Flight

--Theory (ref. 3) - -Theory (ref. 3) .004 .006 .004 .002 - b CDlP CDlP 0 .002 - .002 .010

r

CLIP *m5 0

AoIA ( a ) M , = 0.6, a = O O .

Figure 16.

V a r i a t i o n of CDIP and C L I P w i t h A*/A for p = O o , A ' 3 = 1 1 . 2 ' , and 8 = O O .

0 Pressure 0 Pressure integrated Wind integrated Wind tunnel 0 Force tunnel 0 Force

balance 1

balance 1

0 Pressure integrated, 0 Pressure integrated, flight flight --Theory (ref. 3) .012 .008 .010 .006 CDlP .008 CDlP .004 \\ .006 .002 .004

I o

CLIP .010 c L ~ p * 0 2 0 ~ 0 ;. ~ , .005 .01 5.2 .4 .6 .8 .2 .4 .6 .8 AoIA AoIA (c) M, = 1 . 2 , a = O o . ( d ) M, = 1 . 5 , = 5'.

F i g u r e 1 6 . Concluded.

- Pressure integrated

---- 0 - - - - Force balance

- Pressure integrated

- - + - - Pressure integrated, flight

--- C, - - - Force balance

--e - - Pressure integrated, flight

.005 .006 .004 .005 .003 CDIP, .004 CDIP, counts .002 counts .003 .001 .002 .001

.012 L

- .010 - .0°8 CLIP, counts .006 - CLIP, -014 - counts .012 .004 - .002

.010 "f

1 I I

( b J a = 5 O .

F i g u r e 17.

V a r i a t i o n o f CDIP ani C L I P w i t h M , f o r t w o angles o f a t t a c k a t p = O O .

Pressure integrated 1 Wind

---- o---- tunnel Force balance

-- +-- -- Pressure integrated, flight

.007 = 1.5 .006 .005 . a M CDIP, counts

.003 -

: 0.6 .002 - .001 - 0-2 0 2 4 6 8 10 - 2 0 2 4 6 8 10 ( a ) C D I P .

F i g u r e 18. E f f e c t o f a on CDIP ancl C L I P ; p = 00.

. - Pressure integrated) Wind tunnel ---- 0- - - - Force balance

--+-- - Pressure integrated, flight

.020 I = 0.6 .018

P

.016 .002 .014 .012 .010 -.001 counts - .006 .ma - .005

-

.004 CLIP, .006 counts

-

.003 - .002 .004

I I

.001 .002 E f f e c t of p on CDIP and F i g u r e 1 9 .

I 1 I I 0 CLIP a t M, = 0.6 and a = 0 ' .

0 2 4 6 8 10 - 2 deg ( b ) C L I P .

F i g u r e 18. C o n c l u d e d .

220 { 7.5 -percent scale

0---7.5-percent scale force

2% I+--- 7.5-percent scale

balance data IT 220 {*Flight test 0--- 7.5-percent scale force balance data

.014 r

2% {+Flight test

/=

/

.012 t

9.1 (30,OOo) / /

.w4 ’ 9.1 (30,Ooo)

.010 .003 .008 9.1 (30,Ooo) CLIP CDlP .002 .006 Altitude,

t krn (ft)

.001 .004 .002 -.001 .4 .6 .8 1.0 1.2 1.4 1.6 A .6 .0 1.0 1.2 1.4 1.6 Mm Mm V a r i a t i o n of inlet drag w i t h ( a ) V a r i a t i o n of inlet l i f t w i t h (b) Mach number.

Mach number.

Figure 2 0 . V a r i a t i o n o f inlet drag and l i f t u n c e r t a i n t i e s w i t h Mach number for a and p = O O .

I -

f 20 { +---- 7.5-percent scale

p-- 7.5-percent scale force balance data

f 20 { &- Flight test

f 20 { f---- 7.5-percent scale

.020 [I-- 7.5-percent scale force balance data .018

* 20 1 e Flight test

.016

"r I

. 0 ° 3 t I .014

6.1 (20,000) .012 CDlP 6.1 (20,000) CLIP .010 '.

L

- .W1 '

- 2 0 2 4 6 8 .W6 0, deg .004 ( a ) A n g l e of a t t a c k a s a f u n c - t i o n of i n l e t drag c o e f f i c i e n t ; M, = 0 . 6 , p = 0 ' .

.W2 2 4 6 8 - 2 0 (b) A n g l e of a t t a c k a s a f u n c - t i o n of inlet l i f t c o e f f i c i e n t ; M, = 0 . 6 , p = 0 ' .

F i g u r e 21. E f f e c t of angle of a t t a c k on inlet l i f t and drag coefficients and t h e u n c e r t a i n t y i n t h e d a t a .

2. Government Accasion No. 3. Recipient's Qtalog No.

1. Report No.

NASA TP-2314 6. Report Date 4. Title and Subtitla November 1984 F l i g h t a n d Wind-Tunnel Comparisons of t h e I n l e t / A i r f r a m e 6. Performing Orgonlzetlon Code I n t e r a c t i o n of t h e F-15 A i r p l a n e 7. Author(s) 8. Performing Organization Report No, H-1175 L a n n i e D. Webb, Don Andriyich-Varda, a n d Stephen A. Whitmore 10. Work Unit No.

9. Performing Organization Name and Addreu NASA A m e s R e s e a r c h 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.0. BOX 273 Edwards, C A 93523 13. Type of Report and Period Covered 12. Sponsoring Agency Name end Address T e c h n i c a l P a p e r 14. Sponsoring Agency Code N a t i o n a l A e r o n a u t i c s a n d Space A d m i n i s t r a t i o n Washington, D.C. 20546 RTOP 51 4-54-04 16. Abstract The d e s i g n of i n l e t s a n d n o z z l e s and their i n t e r a c t i o n s w i t h the a i r p l a n e may a c c o u n t for a l a r g e p e r c e n t a g e of t h e total d r a g of modern high-performance a i r c r a f t . This paper d e s c r i b e s t h e 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 program a n d t h e f l i g h t tests c o n d u c t e d a t t h e Dryden F l i g h t R e s e a r c h F a c i l i t y o f the NASA Amee R e s e a r c h C e n t e r . I n l e t - d r a g a n d lift data from a 7.5-percent-scale wind-tunnel model are compared w i t h d a t a from a n F-15 a i r p l a n e w i t h i n s t r u m e n t a t i o n to match the model. P r e s s u r e c o e f f i c i e n t v a r i a t i o n s w i t h v a r i a b l e cowl a n g l e s , c a p t u r e r a t i o s , a n d a n g l e s of a t t a c k are examples o f f l o w i n t e r a c t i o n s p r e s e n t e d . Data are p r e s e n t e d f o r Uach numbers of 0 . 6 , 0.9, 1.2, a n d 1.5.

17. Key Words (Suggasted by Author(s1) 18. Distribution Statement

U n c l a s s i f i e d - U n l i m i t e d

I n l e t d r a g a n d l i f t P r e s s u r e i n t e g r a t i o n F o r c e b a l a n c e 22. Rice' 20. Security Classif. (of this page) 19. Security Classif. (of this report) 21. NO. of Pages U n c l a s s i f i e d U n c l a s e i f i e d 33 A 0 3