Document
NASA Technical Memorandum 85776
Low-Speed Wind-Tunnel Study of the
High-Angle-of-Attack Stability and
Control Characteristics of a Cranked-
Arrow-Wing Fighter Configuration
Sue B. Grafton
LungZey Reseurch Center Humpton, Virginiu National Aeronautics and Space Administration Scientific and Technical Information Branch SUMMARY A n i n v e s t i g a t i o n w a s conducted i n t h e Langley 30- by 60-Foot Tunnel t o study t h e low-speed, high-angle-of -attack s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s of a f i q h t e r confiquration incorporating a cranked arrow wing. The study w a s conducted as part of a NASA/General D y n a m i c s cooperative research-program t o i n v e s t i g a t e the a p p l i c a t i o n of advanced wing designs t o combat a i r c r a f t . T e s t s w e r e conducted on a b a s e l i n e configuration and on s e v e r a l modified configurations. The modifications included a wing apex notch, a wing trailinq-edqe extension, and wing fences. The r e s u l t s of t h e i n v e s t i g a t i o n showed t h a t t h e b a s e l i n e configuration e x h i b i t e d a high l e v e l of maxi- m u m l i f t b u t displayed undesirable l o n q i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s a t hiqh angles of a t t a c k . Various modifications w e r e made which improved the l o n q i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y characteristics of t h e configuration a t high angles of a t t a c k . However, m o s t of the modifications were d e t r i m e n t a l t o maximum l i f t .
INTRODUCTION A cooperative research proqram between t h e NASA Langley Research Center and General Dynamics to i n v e s t i g a t e the a p p l i c a t i o n of advanced wing designs t o combat a i r c r a f t w a s r e c e n t l y completed ( r e f . 1 From the r e s u l t s ' of t h i s e f f o r t , General Dynamics is developing an advanced F-I 6 d e r i v a t i v e , the F-I 6XL, which incorporates a highly s w e p t cranked arrow wing. A s p a r t of t h i s cooperative program, e x p l o r a t o r y low-speed, high-angle-of-attack static s t a b i l i t y and c o n t r o l tests w e r e conducted i n the Langley 30- by 60-Foot Tunnel on a cranked-arrow-wing f i q h t e r configuration t o i d e n t i f y problem areas and d e f i n e modifications f o r improved s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s . Many of the modifications i n v e s t i g a t e d i n the s u b j e c t study w e r e f i r s t developed i n the 1960's f o r use on supersonic t r a n s p o r t designs (ref. 2 ) .
Some of the basic low-speed s t a b i l i t y problems of highly s w e p t arrow wings include ( 1 ) pitch-up c h a r a c t e r i s t i c s a t moderate to high angles of a t t a c k brought on by vortex breakdown over the wing ( r e f s . 2 and 3 ) and ( 2 ) l a t e r a l - d i r e c t i o n a l i n s t a b i l i t y near maximum l i f t caused by asymmetric vortex breakdown under s i d e s l i p .
Modifications to the basic wing which are aimed a t e l i m i n a t i n g o r minimizinq t h e s e problems g e n e r a l l y d e s t r o y vortex l i f t and s i g n i f i c a n t l y reduce t h e low-speed maximum l i f t c o e f f i c i e n t of t h e winq ( r e f s . 4, 5, and 6 ) . Also, such modifications can have a detrimental impact on supersonic aerodynamic performance and would be unacceptable.
Thus, the challenge i n low-speed, high-angle-of-attack research is to provide altera- t i o n s to t h e b a s i c arrow-wing design to give acceptable low-speed s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s by using s e v e r a l options, which have to be evaluated supersonically.
The p r e s e n t i n v e s t i g a t i o n w a s conducted using a 0.15-scale modified F-16A model t o r e p r e s e n t an e a r l y F-16XL confiquration. The model used t h e fuselage and horizon- t a l t a i l of the F-16A and incorporated a highly swept cranked-arrow-wing design ( r e f . 1 ) . The b a s i c c o n f i g u r a t i o n w a s tested with modified leading and t r a i l i n g edges and with fences on the upper surface. A l l t h e configurations w e r e static f o r c e t e s t e d over wide ranges of angle of attack and angle of s i d e s l i p a t a free-stream v e l o c i t y of 75 f t / s e c , which corresponds t o a Reynolds number of 2.15 X IO6. A l i m i t e d number of flow-visualization tests w e r e made to provide information f o r i n t e r p r e t i n g t h e f o r c e - t e s t r e s u l t s .
SYMBOLS Static l o n g i t u d i n a l f o r c e s and moments are r e f e r r e d t o t h e wind-axis system, and s t a t i c l a t e r a l - d i r e c t i o n a l f o r c e s and moments are ref e r r e d t o t h e body-axis system.
A l l f o r c e - t e s t d a t a are r e f e r r e d t o a moment reference c e n t e r located l o n g i t u d i n a l l y a t 45 percent of the wing mean aerodynamic chord. All measurements w e r e reduced to standard c o e f f i c i e n t for& on t h e b a s i s of t h e geometric characteristics of the wing.
b wing span, f t
-
C mean aerodynamic chord, f t drag c o e f f i c i e n t , F ~ / & ~ s cD l i f t c o e f f i c i e n t , F ~ / ~ , , S CL rolling-moment c o e f f i c i e n t , MX/sSb pitching-moment c o e f f i c i e n t , My/qooSc ‘ m yawing-moment c o e f f i c i e n t , MZ/%Sb ‘ n side-f orce c o e f f i c i e n t , Fy&S CY drag f o r c e , l b F D l i f t f o r c e , lb FL s i d e f o r c e , l b F Y moment of i n e r t i a about X body a x i s , s l u q - f t 2 I X moment of i n e r t i a about Z body a x i s , s l u g - f t I Z r o l l i n g moment, f t - l b MX p i t c h i n g moment, f t - l b MY yawing moment, f t-lb MZ f ree-stream dynamic pressure, lb/f t2 % wing area, f t 2 S body reference axes X , Y , Z U angle of a t t a c k , deg anqle of s i d e s l i p , deg
B
Derivatives : ac
ac
acY ‘1 n
‘ n = w =ag
=aS
‘B B B
IZ C s i n a ‘ n
B I dyn = c nB cosa-q ‘B
MODEL The wind-tunnel model c o n s i s t e d of a 0.15-scale modified F-16A fuselaqe, which has a f l o w through duct, and a cranked-arrow-wing planform. A three-view sketch of t h e basic c o n f i q u r a t i o n is presented i n f i g u r e 1. The geometric c h a r a c t e r i s t i c s of t h e test model are l i s t e d i n t a b l e I. Sketches showing a wing leading-edge apex modification, a winq trailing-edge modification, and a wing fence modification are shown i n f i g u r e 2. The f o r c e - t e s t model, constructed p r i m a r i l y of molded f i b e r g l a s s and wood, w a s a 0.15-scale model of t h e proposed f u l l - s c a l e a i r p l a n e . The wings w e r e constructed without c o n t r o l surfaces.
TUNNEL AND APPARATUS The tests w e r e conducted i n t h e Langley 30- by 60-Foot Tunnel. The model support system used i n the f o r c e tests is shown i n f i g u r e 3. A s reported i n reference 3, c e r t a i n model support systems can cause l a r g e flow i n t e r f e r e n c e problems when used t o test c o n f i q u r a t i o n s which produce s t r o n g vortex flows a t high angles of a t t a c k . The model support system shown i n f i g u r e 3 w a s designed to minimize these i n t e r f e r e n c e problems. The d a t a w e r e measured a t a free-stream v e l o c i t y of 75 f t / s e c , which cor- responds t o a Reynolds number of 2.15 x I O 6 based on t h e mean aerodynamic chord of t h e wing. The f o r c e - t e s t s e t u p used a strut-mounted system and incorporated a s i x - component i n t e r n a l l y mounted strain-gage balance. The model engine i n l e t and e x i t w e r e open for a l l tests.
TESTS Static-force tests w e r e conducted over an angle-of-attack range from - 4 O t o 41 O a t s i d e s l i p angles from -13.5O to 10.5O. T e s t s w e r e also conducted to study t h e e f f e c t s of the wing leading- and trailing-edge modifications and wing fences described earlier.
I n a d d i t i o n to t h e s t a t i c tests, a limited number of smoke flow tests w e r e made t o h e l p i d e n t i f y flow behavior over t h e wing. The smoke f l o w tests were used t o document vortex f l o w behavior and to e s t a b l i s h a r e l a t i o n s h i p between vortex flow breakdown as a f u n c t i o n of angle of a t t a c k and angle of s i d e s l i p f o r use i n i n t e r - p r e t i n g the s t a t i c - f o r c e test data.
RESULTS AND DISCUSSION Baseline Configuration
.- The l o n g i t u d i n a l c h a r a c t e r i s t i c s of
t h e b a s e l i n e c o n f i g u r a t i o n with and without the v e r t i c a l t a i l are presented i n f i g - ure 4. The d a t a s h o w that t h e configuration e x h i b i t e d a high l e v e l of l i f t , achieved maximum l i f t near an anqle of a t t a c k of 36O, and e x h i b i t e d a high l e v e l of l o n g i t u d i - n a l i n s t a b i l i t y o r pitch-up a t angles of a t t a c k above about l o o . As expected, t h e t a i l had r e l a t i v e l y l i t t l e e f f e c t on t h e longi- d a t a show t h a t removinq t h e v e r t i c a l t u d i n a l aerodynamic c h a r a c t e r i s t i c s . The severe pitch-up tendency is c h a r a c t e r i s t i c of highly swept wings a t l o w speeds and is w e l l documented i n earlier p u b l i c a t i o n s ( f o r example, r e f s . 2 and 3). The pitch-up i n s t a b i l i t y is mainly a s s o c i a t e d with the formation of h i g h l y concentrated vortex flow emanating from t h e wing apex a t moderate t o high angles of a t t a c k . The vortex core tends t o l i f t off the s u r f a c e of the a f t p o r t i o n of t h e wing while remaining close t o the s u r f a c e of t h e forward portion. The loss i n vortex-induced l i f t over the a f t p o r t i o n of the wing tends t o s h i f t the aero- dynamic c e n t e r forward t o produce t h e p i t c h i n s t a b i l i t y . A t t h e higher anqles of a t t a c k , t h e problem is f u r t h e r aggravated by vortex breakdown, which starts a t the rear of t h e wing and moves forward with i n c r e a s i n g angle of attack. Figure 5 pre- s e n t s photographs of smoke flow s t u d i e s on the b a s e l i n e configuration a t high angles of a t t a c k . The photographs show t h a t vortex breakdown s t a r t e d a t t h e wing t r a i l i n g edge near a = 25O and progressed forward with i n c r e a s i n g angle of a t t a c k u n t i l it reached t h e wing apex near a = 35O. The photographs a l s o show t h a t s i d e s l i p p i n g t h e model increased the angle of a t t a c k f o r vortex breakdown on the leeward wing and reduced t h e angle of a t t a c k f o r vortex breakdown on the windward wing.
L a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s . - The v a r i a t i o n of the s t a t i c l a t e r a l - d i r e c t i o n a l f o r c e and moment c o e f f i c i e n t s with angle of s i d e s l i p f o r t h e b a s e l i n e configuration is presented f o r various values of angle of a t t a c k i n f i g - ure 6. The d a t a show t h a t t h e v a r i a t i o n s of Cn and C, w i t h s i d e s l i p w e r e f a i r l y l i n e a r a t angles of a t t a c k b e l o w 31° and extremely nonlinear a t angles of a t t a c k of 31 O and 36O. The d a t a of f i g u r e s 7 and 8 show t h a t t h e v e r t i c a l t a i l w a s i n e f f e c - t i v e by a = 31° and a c t u a l l y produced highly d e s t a b i l i z i n g yawing-moment and rolling-moment increments a t a = 41O. The d a t a of f i g u r e s 6 and 7 are summarized i n f i g u r e 8 i n terms of the l a t e r a l - d i r e c t i o n a l s t a b i l i t y d e r i v a t i v e s C YB’ ‘ne, and . The d e r i v a t i v e s w e r e obtained from t h e c o e f f i c i e n t d a t a by t h e slope method
c, B
using the d a t a p o i n t s f3 = 4.5O and f3 = -5.5O. The r e s u l t s show that the model with t h e v e r t i c a l tail removed w a s d i r e c t i o n a l l y unstable throughout the a range with the i n s t a b i l i t y i n c r e a s i n g s h a r p l y above a = 30°. The d i h e d r a l e f f e c t
(-“,s>
w a s s t a b l e up t o a = 30° b u t showed an abrupt d e s t a b i l i z i n g break near a = 32O.
The a d d i t i o n of the v e r t i c a l tail produced d i r e c t i o n a l s t a b i l i t y for angles of a t t a c k up t o about 27O. Above a = 21°, t h e increment i n d i r e c t i o n a l s t a b i l i t y provided by the v e r t i c a l tail decreased r a p i d l y , and t h e model became d i r e c t i o n a l l y unstable above a = 27O. The d e s t a b i l i z i n g e f f e c t of t h e v e r t i c a l t a i l above a = 31° i n d i - cates t h a t t h e t a i l w a s i n an adverse sidewash f i e l d . The v e r t i c a l t a i l increased t h e stable d i h e d r a l e f f e c t a t l o w and moderate angles of a t t a c k b u t became l a t e r a l l y d e s t a b i l i z i n g a t hiqher angles of attack. This d e s t a b i l i z i n g e f f e c t r e s u l t e d i n t h e b a s e l i n e configuration e x h i b i t i n g lateral i n s t a b i l i t y i n t h e angle-of-attack range from 33O to 40°.
The lateral i n s t a b i l i t y a t high angles of a t t a c k is a t t r i b u t e d mainly to t h e asymmetrical breakdown of t h e wing leading-edge vortices under s i d e s l i p . The e f f e c t of s i d e s l i p angle on vortex breakdown, shown i n the photographs of f i g u r e 5, is d i s - cussed i n more d e t a i l i n r e f e r e n c e 2. Other i n v e s t i q a t i o n s (see r e f s . 3 and 4 ) have documented the predominant e f f e c t of the asymmetric vortex breakdown on ing- moment c h a r a c t e r i s t i c s a t high angles of a t t a c k f o r c o n f i g u r a t i o n s with l y s w e p t wings. This loss i n lateral s t a b i l i t y combined with the h i g h l y unstable d i r e c t i o n a l - s t a b i l i t y c h a r a c t e r i s t i c s made t h e configuration highly s u s c e p t i b l e t o yaw diver- gence, as shown by the data presented i n f i g u r e 9. Negative values of i n d i c a t e s u s c e p t i b i l i t y to d i r e c t i o n a l divergence, as described i n r e f e r - ence 5. The b a s e l i n e c o n f i g u r a t i o n is seen to have unstable or negative values of C above (x = 32O. These negative values of are caused by t h e
"8 ,dyn c"$ ,dyn
. To correct t h i s
and n e u t r a l or unstable values of unstable values of
c1 8 cn$
l a t e r a l - d i r e c t i o n a l s t a b i l i t y problem and t h e severe pitch-up discussed earlier, a number of configuration modifications w e r e s t u d i e d , and t h e r e s u l t s are discussed i n the subsequent s e c t i o n .
Modified Configurations Longitudinal aerodynamic c h a r a c t e r i s t i c s . - In an attempt t o a l l e v i a t e t h e p i t c h and l a t e r a l - d i r e c t i o n a l i n s t a b i l i t i e s discussed above, the v e r t i c a l t a i l w a s removed from the b a s e l i n e configuration, and various a l t e r a t i o n s w e r e made i n order t o develop a modified configuration. The modifications included a wing apex notch, a wing trailinq-edge extension, and wing fences ( f i g . 2 ) . Because the i n s t a b i l i t i e s w e r e a d i r e c t r e s u l t of the wing leading-edge vortex flow c h a r a c t e r i s t i c s , each modi- f i c a t i o n w a s designed t o favorably alter these c h a r a c t e r i s t i c s .
Presented i n f i g u r e 10 are t h e r e s u l t s of tests t o study the e f f e c t of the wing apex notch modification on t h e l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s of the con- f i g u r a t i o n . This modification, shown i n f i g u r e 2 ( a ) , produced a l a r g e s t a b i l i z i n g e f f e c t and s i g n i f i c a n t l y reduced the i n s t a b i l i t y above (11 = 1 1 O . The modification caused some reduction i n l i f t c o e f f i c i e n t above (x = 15O. This reduction i n l i f t c o e f f i c i e n t w a s apparently a r e s u l t of the apex notch, which weakened the wing leading-edge vortex.
Another modification made s p e c i f i c a l l y to minimize p i t c h i n s t a b i l i t y w a s the wing t r a i l i n g - e d g e extension. (See f i g . 2 ( b ) . ) The d a t a of f i g u r e 11 show t h a t t h e trailing-edge extension w a s e f f e c t i v e i n reducing the p i t c h i n s t a b i l i t y and, as expected, increased t h e maximum l i f t because of t h e increased wing area. Presented i n f i g u r e 1 2 are d a t a obtained from tests t o determine the e f f e c t i v e n e s s of t h e com- b i n a t i o n of apex notch and wing trailing-edge extension on t h e l o n g i t u d i n a l aerody- namic c h a r a c t e r i s t i c s . The d a t a of f i g u r e 12 show that the combined e f f e c t s of apex notch and wing t r a i l i n g - e d g e extension produced a very favorable e f f e c t on p i t c h i n g moment and maximum l i f t . The maximum unstable pitching-moment c o e f f i c i e n t w a s reduced from about 0.10 t o 0.023, and t h e maximum l i f t c o e f f i c i e n t w a s increased f r o m 1.55 t o 1.65.
Another modification s t u d i e d w a s the a d d i t i o n of wing fences. (See f i g . 2 ( c ) . ) This modification w a s aimed p r i m a r i l y a t improving t h e lateral s t a b i l i t y a t high angles of a t t a c k by f o r c i n g the v o r t i c e s on t h e r i g h t and l e f t wings t o break down more symmetrically and thereby minimize the adverse e f f e c t of asymmetrical vortex breakdown on lateral s t a b i l i t y . The r e s u l t s of adding the wing fences on longitudi- n a l characteristics are presented i n f i g u r e 13 and show that wing fences reduced t h e l e v e l of p i t c h i n s t a b i l i t y above a = 25O b u t w e r e very d e t r i m e n t a l to vortex l i f t , as i n d i c a t e d by the reduction i n the l i f t - c u r v e slope and loss of maximum l i f t .
The r e s u l t s o f tests t o study t h e combined e f f e c t s of t h e wing apex notch, wing trailing-edge extension, and wing fences on t h e l o n g i t u d i n a l aerodynamic c h a r a c t e r i s - tics are summarized i n f i g u r e 14. The d a t a show t h a t t h e combined modifications g r e a t l y reduced the p i t c h i n s t a b i l i t y , although the configuration still showed a mild pitch-up near a = 11O. The modification also reduced the l i f t c o e f f i c i e n t above a = 16O. Comparing f i g u r e 14 with f i g u r e 15 shows that the a d d i t i o n of the v e r t i c a l t a i l had l i t t l e e f f e c t on these r e s u l t s .
L a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s . - Presented i n f i g u r e 16 are the l a t e r a l - d i r e c t i o n a l s t a b i l i t y characteristics of t h e b a s e l i n e configuration compared with those of the modified c o n f i g u r a t i o n with t h e apex notch. The d a t a show some s l i g h t d e t r i m e n t a l e f f e c t s of t h e notch on the d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s .
However, the notch appeared t o provide some improvement i n lateral s t a b i l i t y near a = 36O. The e f f e c t of t h e wing trailing-edqe extension on t h e l a t e r a l - d i r e c t i o n a l s t a b i l i t y characteristics is shown i n the d a t a of f i g u r e 17. The d a t a show very l i t t l e e f f e c t of t r a i l i n g - e d g e extension on t h e d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s - tics but a d e t r i m e n t a l e f f e c t on lateral s t a b i l i t y above a = 31". I n f a c t , t h e b e n e f i c i a l e f f e c t of t h e apex notch a t a = 36O w a s o f f s e t by the d e t r i m e n t a l e f f e c t of the trailing-edge extension. It should be remembered, however, that the apex notch and trailing-edge extension modifications w e r e aimed primarily a t reducing the pitch-up c h a r a c t e r i s t i c s of t h e model and t h a t t h e i r influence on l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s w a s of secondary i n t e r e s t . A s shown i n f i g u r e 18, t h e l a t e r a l - d i r e c t i o n a l s t a b i l i t y characteristics of the configuration with t h e combina- t i o n of t h e apex notch and trailing-edge extension w e r e somewhat i n f e r i o r t o those of t h e b a s e l i n e configuration.
Presented i n f i g u r e 19 are the r e s u l t s of tests to determine the incremental e f f e c t on l a t e r a l - d i r e c t i o n a l s t a b i l i t y of adding t h e wing fences. A s noted i n t h e earlier discussion of the l o n g i t u d i n a l c h a r a c t e r i s t i c s , the fences w e r e conceived as devices f o r d e l i b e r a t e l y f o r c i n g vortex breakdown i n a more symmetrical fashion, p a r t i c u l a r l y under s i d e s l i p conditions, and for a l l e v i a t i n g t h e lateral i n s t a b i l i t y e x h i b i t e d by t h e b a s e l i n e configuration i n t h e region of maximum lift. Although they w e r e d e t r i m e n t a l t o vortex l i f t , the fences accomplished t h e i r primary purpose and eliminated t h e lateral i n s t a b i l i t y near a = 36O. Figure 20 compares t h e lateral- d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of the b a s e l i n e configuration with those of the configuration i n c o r p o r a t i n g a l l t h r e e modifications. The d a t a show t h a t t h e combina- t i o n of modifications w a s g e n e r a l l y d e t r i m e n t a l t o d i r e c t i o n a l s t a b i l i t y throughout lateral s t a b i l i t y below a = 33O while t h e angle-of-attack range and reduced t h e improving it a t higher angles of a t t a c k .
Presented i n f i g u r e 21 are the l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of the b a s e l i n e and modified confiqurations with the v e r t i c a l t a i l on. The combination of modifications reduced the d i r e c t i o n a l s t a b i l i t y a t l o w angles of a t t a c k and r e s u l t e d i n t h e model becoming d i r e c t i o n a l l y unstable near a = 24O. Above a = 31°, the modifications improved d i r e c t i o n a l s t a b i l i t y compared with the b a s e l i n e configu- r a t i o n . The modifications w e r e unfavorable t o lateral s t a b i l i t y below a = 31° b u t produced favorable effects a t higher angles of a t t a c k . The overall impact of the modifications a t high angles of a t t a c k can be summarized by using t h e Cn $ ,dyn parameter. The Cn d a t a f o r t h e b a s e l i n e and modified configurations are pre- $ ,dYn sented i n f i g u r e 22. The d a t a show t h a t t h e modifications reduced the magnitude of C a t moderate angles of a t t a c k ; however, the modifications extended the angle- "$ ,dyn o f - a t t a c k range f o r s t a b l e values of up t o 36O.
CONCLUDING REMARKS The r e s u l t s of a low-speed, e x p l o r a t o r y wind-tunnel study of a cranked-arrow- wing f i g h t e r confiquration can be summarized as follows: 1 . The b a s e l i n e configuration e x h i b i t e d a hiqh l e v e l of maximum l i f t but d i s - played unstable l o n q i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s a t moderate t o high angles of a t t a c k . These c h a r a c t e r i s t i c s w e r e dominated by the s t r o n g vortex flows generated by the highly s w e p t wing a t these conditions.
2. The combination of wing apex notch, wing trailing-edge extension, and wing fences improved t h e l o n q i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of t h e b a s e l i n e configuration a t hiqh angles of a t t a c k b u t w e r e detrimental t o maximum l i f t c o e f f i c i e n t .
Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 A p r i l 20, 1984 REFERENCES 1. M i l l e r , David S.; and Schemensky, Roy T.: Design Study R e s u l t s of a Supersonic Cruise Fighter Wing. A I A A Paper 79-0062, Jan. 1979.
2. Freeman, D e l m a C.,. Jr.: Low Subsonic F l i g h t and Force I n v e s t i g a t i o n of a Super- sonic Transp6rt Model With a Highly Swept A r r o w Wing. N A S A TN D-3887, 1967.
3. Johnson, Joseph L., Jr.; Grafton, Sue B.; and Y i p , Long P.: Exploratory I n v e s t i - g a t i o n of t h e E f f e c t s of Vortex Bursting on the High Angle-of-Attack Lateral- D i r e c t i o n a l S t a b i l i t y C h a r a c t e r i s t i c s of Highly-Swept Wings. A Collection of Technical Papers - AIAA 11th Aerodynamic Testing Conference, Mar. 1980, pp. 282-297. (Available as AIAA-80-0463.)
4. H u m m e l , D i e t r i c h (MacAdam, E. J., t r a n s l . ) : Research on Vortex Breakdown on Slender D e l t a Wings. A.R.A. Lib. Transl. N o . 1 2 , A i r c r a f t R e s . Assoc. Ltd., O c t . 1965.
5. Fennell, L. J.: vortex Breakdown - Some Observations i n F l i g h t on the Kp 115
A i r c r a f t . R. & M. No. 3805, B r i t i s h A.R.C., 1977.
6. Chambers, Joseph R.; and Anglin, Ernie L.: Analysis of Lateral-Directional Sta- b i l i t y C h a r a c t e r i s t i c s of a Twin-Jet Fighter Airplane a t High Angles of Attack.
N A S A TN D-5361, 1969.
TABLE I.- GEOMETRIC CHARACTERISTICS OF BASELINE MODEL Fuselage :
Length. f t ................................................................... 7.34
Wing:
A r e a . f t 2 .................................................................... 13.5
Span. f t .................................................................... 4.86
Aspect r a t i o ................................................................. 1.75
Taper ratio ................................................................. 0.10
Leading-edge sweep: Outboard. deg ..............................................................
Trailing-edge sweep. deg ..................................................... 18.5
Mean aerodynamic chord. f t ................................................... 3.71
A i r f o i l ............................................................. NACA 64A003.2
Root chord. f t ............................................................... 6.93
............................................................... 0.58
Tip chord. f t V e r t i c a l tail:
A r e a . f t 2 .................................................................... 0.55
.................................................................... 0.85
Span. f t
Aspect r a t i o ................................................................. 1.30
Taper ratio ................................................................. 0.30
A i r f oi 1 :
R o o t ....................................................... 6 percent biconvex
Tip ....................................................... 3.5 percent biconvex
( a ) Wing leading-edge apex notch.
Figure 2.- Modifications to b a s e l i n e configuration. Dimensions are given i n f e e t .
1 1 (b ) Wing t r a i li ng-e dge e x t e n sion.
Figure 2.- Continued.
I
t I
1.51
T ''Wing Fence
(c ) Wing fences.
Figure 2 . - Concluded.
Figure 3.- Model support system for static-force test.
.08
Vertical tail
o On
I3 Off
- . 0 q
1.6 1.2 1.0
- -2
-10 0 10 20 30 50 .12 -08 .OLf 0 -.oq
a, deg
cm
Figure 4, - S t a t i c l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s of b a s e l i n e configuration.
#3 +40 f3 = oo a = 30° a = 35O L-84-43 Fiqure 5.- Smoke flow v i s u a l i z a t i o n showing vortex breakdown on baseline configuration.
-2
CY
-
- . 2 ~ I I I I I I I I I r i I
0 , deg
.02 0 1
c*
0 2 1
A 26 - -02
- 04
- -02 -.OY
- -06
-15 -10 -5 0 5 10 15
B , deg
Figure 6 . - Variation of static lateral-directional coefficients with sideslip angle. Baseline configuration.
.2
CY
- 92 1 1 1 1 1 1 1 1 1 1 1 1 1
.02
- -02
- 9 0 1 1 -15 -10 -5 0 5 10 15
8, deg
Figure 6.- Concluded.
. 2
- 2 1 1 I I I I I I 1 1 I I I
* 06
9 0 1 1 Vertical tail
o On
4 2 0 Off -.02 - . 0 1 1 .02
G I 0
- a 0 2 -15 -10 -5 0 5 10 15
B , deg
( a ) a = 3 1 O .
Figure 7 . - Variation of s t a t i c lateral-directional coefficients for baseline configuration with vertical t a i l off and on.
. 2
cy
-21 I I I I I I I I 1 I I 1
- 06
Vertical tai I
0 On Off -.02 -15 -10 -5 0 5 10 15
B , deg
(b) 01 = 3 6 O .
Figure 7 . - Continued.
Vert ica I tail
o On
0 Off -.02 - . 0 q (c) a = 41°.
Figure 7.- Concluded.
Vertical tail
,002 - .ooq - -006 - .008 -002
- a002
- .OOY
0 10 20 30 YO 50
a, deg
Figure 8.- Effect of vertical t a i l on s t a t i c lateral-directional s t a b i l i t y characteris- tics of baseline configuration.
rn 02
Vertical tai I
-01
0 Off
0 On
-mol
- -02
-10 0 10 20 30 1 1 0 so
a, deg
Figure 9.- Variation of C with angle of a t t a c k . Baseline configuration.
$ ,Wn 0 12 * 08
cm
0 OLf
Configuration
-.m
0 Baseline
0 Apex modification
1 =6
1 . 4
1 . 2
1 *o
CL
d 8
CD &
e 4
.2
- . 2
0 -.w
-10 0 10 20 30 50
a, deg
=12 *08 &
Figure 10.- K f e c t of apex modification on s t a t i c longitudinal aerodynamic characteristics. Vertical t a i l off.
cm *OV
Wing t railingedge extension
0 Off
0 On
- 0 0 4
1 *8 1 e 6
1 . Q
1.2
CL 1 . 0
*8
CD
*6 0 2
- . 2
”
-1 0 0 10 30 YO 50 20 - -04 0 0 8 .OY
a, deg
%lo
Figure 11. - E f f e c t of wing trailing-edge extension on static l o n g i t u d i n a l
aerodynamic c h a r a c t e r i s t i c s . Vertical t a i l o f f ; apex modification on.
. 1 2
Cm
OY
Configuration
0 Baseline
0 Apex modification
0 Apex modification +
- - 0 1 1
wing trailing-edge
ext en si0 n
1 . 8 1-6
1 . Y
1 . 2
CL 1.0
-8
%
*6 . 2
- e 2
-10 0 10 20 30 1 1 0 50
a, deg
Figure 12.- E f f e c t of apex notch and wing t r a i l i n g - e d g e extension on static l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s .
0 oq
Clm 0
Wing fences
- o m I I I I I I I I h I I
0 Off
0 On
1 *8 1.6 1 *q 1.2
GL 1.0
*8
CD
*6 e 2
- 02
-10 0 10 20 30 q0 50 0 q
a, deg
c
Figure 13.- E f f e c t of wing fences on s t a t i c l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s . V e r t i c a l t a i l o f f ; apex modification on; wing t r a i ling-edge extension on.
.12 *08
Configuration
- *OY
0 Baseline
0 Modifications
co rn bi n ed
1 *6 1 .Y 1 e 2 .2
- 02
0 - a 0 9 -10 0 10 20 30 90
a, deg
*12 ‘08 c
Figure 14.- E f f e c t of combination of apex notch, wing trailing-edge extension, and fences on s t a t i c l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s . Vertical t a i l o f f .
* 12 .08
c m
. OY
Configuration
- .OY
0 Baseline
0 Modifications
1.6 1 .Ll 1.2
.o
CL
.8 .6
CD
.Y .2 - .2 .12 .08 .OY 0 -.OY -10 0 10 20 30 YO
a, deg
Crn
Figure 15. - E f f e c t of combination of apex notch, wing t r a i l i n g - e d g e extension, and fences on s t a t i c l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s . V e r t i c a l t a i l on.
cy,
- .02
Configuration
Apex modification
c
" B - .002
- .004
.002 - -002 - .ooLi 0 10 20 30 90 50
a, deg
Figure 16.- Effect of apex modification on l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of baseline configuration.
- .02
Wing t ra ii i ng-edge extension
- -006
- -002
- .009
0 10 20 30 LfO 50
a, deg
Figure 17.- E f f e c t of wing trailing-edge extension on l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of base- l i n e configuration. Apex modification on; v e r t i c a l t a i l o f f .
- =02 Configuration
Base I in e
a Apex modification
Apex modification t
wing trailing-edge
c -moo2 exten s ion
"B
- mooq
- no06
,002
- - 002
- 0004
0 10 20 30 q0 50
a, deg
Figure 18.- Effect of combination of wing apex notch and wing t r a i l i n g - e d g e extension o n l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of test con- f i g u r a t i o n . Vertical t a i l off.
- -02
Wing fences
(-J -.002
" B
- .ooq
- 9006
0 002
- -002
- .ooq
0 10 20 30 LiO 50
a, deg
Figure 19.- E f f e c t of wing fences on l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of test configuration.
V e r t i c a l t a i l o f f ; apex modification and wing t r a i l i n g - edge extension on.
Configuration
Combined modifications
C
B- ‘002
- .ooq
,002
- ,002
- .009
0 10 20 30 q0 50
a, deg
Figure 20.- E f f e c t of combination of wing apex notch, wing t r a i l i n g - e d g e extension, and wing f e n c e s on lateral- d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of b a s e l i n e con- f i g u r a t i o n . V e r t i c a l t a i l o f f .
0 02
cys 0
- -02
Configuration
.002
Basel in e
E l
Mod if i cat io n s
co r n bi n ed
- -004
- .OO6
.002 - .002
- -004
0 10 20 30 40 50
a, deg
Figure 21.- E f f e c t of combination of wing apex notch, wing t r a i l i n g - e d g e extension, and wing fences on lateral- d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of b a s e l i n e con- f i g u r a t i o n . V e r t i c a l t a i l on.
.02
Configuration
Base I i n e
CI Mod i f k a t ions
combined
- .01
-10 0 10 20 30 40 50
a, deg
Figure 22.- E f f e c t of combined modifications on v a r i a t i o n of C with a n g l e of a t t a c k . V e r t i c a l t a i l on. ”$ ,dyn 1. Report No. 2. Government Accession No. 3. Recipient’s Catalog No.
N A S A TM-85776
I
I 4. Title and Subtitle 5. Report Date LOW-SPEED WIND-TUNNEL STUDY O F THE HIGH-ANGLE-OF-ATTACK May 1984 ~ STABILITY AND CONTROL CHARACTERISTICS OF A CRANKED- 6. Performing Organization Code ARROW-WING FIGHTER CONFIGURATION 505-43-1 3-01
I
7. Author(s) 8. Performing organization Report No.
Sue B. Graf ton L-15762 10. Work Unit No.
9. Performing Organization Name and Address N A S A Langley Research Center 11. Contract or Grant No.
Hampton, VA 23665 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technic a 1 Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546
I I
15. Supplementary Notes 16. Abstract The low-speed, high-angle-of-attack s t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s of a f i g h t e r c o n f i g u r a t i o n incorporating a cranked arrow wing w e r e i n v e s t i g a t e d i n t h e Langley 30- by 60-Foot Tunnel. The study w a s conducted a s p a r t of a NASA/General Dynamics cooperative research program t o i n v e s t i g a t e t h e a p p l i c a t i o n of advanced wing designs t o combat a i r c r a f t . Tests w e r e conducted on a b a s e l i n e configuration and on s e v e r a l modified configurations. The r e s u l t s of t h e i n v e s t i g a t i o n showed t h a t t h e b a s e l i n e configuration e x h i b i t e d a high l e v e l of maximum l i f t but displayed undesir- a b l e l o n g i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s a t high angles of a t t a c k . Various wing modifications w e r e made which improved t h e l o n g i t u d i n a l and l a t e r a l - d i r e c t i o n a l s t a b i l i t y c h a r a c t e r i s t i c s of t h e configuration a t high angles of a t t a c k . However, m o s t of the modifications were d e t r i m e n t a l t o m a x i m u m l i f t .
I 17. Key Words (Suggested by Author(s)) 18. Distribution Statement
Unclassified - Unlimited
Arrow wings S t a t i c Fighter S t a b i l i t y Longitudinal Lateral d i r e c t i o n a l Subject Category 08 A03 Unclassified Unclassified 37 For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA-Langley, 1984