Document
ANALYTIC DEVELOPMENT OF AN IMPROVED SUPERSONIC CRUISE AIRCRAFT BASED ON WIND TUNNEL DATA* R. L. Roensch and G. S. Page Douglas A i r c r a f t Company McDonnell Douglas Corporation SUMMARY Data obtained from t h e MDC/NASA c o o p e r a t i v e wind t u n n e l program w e r e used t o develop e m p i r i c a l c o r r e c t i o n s t o theory. These methods were t h e n used t o develop a 2.2M Supersonic C r u i s e A i r c r a f t C o n f i g u r a t i o n w i t h a c r u i s e trimmed L/D of 10.2. The e m p i r i c a l c o r r e c t i o n s t o t h e t h e o r y are reviewed, and maximum t h e c o n f i g u r a t i o n a l t e r n a t i v e s examined i n t h e development of t h e c o n f i g u r a t i o n a r e presented. The b e n e f i t s of designing f o r optimum trimmed performance, i n c l u d i n g t h e e f f e c t s of t h e n a c e l l e s , are discussed.
INTRODUCTION A c o o p e r a t i v e MDC/NASA wind t u n n e l test program f o r an MDC designed supersonic c r u i s e a i r c r a f t c o n f i g u r a t i o n w a s conducted i n 1975. T e s t i n g w a s conducted i n t h e NASA Ames Research Center 9- by 7-foot s u p e r s o n i c wind t u n n e l a t Mach numbers from 1 . 6 t o 2.4, and i n t h e Ames 11- by 11-foot t r a n s o n i c wind t u n n e l a t Mach numbers from 0.5 t o 1 . 3 . A complete descrip- t i o n of t h e t e s t i s p r e s e n t e d i n r e f e r e n c e 1.
The c o n f i g u r a t i o n f o r t h e MDC/NASA tests w a s t h e McDonnell Douglas D3230-2.2-5E advanced s u p e r s o n i c t r a n s p o r t c o n f i g u r a t i o n shown i n f i g u r e s l ( a ) and l ( b ) . The c o n f i g u r a t i o n employs a modified arrow wing w i t h 71-degrees leading-edge sweep inboard and 57 degrees leading-edge sweep outboard. The design c r u i s e p o i n t i s 2.2M.
SUMMARY OF PREVIOUS W I N D TUNNEL TEST The d a t a from t h e 9- by 7-fOOt t u n n e 1 , s h o w n i n f i g u r e s 2 , 3, and 4 , w e r e presented a t t h e 1976 SCAR Conference ( r e f e r e n c e 2 ) . The estimates shown were based on Woodward l i f t i n g s u r f a c e theory ( r e f e r e n c e 3 ) , combined w i t h wave drag from a s u p e r s o n i c area r u l e theory ( r e f e r e n c e 4 ) , and s k i n f r i c t i o n d r a g estimates. E x c e l l e n t agreement i s shown between t h e estimated and experimental minimum d r a g i n f i g u r e 2 f o r a l l Mach numbers. The esti- mated and experimental d r a g p o l a r shapes d i f f e r , causing t h e wing body drag-due- t o - l i f t t o be o v e r p r e d i c t e d below 2.OM, underpredicted above 2.OM and t o a g r e e at 2.OM. Agreement i n l i f t curve s l o p e s , as shown i n f i g u r e 3 , is
- - - - - _
*This work w a s performed under NASA Contract NAS1-14621 excellent at the lower Mach numbers, but the agreement decreases at the higher Mach numbers. The estimated and experimental pitching moments shown in figure 4 agree well considering the difficulty of predicting pitch- ing moments for cambered, three-dimensional configurations. This character- istic of Woodward-calculated pitching moments is observed for other slender configurations.
The results of the MDC/NASA test justified the basic design and analysis of the MDC supersonic transport configuration. Although some discrepancy exists in the drag-due-to-lift, the overall data agreement was excellent and the test served as a good base for the methods and configuration development detailed in this paper.
DEVELOPMENT OF IMPROVED ANALYSIS METHODS DRAG-DUE-TO-LIFT When compared to the wind tunnel data, the basic Woodward theory underpre- dicts the drag-due-to-lift at Mach numbers greater than 2.0 as seen in figure 2. The comparison of data to theory also shows that the theory does not accurately predict the lift-curve slope at Mach numbers greater than 2.0 as seen in figure 3. The discrepancy in lift curve slope is also seen to increase with increasing Mach number. A correction to the Woodward-theory drag was developed based on the error in predicted lift curve slope and assuming no leading-edge suction. From the discrepancy in estimated and experimental lift curve slopes, a difference in angle-of-attack at constant CL can be calculated. The change in angle-of-attack, Aa, is calculated by equation 1.
EXP . THEORY
The supersonic flat plate (no leading-edge suction) drag term based on the angle shift, from equation 2, is then applied to the Woodward drag estimates 5.
as shown in figure ACD = CL2(e) Analysis of three additional wing planforms for which experimental data were available (references 5 and 6 ) showed similar trends in lift-curve-slope and drag estimates. A generalized correction factor, Acx/CL (equation ( l ) ) , was determined and the results are shown in figure 6 . The correction term is a function of the Mach number normal to a nominal leading-edge sweep, AED, which was chosen to represent a multi-segment leading edge by a single leading-edge sweep value. This correction to the Woodward drag estimates, the transonic leading edge (TLE) correction, shows excellent agreement with the experimental data as shown in figure 7.
NACELLE-WING INTEGRATION The Woodward program d i d n o t a c c u r a t e l y p r e d i c t t h e changes i n drag-due-to- l i f t and p i t c h i n g moment due t o n a c e l l e a d d i t i o n . The problem w a s i n t h e i n a b i l i t y of t h e Woodward program t o model t h e f l o w d i v e r t e r (pylon) and t h e i n t e r a c t i o n between the n a c e l l e - s h o c k and t h e wing-boundary-layer. A s a r e s u l t , t h e Woodward program d i d n o t a c c u r a t e l y p r e d i c t t h e nacelle-on-wing i n t e r f e r e n c e p r e s s u r e s . The measured nacelle-on-wing i n t e r f e r e n c e p r e s s u r e s are i l l u s t r a t e d i n f i g u r e 8. To c o r r e c t t h e Woodward a n a l y s i s , t h e program w a s modified t o a l l o w t h e i n c l u s i o n of t h e e x p e r i m e n t a l i n t e r f e r e n c e p r e s s u r e s on t h e wing. The a b i l i t y o f t h e program t o p r e d i c t p i t c h i n g moments and induced d r a g w a s s i g n i f i c a n t l y improved, as s e e n i n f i g u r e s 9 and 10.
DEVELOPMENT OF AN IMPROVED PERFORMANCE W I N G WING PLANFORM STUDY A wing planform s t u d y w a s conducted u s i n g t h e improved methods developed above. The a n a l y s i s o f c a n d i d a t e planforms w a s conducted under t h e f o l l o w i n g c o n s t r a i n t s : (1) Constant d i n g Area ( 2 ) Constant Aspect R a t i o (3) Constant Tip Chord ( 4 ) Constant t / c D i s t r i b u t i o n (5) Constant Design CL (6) Nacelle Induced Drag Not I n c l u d e d (7) 0 d e g r e e t r a i l i n g - e d g e sweep i n b o a r d of 31% semi-span The wing camber s u r f a c e w a s d e s i g n e d u s i n g t h e Woodward program o p t i m i z a t i o n o f a n i s o l a t e d wing. The wing w a s t h e n i n t e g r a t e d t o t h e f u s e l a g e by modi- f y i n g t h e r o o t a i r f o i l i n c i d e n c e . A f o u r d e g r e e r o o t i n c i d e n c e w a s used f o r a l l cases. The wing-body combination w a s a n a l y z e d f o r l i f t i n g e f f e c t s u s i n g t h e Woodward program and i n c o r p o r a t i n g t h e TLE , c o r r e c t i o n d e r i v e d above. Each c o n f i g u r a t i o n w a s o p t i m i z e d f o r minimum zero-lift-wave-drag u s i n g t h e A r b i t r a r y Body program ( r e f e r e n c e 4 ) . The c o n f i g u r a t i o n s w e r e trimmed at t h e c.g. l o c a t i o n f o r maximum trimmed L/D.
The planform s t u d y i n c l u d e d v a r i a t i o n s i n g e o m e t r i c p l a n f o r m and wing camber.
The geometry o f t h e planforms is shown i n t a b l e 1. Although wings W38 and W40 had good L / D ' s , as s e e n i n t a b l e 2 , t h e y w e r e dropped from t h e a n a l y s i s because of e x c e s s i v e wing l e n g t h which r e s u l t e d i n t h e wing o v e r l a p p i n g t h e h o r i z o n t a l t a i l . Wings W36 and W37 w e r e n o t r e t a i n e d f o r t h e f u l l a n a l y s i s due t o t h e i r l o w L/D v a l u e s . The d a t a i n t a b l e 2 p r e s e n t s t h e L / D v a l u e s f o r several s t e p s i n t h e a n a l y s i s p r o c e s s t o show t h e t r a d e s f o r v a r i o u s wings.
The g r o s s wing L / D v a l u e i s o b t a i n e d from t h e wing-alone induced d r a g d a t a , as produced by t h e o p t i m i z e d wing camber. A r e p r e s e n t a t i v e l i f t - i n d e p e n d e n t as p r e v i o u s l y e s t i m a t e d f o r t h e b a s e l i n e a i r c r a f t , i s added t o a d j u s t d r a g , t h e d a t a t o t h e p r o p e r L/D r a n g e f o r c o r r e l a t i o n w i t h t h e complete a i r c r a f t performance d a t a . The wing-body induced d r a g d a t a i n c l u d e t h e e f f e c t s o f r o t a t i n g t h e wing-root i n c i d e n c e t o f o u r d e g r e e s and a d d i n g t h e f u s e l a g e .
The r e p r e s e n t a t i v e l i f t - i n d e p e n d e n t d r a g used above i s r e t a i n e d . The wing- t h e e f f e c t of t r i m d r a g on t h e wing-body trimmed L/D i n c o r p o r a t e s body, d a t a w i t h t h e c.g. l o c a t e d t o achieve t h e maximum L/D, w h i l e m a i n t a i n i n g t h e r e f e r e n c e l i f t - i n d e p e n d e n t drag. A t t h e optimum c.g. l o c a t i o n , t h e t a i l l o a d i s up, so t h e trimmed L/D i s g r e a t e r than t h e wing-body L / D (CD, o f t h e t a i l is included i n t h e r e f e r e n c e l i f t - i n d e p e n d e n t d r a g ) . The complete a i r c r a f t L / D c o r r e c t s t h e wing-body trimmed L / D f o r t h e d i f f e r e n c e s i n the s k i n f r i c t i o n and z e r o l i f t wave d r a g of t h e a c t u a l a i r c r a f t c o n f i g u r a t i o n .
The planform s t u d y , u s i n g t h e complete c o n f i g u r a t i o n , showed a r e l a t i o n of t h e wing both drag-due-to-lift and c o n f i g u r a t i o n wave-drag-due-to-volume t o trailing-edge-sweep (notch r a t i o ) , w i t h t h e wave d r a g bounding t h e optimiza- t i o n p r o c e s s . When t h e t r a i l i n g edge sweep a n g l e a p p r o a c h e s t h e Mach angle, t h e wing area d i s t r i b u t i o n , c a l c u l a t e d by t h e Mach c u t t i n g p l a n e s , experiences r a p i d changes i n c r o s s - s e c t i o n a l area. A s a r e s u l t , t h e c o n f i g u r a t i o n wave drag-due-to-volume i n c r e a s e s a t h i g h t r a i l i n g - e d g e sweep a n g l e s , c a n c e l i n g t h e drag-due-to-lift b e n e f i t s a s s o c i a t e d w i t h h i g h t r a i l i n g - e d g e sweeps ( o r l a r g e notch r a t i o s ) . This produces an "optimum" t r a i l i n g - e d g e sweep a t This approximately one-half of t h e Mach cone a n g l e as s e e n i n f i g u r e 11.
e f f e c t made t h e h i g h t r a i l i n g edge sweep of wing W33 and W39 less b e n e f i c i a l than t h e g r o s s wing d a t a of t a b l e 2 i n d i c a t e d , showing t h e importance of a n a l y z i n g t h e complete a i r c r a f t when s e l e c t i n g t h e optimum wing planform.
The f o u r most promising wings from t h e planform s t u d y are shown i n f i g u r e 1 2 .
Based on t h e c r u i s e L/D and c o n s i d e r a t i o n of s t r u c t u r a l weight, t r a i l i n g edge f l a p s , and a i l e r o n e f f e c t i v e n e s s , wing W35 w a s chosen f o r f u r t h e r a n a l y s i s .
WING ASPECT RATIO STUDY An a s p e c t r a t i o s t u d y w a s conducted based on t h e wing W35 planform.
Three a l t e r n a t e methods f o r v a r y i n g t h e a s p e c t r a t i o were i n v e s t i g a t e d . They were: (1) c o n s t a n t t r a i l i n g - e d g e sweep o r notch r a t i o (inboard p a n e l L.E. sweep is allowed t o v a r y ) ; ( 2 ) c o n s t a n t leading-edge sweep (T.E. sweep i s allowed t o v a r y ) ; (3) c o n s t a n t leading- and t r a i l i n g - e d g e sweeps ( t i p chord i s allowed t o v a r y ) . The geometry of t h e s t u d y wings i s given i n t a b l e 3. The r e s u l t a n t L / D ' s f o r each approach, summarized i n f i g u r e 1 3 , are p r e s e n t e d below f o r each t y p e of planform c o n s t r a i n t .
(1) Trailing-Edge Sweep Constant: A s t r a i l i n g - e d g e sweep w a s t h e key parameter f o r d r a g as shown i n f i g u r e 11, a n a s p e c t r a t i o s t u d y w a s conducted a t c o n s t a n t t r a i l i n g - e d g e sweep.
AR COMMENTS 1 . 7 0 9.25 i n c r e a s e d induced d r a g 1.84 9.60 b a s e case 2.08 9.05 wave d r a g and induced d r a g p e n a l t y due t o decreased L.E. sweep.
(2) Leading-Eage Sweep Constant: To e v a l u a t e t h e p e n a l t y shown f o r t h e h i g h a s p e c t r a t i o wing w i t h f i x e d t r a i l i n g - e d g e sweep, t h e a n a l y s i s w a s r e p e a t e d f o r c o n s t a n t leading-edge sweep: COMMENTS AR base case 1.84 2.08 40 degrees trailing edge sweep may cause degraded flap and aileron authority, additional low speed analysis required ( 3 ) Leading-Edge Sweep Constant and Trailing-Edge Sweep Constant: Due to the strong impact of both leading- and trailing-edge sweeps in.theprevious analysis, a case was run holding all sweeps constant: AR L/DTRIMMED COMMENTS 1.61 9.27 increased induced drag 1.84 9.60 base case 2.09 9.47 wave drag penalty due to wing volume and induced drag penalty due to ' short tip chords.
The base case aspect ratio was near the optimum in all three studies, so the base aspect ratio of 1.84 was retained for the subsequent analyses.
WINGNACELLE INTEGRATION STUDY The classical approach to nacelle integration (reference 7 ) for supersonic aircraft is to reflex the wing trailing edge in the region of influence of the nacelle interference pressures as shown in figure 14. The reflex is designed to cancel the change in wing loading generated by the nacelle-on- wing interference pressure. This approach attempted to eliminate the change in drag-due-to-lift produced by the nacelle interference, but did not fully consider that there may be a benefit in the trimmed configuration performance due to the change in pitching moment produced by the nacelle installation.
Results of the 1975 MOC/NASA wind tunnel test (ref. 1) showed the reflex tested did not produce a favorable nacelle interference for the trimmed aircraft configuration. The loss in pitching moment with the nacelles installed created a signficicant loss in trimmed L/D for the design c.g. location. An improved wing-nacelle integration procedure was developed which includes the effect of the nacelle installation on the configuration pitching moment in addition to the effect on drag-due-to-lift.
The current procedure for wing-nacelle integration is based on the selection of the wing camber which will produce the maximum trimmed L/D for a specified c.g. location. The relation of maximum trimmed L/D to wing camber (referenced by the zero-lift pitching moment coefficient) and c.g. location is shown in figure 15. In figure 15, the maximum trimmed L/D attainable for a given c.g.
location is shown by the envelope curve created from the plots of trimmed L/D as a function of c.g. location for the individual pitch-constrained wings.
Each point on the envelope is a specific pitch-constrained wing. Therefore, f o r any d e s i g n c.g. l o c a t i o n a wing can be d e f i n e d which produces t h e maximum trimmed L/D.
The e f f e c t of n a c e l l e a d d i t i o n on a f i x e d geometry wing is shown i n f i g u r e 1 6 .
It i s seen t h a t i f t h e d e s i g n c.g. l o c a t i o n i s n e a r t h e c.g. l o c a t i o n f o r maximum trimmed L / D f o r a s p e c i f i e d wing geometry, a f a v o r a b l e n a c e l l e i n t e r f e r e n c e i s o b t a i n e d . I f t h e design c.g. is s u f f i c i e n t l y forward of t h e optimum c.g. l o c a t i o n , a n a c e l l e i n s t a l l a t i o n p e n a l t y may occur.
For c a s e s where t h e d e s i g n c.g. is forward of t h e optimum c . g . f o r t h e L/D envelope, shown i n f i g u r e 15, a l o c a l wing r e f l e x can b e added which w i l l r e s u l t i n a trimmed L/D g r e a t e r than t h a t f o r t h e non-reflexed wing. A s seen i n f i g u r e 1 7 , a g r e a t e r amount of r e f l e x is d e s i r e d as t h e c.g. l o c a t i o n i s moved f a r t h e r forward. The r e f l e x e s shown on f i g u r e 1 7 are simple geometric r e f l e x e s ( s e e i n s e t , f i g u r e 1 4 ) t h a t c a n c e l approximately 50 percent and 100 p e r c e n t of t h e n a c e l l e induced wing loading.
The combination of re-camber and/or r e f l e x r e s u l t s i n t h e maximum L / D envelopes shown i n f i g u r e 18. The amount of r e f l e x used f o r t h e r e f l e x e d wing envelope i n c r e a s e s as t h e c.g. moves forward u n t i l 100 p e r c e n t allevi- a t i o n of t h e n a c e l l e induced load i s achieved. Note t h a t i f t h e d e s i g n c.g.
l o c a t i o n i s n o t c o n s t r a i n e d t o b e forward of t h e c.g. l o c a t i o n f o r maximum L/D of t h e re-cambered wing envelope, then t h e r e is no i n c r e a s e i n L/D a v a i l a b l e f o r a r e f l e x e d and re-cambered wing. S i n c e f u e l pumping can b e used f o r c.g. c o n t r o l , t h e re-cambered wing without r e f l e x w a s s e l e c t e d f o r t h e a i r c r a f t . The r e s u l t a n t c.g. l o c a t i o n a t 37 p e r c e n t MAC is e q u i v a l e n t t o z e r o s t a t i c margin f o r t h e r i g i d wing.
HORIZONTAL TAIL OPTIMIZATION Since t h e MDC AST c o n f i g u r a t i o n u s e s a t a i l upload f o r t r i m t o o b t a i n a f a v o r a b l e t r i m d r a g , i t i s a p p r o p r i a t e t o c o n s i d e r o p t i m i z i n g t h e h o r i z o n t a l t a i l f o r i t s t r i m loading. The h o r i z o n t a l t a i l used i n t h e 1975 MDC/NASA test w a s f l a t (no camber o r t w i s t ) w i t h a biconvex a i r f o i l s e c t i o n and, as such,was n o t optimized f o r minimum drag-due-to-lift a t its t r i m CL. The experimental t a i l - o n d a t a are shown i n f i g u r e 19. The experimental t a i l d r a g p o l a r s (with c o e f f i c i e n t s based on wing a r e a ) f o r t h r e e a i r p l a n e a n g l e s of a t t a c k are shown i n f i g u r e 20. (The estimated p o l a r w a s c a l c u l a t e d f o r t h e uncambered t a i l without t h e wing induced f l o w f i e l d . ) A s shown, t h e e s t i m a t e d and experimental p o l a r shapes are i n good agreement. The CL f o r minimum d r a g , C L ~ , shows a s h i f t i n t h e experimental p o l a r r e l a t i v e t o t h e estimate. The s h i f t i n CL i s due t o t h e presence of a wing-induced f l o w f i e l d which c r e a t e d an adverse, non-uniform o n s e t flow a t t h e t a i l . The r e s u l t i n g n e g a t i v e CL of t h e experimental d a t a h a s an adverse e f f e c t on trimmed L/D.
The L/D p o t e n t i a l f o r an optimum t a i l w a s , a s s e s s e d by a n a l y s i s of a series of t a i l s w i t h v a r i e d CL An approximation v a l u e s i n a l i n e a r t r i m drag program.
of t h e camber d r a g expected f o r t h e t a i l w a s included. The a n a l y s i s showed a 0.2 improvement i n trimmed L / D f o r t h e optimum t a i l , as shown i p i f i g u r e 21.
An optimum t a i l has n o t been designed due t o t h e i n a b i l i t y of t h e Woodward program t o adequately analyze a t a i l i n t h e presence of t h e wing f l o w f i e l d .
21 0 CONCLUSION R e s u l t s o f t h e d e s i g n s t u d i e s described above, summarized i n f i g u r e 22, have been used t o develop a r e f i n e d AST c o n f i g u r a t i o n w i t h a n e s t i m a t e d L/D of 10.18.
The changes i n c o r p o r a t e d i n t h e r e f i n e d c o n f i g u r a t i o n are i l l u s t r a t e d i n f i g u r e 2 3 , along with t h e 1975 MDC/NASA test c o n f i g u r a t i o n . The r e f i n e d c o n f i g u r a t i o n is designated a s t h e model D3232-2.2-3 and is shown i n f i g u r e 24.
A c o o p e r a t i v e MDC/NASA wind t u n n e l test i s c u r r e n t l y being planned t o v e r i f y t h e performance estimated f o r t h e r e f i n e d c o n f i g u r a t i o n d e s c r i b e d above. The e x i s t i n g model f u s e l a g e and t a i l s w i l l b e r e t a i n e d , s o t h e e f f e c t s of f u s e l a g e shaping and t h e optimum t a i l design w i l l not be v e r i f i e d . The primary o b j e c t i v e s of t h e t e s t are: o Verify TLE c o r r e c t i o n o Confirm performance improvements f o r W35 o V a l i d a t e new n a c e l l e i n s t a l l a t i o n procedure o Obtain expanded nacelle-on-wing i n t e r f e r e n c e p r e s s u r e d a t a b a s e f o r u s e i n developing a n a l y t i c a l p r e d i c t i o n methods o Obtain expand,ed h o r i z o n t a l t a i l d r a g d a t a b a s e t o v a l i d a t e f u t u r e wing-body-tail a n a l y s i s and design methods The test is expected t o be conducted i n a NASA f a c i l i t y i n 1980.
SYMBOLS AND ABBREVIATIONS a a n g l e of a t t a c k A a
- c o r r e l a t i o n f a c t o r f o r t h e TLE c o r r e c t i o n
cL rl span f r a c t i o n A sweep a n g l e e q u i v a l e n t d e r i v e d sweep a n g l e AED l e a d i n g edge sweep a n g l e ALE t r a i l i n g edge sweep a n g l e 'TE 4 a n g u l a r change i n s l o p e of t h e wing camber s u r f a c e AR wing aspect r a t i o AST Advanced Supersonic Transport d r a g c o e f f i c i e n t cD C l i f t independent d r a g c o e f f i c i e n t DO l i f t c o e f f i c i e n t cL C l i f t curve s l o p e L a C l i f t c o e f f i c i e n t f o r minimum d r a g L O Cm p i t c h i n g moment c o e f f i c i e n t 21 1 zero l i f t p i t c h i n g moment c o e f f i c i e n t %I c.g. c e n t e r of g r a v i t y dz
-
wing camber s u r f a c e s l o p e i n t h e f r e e s t r e a m d i r e c t i o n dx h o r i z o n t a l t a i l i n c i d e n c e iH leading-edge L.E.
l i f t t o drag r a t i o L I D M Mach number f r e e s t r e a m Mach number MO MAC mean aerodynamic chord MDC McDonnell Douglas Corporation t h i c k n e s s t o chord r a t i o t l c t r a i l i n g - e d g e T.E.
TLE t r a n s o n i c l e a d i n g edge REFERENCES 1. R. L. Radkey, H . R. Welge, and J. E. F e l i x : Aerodynamic C h a r a c t e r i s t i c s of a Mach 2.2 Advanced Supersonic C r u i s e A i r c r a f t Configuration a t Mach . Numbers From 0.5 t o 2.4. N A S A CR-145094, 1977.
2. R. L. Roensch: Aerodynamic V a l i d a t i o n o f a SCAR Design. Proceedings o f t h e SCAR Conference - P a r t 2 , N A S A CP-001, 119771, pp. 155-168.
3. F. A. Woodward, E. N. Tinoco, and J. W. Larsen: Analysis and Design of Supersonic Wing-Body Combinations, I n c l u d i n g Flow P r o p e r t i e s i n t h e Near F i e l d . P a r t I - Theory and Application. N A S A CR-73106, 1967.
4. A. E. Gentry, D. N. Smyth, and W. R. Oliver: The Mark I V Supersonic- Hypersonic A r b i t r a r y Body Program. AFFDL-TR-73-159, 1973.
5. 0. A. Morris, D. E. F u l l e r , and C. B. Watson: Aerodynamic C h a r a c t e r i s t i c s of a Fixed Arrow-Wing Supersonic Cruise A i r c r a f t a t Mach Numbers of 2.30, 2.70, and 2.95. N A S A TM-78706, 1978.
6. R. B. S a v e l l e 111, and E. J. Landrum: T h e o r e t i c a l and Experimental Study of Twisted and Cambered Delta Wings Designed f o r a Mach Number of 3.5.'
N A S A TN D-8247, 1976.
7. R. J. Mack: A Numerical Method f o r Evaluation and U t i l i z a t i o n of Supersonic Nacelle-Wing I n t e r f e r e n c e . N A S A TN D-5057, 1969.
8. E. Bonner, M. H. Roe, R. M. Tyson, and R. Y. Mairs: I n f l u e n c e of Propulsion System S i z e , Shape, and Location on Supersonic A i r c r a f t Design. N A S A CR-132544, 1974.
21 2 TABLE 1.- WING PLANFORM GEONETRY SUIIMARY ~ TRAlLlf LEADING EDGE EDGE PLANFORM y BREAK A OUTBOARD REFERENCE A INBOARD y BREAK A OUTBOARD
(DEGREES) (% SEMISPAN) (DEGREES) (W SEMISPAN) (DEGREES)
NUMBER N/A 30 46 w33 71 NONE 71 57 30 17 w34* 63.6 71 61.5 30 3 1 w35 70 NONE 0 61 NONE N/A W36 30 18 w37 65 NONE N/A N/A 30 62 W38 74 NONE 30 43 w39 74 70 62 62 30 25 W40 74 55 TABLE 2.- W I N G PLANFORM PERFORMANCE SUMMARY DATA USING BASELINE" AIRCRAFT PLANFORM
s IN FRICTION AND AVE DRAG COMPLETE
REFERENCE GROSS WING WING/ BODY WING BODY, TRIMMED AIRCRAFT NUMBER L/ D L/ D L/ D L/ D w33 9.75 9.91 10.10 9.75 W34" 8.69 8.76 9.10 9.10 w35 9.09 9.25 9.64 9.60 W36 8.32 8.39 8.66 w37 8.61 W38 10.50** w39 9.64 9.60 9.80 9.75 W40 9.18*'k *BASELINE **DROPPED DUE TO STRUCTURAL LIMITATIONS 21 3 TABLE 1.- WING PLANFORMS FOR ASPECT RATIO STUDY PLANFORM LEADING EDGE TRAILING EDGE
A INBOARD I LEADING EDGE BREAK I AOUTBOARD A OUTBOARD REFERENCE ASPECT
NUMBER RATIO
(DEGREES) I (%SEMISPAN) 1 (DEGREES) I (DEGREES)
w35 1.84 71 70 61.5 W41 2.08 71 70 62 W42 2.08 67 70 62 30 w44 1.70 72 70 62 30 w45 2.09 71 65 61.5 31 w47 1.61 71 75 61.5 31 21 4 90.5 M (310 FT) - 4 I 16.7 M (54.8 FT) . ... . . .. . . .... , ., , . ...... .
t - (a) Conf i.guration details.
X X=84.166 (33.136) Z= 1.722 (0.678) - - - - = - . I - -.._
-- --
&---FRp-------- - -~ a - X (b) High-speed wind tunnel model details.
Figure 1 . - McDonnell Douglas D3230-2.2-5 configuration and model details.
21 5 + 38 37 36 35 34 1ST ENTRY 0’3-
U 2 1 2 213 211 209 210 ZNOENTRY *+\@
_ _ _ ESTIMATED - 0.2 - 0.1 0 - I I 1 1 I -0.1 Figure 2 . - Comparison of estimated and experimental drag polars for B1W2, Mach 1 . 6 to 2.4.
RUN NO.
SYM 1.6 1.8 2.0 2.2 2.4 MACH NO.
+ 38 37 36 35 34 lSTENTRY + 212 213 211 209 210 PNDENTRV -5 0 0 OI Figure 3 . - Comparison of estimated and experimental lift curves for B1W2, Mach 1 . 6 to 2.4.
21 6 0.20 0.1 5 0.10 C L 0.05 " 0,04 0.03 0.02 OD1 0.01 0.01 0.01 0.01 0 -0.01 cnl Figure 4 . - Experimental and estimated supersonic pitching moments for B1W2.
0.3 r 0 EXPERIMENT
- WOODWARD
--
WOODWARD + C L 2 (e)
0 - 0 0.08 0.16 a CL Figure 5 . - Derivation of the transonic leading edge correction; 2.2 M.
21 7 0 SCAT 15F-9898 h E D = 7 2 9 - 0 D-32302.2 5E A E ~ = 67.5' 0 . 0 8 V 680 DELTA AED = 680 Figure 6 . - Transonic leading edge (TLE) correction. (Semi-empirical correction of Woodward for improved drag prediction.)
0.24 0.18 0.12 C L 0.06 -0.06 0 0.01 0.02 0.03 0.04 C D Figure 7.- Effect of TLE correction on estimated drag polars for B1W2, Mach 2.0 to 2.4.
21 8 Figure 8 . - Pictorial representation of nacelle-on-wing interference pressures.
Figure 9 . - Comparison of Woodward with nacelle interference modifications and experimental pitching moments; tail off, 2.2 M.
21 9 0.02a 0 . 0 1 8 0 . 0 1 6 0.014 EXPERIMENT -WOODWARD 0 . 0 1 2 --MODIFIED WOODWARD 'D o.oia M,= 2.2 0.008 CONFIGURATION: W, B1 N1 0.006 rn DATA AT TUNNEL REYNOLDS NUMBERS 0.004 BOTH WOODWARD THEORIES INCLUDE THE TLE CORRECTION 0.002 0 . m - 0 0 . 0 2 0.04 0.060 . 0 8 0 . 1 0 . 1 2 0 . 1 4 0.16 C L F i g u r e 10.- Comparison of Woodward w i t h n a c e l l e i n t e r f e r e n c e m o d i f i c a t i o n s and e x p e r i m e n t a l d r a g p o l a r s ; 2.2 M.
0.010 M = 2.2 I 1/2 M A C H A N G L E AR = 1.84
8 2 M A C H A N G L E
CL = 0.1 WING A R E A = C O N S T A N l 0 10 20 30 0 4 0 50 60 ATE 57 62 7 1 ALE >70 % 77 F i g u r e 11.- E f f e c t of t r a i l i n g - e d g e sweep on induced drag and wave drag.
W35 W34 w33 W39 MAX TRIMMED L/D BASIC ANALYSIS 9.45 9.80 9.75 9.70 (NO N E CORRECTION) REFINED ANALYSIS 9.10 9.75 9.75 960 (WITH TLE CORRRECTION) F i g u r e 12.- Wing planform s t u d y , summary of s e l e c t e d wings; 2.2 M.
61.5' LID = 9.47 CONSTANT LEADING AND 4 = 2.09 TRAILING EDGE SWEEPS, WINGS W45. W47 CONSTANT LEADING EDGE SWEEP, WING W41 710 6 1 . 5 ' - 43 = 1.84 BASELINE, WING W35
I
CONSTANT TRAILING EDGE SWEEP, WINGS W42, W44 4 = 1.7 d = 2.08 F i g u r e 13.- L/D v a r i a t i o n s w i t h a s p e c t r a t i o ; 2.2 M.
2 21 DEFINITION OF REFLEX ANGLE
-. --. .- .- /
Figure 14.- Reflex i n r e g i o n of n a c e l l e i n t e r f e r e n c e .
L/D,, ENVELOPE
---- CM,= 0.02424
CM0 = 0.01394
-.- CMo = 0.01120 MAXIMUM
TRIMMED Cu0 = 0.00453 L/D NOTE 9.2
-
INCLUDES NACELLE \ WAVE DRAG-WE-TOVOLUME \ \ \ \ 9.0 \ ~ ~ ~ ~ 0 10 2 0 30 40 50 60 70 CG LOCATION (PERCENT MAC) Figure 15.- S e l e c t i o n of wing p i t c h i n g moment f o r optimum trimmed L/D; 2.2 M, n a c e l l e s o f f .
10.0 r CM0 = 0.01120 NOTE: (11 NACELLE WAVE DRAG-DUE-TO-VOLUME IS INCLUDED FOR NACELLE ON AND NACELLE OFF CASES NACELLE SKIN FRICTION DRAG IS (2) INCLUDED IN INSTALLED ENGINE PERFORMANCE
9.2 1 I I I I I
10 20 30 40 50 60 CG (PERCENT MAC) Figure 16.- E f f e c t of n a c e l l e a d d i t i o n on a p i t c h c o n s t r a i n e d wing; 2 . 2 M.
-
NO REFLEX Cmo=0.01120
r
--- PARTIAL REFLEX A dx dz = 0.01
I
MAXIMUM t
TRIMMED ~~ L/D .
9.6
I
NACELLE SKIN FRICTION DRAG INCLUDED IN INSTALLED ENGINE PERFORMANCE I I I I I 1 9.4 I I 0 1 0 20 30 40 50 60 70 CG LOCATION (PERCENT MAC) Figure 17.- E f f e c t of r e f l e x f o r n a c e l l e s on a p i t c h c o n s t r a i n e d wing; 2.2 M , n a c e l l e s on.
RECAMBER AND NACELLES ON REFLEX WING NOTE 0 NACELLE WAVE DRAG INCLUDED 0 NACELLE SKIN FRICTION DRAG INCLUDED IN INSTALLED ENGINE PERFORMANCE WING TRIMMED
L/D I /
I 1 I I I I I 2 0 30 40 50 60 70 CG LOCATION (PERCENT MAC) Figure 18.- Design LID envelopes f o r n a c e l l e a d d i t i o n w i t h wing r e f l e x and recamber; 2 . 2 M.
0 . 1 6 0 . 1 4 0.12 0 . 1 0 -30 -4- CL 0.08 oo 30 --o-- TAILOFF --A-- 0 . 0 6 0.04 0.02 0.006 0.008 0 . 0 1 0 0 . 0 1 2 0 . 0 1 4 0 . 0 1 6 0 . 0 1 8 OD20 0 . 0 2 2 0 . 0 2 4 C D Figure 1 9.- Experimental t a i l on and o f f drag p o l a r s ; 2 . 2 M.
I . . r . .
AC LTAIL OTAIL
---
ESTIMATED - 0.005 \ 0 u = 40 \ . \ -0.010 I 0.0010 0.0020 0.0030 0.0040 F i g u r e 20.- H o r i z o n t a l t a i l d r a g p o l a r s ; 2.2 M, c o e f f i c i e n t s based on wing area.
10.2 /-' NACELLES ON WITH / \OPTIMIZED TAIL 10.0 NACELLES ON (NO REFLEX)
f"
NOTE
MAXIMUM -
- TRIMMED 9.8 0 NACELLE WAVE DRAG INCLUDED IN ALL L'D CASES 0 NACELLESKIN - FRICTION DRAG 9.6 INCLUDED IN INSTALLED ENGINE NACELLES OFF PERFORMANCE 9.4 I I I I I I I 20 30 40 50 60 70 CG LOCATION (PERCENT MAC) F i g u r e 21.- E f f e c t of optimized t a i l on d e s i g n L/D e n v e l o p e s ; 2.2 P Z .
TLE CORRECTION USED TO IMPROVE WOODWARD ESTIMATES WING W35 SELECTED AS NEW PLANFORM MODIFIED WOODWARD PROGRAM ACCURATELY PREDICTS EFFECT OF NACELLES WING RECAMBER PRODUCES FAVORABLE NACELLE INTERFERENCE WING REFLEX NOT NEEDED IF CG CAN BE ALLOWED TO VARY HORIZONTAL TAIL SHOULD BE OPTIMIZED FOR ITS TRIM LIFT IMPROVED METHODS ARE REQUIRED TO PROPERLY DESIGN AN OPTIMIZED TAIL F i g u r e 22.- Conclusions.
1975MDC/NASA TEST CONFIGURATION CURRENT REFERENCE CONFIGURATION L/D = 9.09 L/D = 10.18 (BASED ON WIND TUNNEL DATA) (BASED ON METHODS WHICH MATCH WIND TUNNEL DATA) KEY ITEMS PLANFORM MODIFICATIONS DETAILED FUSELAGE SHAPING WING THICKNESS DISTRIBUTION DETAILED NACELLE INTEGRATION HORIZONTAL TAIL OPTIMIZATION
L
'i- 2.5% F i g u r e 23.- Refined aerodynamic c o n f i g u r a t i o n ; MDC/NASA test c o n f i g u r a t i o n compared w i t h c u r r e n t r e f e r e n c e c o n f i g u r a t i o n .
~~ F i g u r e 24.- Details of McDonnell Douglas D3232-2.2-3 c o n f i g u r a t i o n .