FULL-SCALE WIND-TUNNEL INVESTIGATION OF A FLEXIBLE-WING MANNED TEST VEHICLE

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D-1946

FULL-SCALE WIND-TUNNEL INVESTIGATION OF A FLEXIBLE-WING MANNED TEST VEHICLE By Joseph L. Johnson, Jr., and James L. Hassell, Jr.

Langley Research Center Langley Station, Hampton, Va.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION August 1963 WASHINGTON NATIONAL AERONAUTICS AND SPACE ADMINISTRATION NOTE D-1946 TECHNICAL FULL-SCALE WIND-TUNNEL DIVESTIGATION O F A FLEXIBLE-WING MANNED TEST VEHICLI3 By-Joseph L. Johnson, Jr., and James L. Hassell, JI-.

SUMMARY An i n v e s t i g a t i o n has been conducted i n .;he Langley f u l l - s c a l e t u n n e l t o determine t h e performance and s t a t i c 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 flexible-wing manned t e s t vehicle. This a i r p l a n e i s a s i m p l i f i e d research machine, which c o n s i s t s b a s i c a l l y of a cargo platform attached t o a parawing , by means of an overhead t r u s s arrangement. In a d d i t i o n t o t h e b a s i c c o n t r o l t e s t s , a f e w t e s t s were made t o evaluate s e v e r a l a l t e r n a t e c o n t r o l systems which involved d e f l e c t i o n s of t h e a f t portion of t h e parawing k e e l and wing t i p s .

The t u n n e l t e s t s showed t h a t t h e maximum l i f t c o e f f i c i e n t of t h e a i r p l a n e occurred at a k e e l angle of a t t a c k o f 420 and w a s 1.24 with power o f f and 1 . 3 3 with power on. The maximum l i f t - d r a g r a t i o vas about 5 . 5 . With s t i c k f i x e d , a t k e e l angles of t h e a i r p l a n e had about n e u t r a l s t a t i c 1 o n g i t J d i n a l s t a b i l i t y a t t a c k below 2 0 ° , a moderate degree of s t a b i l i t y from 2 0 ° t o 3 5 O , and longitudi- n a l i n s t a b i l i t y , o r pitch-up, from 3 5 O t o 420. A t a k e e l angle of a t t a c k of 42O t h e a i r p l a n e again became stable. With t h e s t i c k f r e e , t h e l o n g i t u d i n a l s t a b i l i t y 1 w a s g e n e r a l l y worse with t h e a i r p l a n e being unstable a t t h e lower angles of a t t a c k , about n e u t r a l l y s t a b l e i n t h e intermediate range, and unstable a t t h e ' higher angles of a t t a c k . The a i r p l a n e , i n general, w a s d i r e c t i o n a l l y s t a b l e and had p o s i t i v e e f f e c t i v e dihedral throughout t h e angle-of-attack range investigated.

The l a t e r a l c o n t r o l provided by banking t h e wing d i d not appear t o be s a t i s f a c - , t o r y because of inadequate r o l l i n g moments and excessively high s t i c k f o r c e s .

Analysis of t h e tunnel d a t a i n d i c a t e d t h a t t h e rudder w a s g e n e r a l l y a b e t t e r r o l l - c o n t r o l device with power on (inasmuch as t h e rudder i s i n t h e s l i p s t r e a m of t h e pusher p r o p e l l e r ) than t h e wing-bank c o n t r o l system provided on t h e a i r p l a n e .

The rudder w a s not very e f f e c t i v e with power o f f . The hinged wing-tip c o n t r o l device tested on t h e a i r p l a n e (which had been developed e a r l i e r a t t h e Langley Research Center i n small-scale model t e s t s ) appeared promising i n t h a t it pro- vided higher r o l l i n g e f f e c t i v e n e s s and lower estimated s t i c k f o r c e s than those of t h e wing-bank c o n t r o l system provided on t h e a i r p l a n e .

INTRODUCTION For t h e p a s t f e w years, t h e National Aeronautics and Space Administration has been conducting a general i n v e s t i g a t i o n t o provide some b a s i c information on configurations employing the parawing concept. (For example, see refs. 1 to 3. ) This early work eventually led to the design and construction of a flexible-wing airplane configuration, which was proposed as a test vehicle to demonstrate flight characteristics of the parawing concept as well as to provide a prototype for the development of a manned combat utility vehicle. This airplane is a simplified research machine which consists basically of a cargo platform attached to a para- wing by means of an overhead truss arrangement. The vehicle is powered by a pusher propeller located at the aft end of the platforni and has a cockpit located at the front. Control is obtained by banking or pitching the wing with respect to the cargo platform. A rudder operating in the propeller slipstream provides directional control. A model generally similar in design to the vehicle of this investigation was flight tested in the Langley full-scale tunnel, and the results

of this investigation are reported in reference 4. A preliminary flight evalua-

tion of the full-scale configuration of the present investigation is given in ref-

erence 5. A s part of the overall research effort on the parawing concept, a

force-test investigation has been conducted in the Langley full-scale tunnel on the full-scale vehicle to determine static stability and control and performance parameters for correlation with the earlier flight tests as well as to extend the present research program to include wind-tunnel data on large-scale parawings.

The present investigation consisted of static tests to determine the basic aerodynamic and longitudinal and lateral stability and control characteristics of the airplane over an angle-of-attack range of the keel from about 14O to 44O with power off and on. These tests were conducted at several different values of dynamic pressure to evaluate the effects of aerodynamic loading on the character- istics of the wing up to simulated steady level flight ( l g ) conditions. Included in the investigation were tests of the airplane with the rudder off and on and with the wing off. In addition, tests were made to study the effects of boltrope and batten modifications to the parawing trailing edge. A few tests were also made to evaluate several alternate means of providing control. Comparisons of the wind-tunnel data with flight-test data obtained on the airplane (ref. 5) have been made where possible.

SYMBOLS A l l forces, moments, and velocities are presented with respect to the I stability-axis system originating at the reference center-of-gravity position shown in figure 1. All measurements are reduced to coefficient form and are based on the dimensional characteristics of the flat plan geometry of the wing I (45' leading-edge sweep).

S wing area, sq ft b wing span, ft keel length, ft ' k

v free-stream velocity, fps

I

free-stream dynamic pressure, l b / s q f t ' 9 angle of a t t a c k of keel, deg uk angle of a t t a c k of platform, deg

' a p

angle of incidence of parawing k e e l with respect t o platform, iW

Qir - "p., deg

angle of s i d e s l i p , -+, deg

P angle of yaw, deg I * angle of roll, p o s i t i v e r i g h t wing t i p down, deg weight, l b l i f t , l b drag, l b l i f t - d r a g r a t i o wing loading, l b / s q f t s i d e force, l b t h r u s t , l b hinge moment ( p o s i t i v e when M h -;ends t o d e f l e c t k e e l o r wing-tip t r a i l i n g edge downward i n t h e XZ-plane o r wing-tip t r a i l i n g edge outward i n t h e XY-plane), f t - l b pitching moment, f t - l b r o l l i n g moment, f t - l b yawing moment, f t - l b l i f t c o e f f i c i e n t , L/qS drag c o e f f i c i e n t , D/qS t h r u s t c o e f f i c i e n t ,

[CD (power On) - C D (power off, p r o p e l l e r stoppedflap=Oo

hinge-moment c o e f f i c i e n t , %/qSb f o r r o l l , %/qSck, f o r p i t c h 'h lateral-force coefficient, Fy/qS Cm pitching-moment coefficient, My/qSck pitching-moment coefficient at zero lift Cm. o Cn yawing-moment coefficient, %/qSb rolling-moment coefficient, Mx/qSb Cl slope of pitching-moment curve with lift coefficient

c = - per deg

YP aP'

32 ft/sec2 acceleration due to gravity, g P rolling velocity, radians/sec 6, rudder deflection, deg 6t wing-tip deflection, deg I horizontal and vertical distances from airplane center of gravity to X?z wing pivot, respectively, ft AIRPLANE AND APPARATUS

I

A three-view drawing of the airplane and photographs of the airplane mounted for force testing in the Langley full-scale tunnel are presented in figures 1 and 2, respectively. Characteristics of the airplane are presented in table I. The parawing used on the vehicle consisted of a dural box-beam keel A and two airfoil-shaped leading edges hinged together at the apex of the wing.

fixed leading-edge sweep angle of 50° was maintained by a spreader bar which was attached to the parawing leading edges and to the keel at approximately the The f a b r i c used t o form t h e membrane of' t h e parawing w a s 35-percent k e e l s t a t i o n .

made of 7-ounce-per-square-yard dacron impregnated w i t h weather-resistant poly- e s t e r . The warp of t h e c l o t h w a s p a r a l l e l t o t h e k e e l . The t r a i l i n g edge of t h e parawing was scalloped and had b a t t e n s and a boltrope (3/32-inch a i r c r a f t c a b l e ) i n s t a l l e d . The boltrope had a 1-inch asymme1,ric s e t t i n g t o provj.de l a t e r a l t r i m .

These modifications t o t h e t r a i l i n g edge were made i n t h e prelimLnary f l i g h t evaluation program (see r e f . j ) and were considered as p a r t of t h e b a s i c a i r p l a n e configuration.

The wing could be pitched o r r o l l e d about t h e p i v o t p o i n t through a system of bellcranks, cables, and push-pull rods. A n e l e c t r i c a c t u a t o r mounted on t h e para- wing k e e l w a s used t o p o s i t i o n t h e wing forward o r rearward with r e s p e c t t o t h e pivot p o i n t i n order .to provide a means of t:rimmi.ng t h e machine l o n g i t u d i n a l l y .

Power f o r t h e vehicle w a s supplied by a 180-horsepower engine and a f i x e d p i t c h propeller. A rudder w a s mounted t o t h e platform s t r u c t u r e and t o t h e wing k e e l d i r e c t l y behind t h e pusher p r o p e l l e r t o provide d i r e c t i o n a l c o n t r o l .

The a i r p l a n e w a s mounted f o r f o r c e t e s t i n g by a t t a c h i n g t h e t u n n e l support s t r u t s r i g i d l y t o t h e f r o n t and r e a r of t h e platform. (See f i g . 2 . ) "he wing w a s remotely pitched o r r o l l e d with respect t o t n e platform through a i r c r a f t a c t u a t o r s which were i n s t a l l e d i n t h e l o n g i t u d i n a l and l a t e r a l c o n t r o l systems of t h e air- plane. The wing d e f l e c t i o n angles were measured at t h e wing pivot p o i n t . The gearing r a t i o of c o n t r o l column, o r wheel d e f l e c t i o n t o wing d e f l e c t i o n , w a s 8.5 i n p i t c h and 7.9 i n roll.

Several a l t e r n a t e c o n t r o l systems, which required some modifications t o t h e b a s i c s t r u c t u r e , were tested on t h e a i r p l a n e . These i n s t a l l a t i o n s a r e i l l u s t r a t e d i n f i g u r e 3 and shown i n photographs of a c t u a l t e s t setups i n f i g u r e 4. One modi- f i c a t i o n consisted of removing t h e f a b r i c from t h e rear portion of t h e k e e l and r e a t t a c h i n g t h e f a b r i c t o a rectangular lightweight framework which w a s hinged t o t h e k e e l so t h a t it could be d e f l e c t e d up and down f o r p i t c h i n g c o n t r o l . Another modification consisted of removing t h e f a b r i c from one wing t i p and r e a t t a c h i n g t h e f a b r i c t o a c o n t r o l arm which w a s hinged and allowed t o move inward and out- w a r d f o r roll c o n t r o l (designated wing-tip c o n t r o l A ) . A t h i r d modification involved t h e i n s t a l l a t i o n of a hinged l o n g i t u d i n a l m e m b e r near t h e wing t i p which could be d e f l e c t e d i n a v e r t i c a l plane t o change t h e b a s i c wing contour a t t h e t r a i l i n g edge t o provide c o n t r o l (designated. wing-tip c o n t r o l B ) . I n c o n t r o l system B, t h e f a b r i c remained attached t o t h e wing l e a d i n g edge, and as t h e hinged l o n g i t u d i n a l m e m b e r w a s d e f l e c t e d downward, t h e wing f a b r i c w a s permitted t o seek i t s own p o s i t i o n under t h i s m e m b e r . ( I n o t h e r words, t h e f a b r i c w a s not attached t o t h e hinged m e m b e r and t h e r e f o r e d i d n o t transmit any l a t e r a l hinge moments.) I n a l l of t h e s e modifications, provision w a s made f o r t h e i n s t a l l a t i o n of s t r a i n gages t o allow f o r t h e determination of hinge-moment and s t i c k - f o r c e c h a r a c t e r i s t i c s .

I TESTS The i n v e s t i g a t i o n w a s conducted i n t h e Langley f u l l - s c a l e t u n n e l .

A complete d e s c r i p t i o n of t h e tunnel and t e s t apparatus i s given i n reference 6. The s t a t i c l o n g i t u d i n a l and l a t e r a l 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 t h e a i r p l a n e were determined from f o r c e measurements obtained from t h e t u n n e l scale-balance system f o r a range of angles of a t t a c k of t h e k e e l from about 14' t o 44O f o r sev- e r a l values of wing incidence, dynamic pressure, and power s e t t i n g s .

The power-off t e s t s were made with t h e p r o p e l l e r stopped, and no t e s t s were made with t h e p r o p e l l e r re!iisved. Since t h e drag of t h e stopped p r o p e l l e r w a s probably very s m a l l , it w a s assumed t h a t t h e t h r u s t c o e f f i c i e n t Tc w a s zero f o r t h e power-off t e s t s . The power-on t e s t s were made by holding t h e p r o p e l l e r r o t a - t i o n a l speed constant over t h e angle-of-attack range investigated. Several power- on runs were made i n t h i s manner t o cover a range of t r i m conditions. I n both t h e power-off and power-on cases, tests w e r e made at wing-incidence angles of 22-, which correspond approximately t o t h e wing-incidence range i n v e s t i - 25O, and 28-, gated i n t h e f l i g h t t e s t s of reference 5.

The 2 5 O incidence w a s considered t h e b a s i c condition, however, and most of t h e t e s t s were made a t t h i s condition. Most of t h e t e s t s were conducted a t a d.pamic pressure of 3.07 pounds per square f o o t .

Included i n t h e i n v e s t i g a t i o n , however, were tests a t s e v e r a l d i f f e r e n t values of dynamic pressure t o evaluate t h e e f f e c t s Of aerodynamic loading on t h e aerodynamic c h a r a c t e r i s t i c s of t h e configuration up t o simulated steady l e v e l f l i g h t ( l g ) conditions.

The lateral s t a b i l i t y tests were made a t s i d e s l i p angles of 5 O and -5O, and t h e l a t e r a l c o n t r o l t e s t s were made at wing r o l l angles of 5 ' and - 5 O and at rud- der d e f l e c t i o n angles ranging from -200 t o 2 0 ° .

Included i n t h e i n v e s t i g a t i o n w e r e t e s t s of t h e a i r p l a n e with t h e rudder o f f and on and with t h e wing o f f . I n addition, a few t e s t s were made t o o b t a i n some information concerning t h e e f f e c t of t r a i l i n g - e d g e boltrope tension and changes i n b a t t e n geometry ( l e n g t h and arrangement) on t h e s t a b i l i t y and c o n t r o l charac- t e r i s t i c s of t h e a i r p l a n e . A number of tests were a l s o made t o evaluate s e v e r a l proposed a l t e r n a t e c o n t r o l systems which included a hinged-keel ( t r a i l i n g edge) c o n t r o l system f o r p i t c h c o n t r o l and two hinged-wing-tip c o n t r o l systems f o r l a t - e r a l c o n t r o l .

The range of dynamic pressures used i n t h e i n v e s t i g a t i o n v a r i e d from about

I

~ 1.60 t o 5.60 pounds p e r square f o o t , which corresponds t o an airspeed range from about 37 t o 69 feet per second a t standard sea-level conditions and t o a ReJpolds number range from about 6 . 6 x lo6 t o 12.4 X 10 based on t h e parawing k e e l length of 28 feet.

CORRECTIONS The f o r c e and moment d a t a presented have been corrected f o r airstream- misalinement, jet-boundary, and blockage e f f e c t s . Because of t h e l a r g e s i z e of t h e a i r p l a n e with respect t o t h e tunnel t e s t section, it w a s necessary t o mount t h e a i r p l a n e f a i r l y c l o s e t o t h e ground board ( r a t i o of height of wing pivot above t h e ground board t o wing span i s 0.50) i n o r d e r t o allow high angles of a t t a c k .

I n order t o properly represent f l i g h t out of ground e f f e c t , these d a t a should be corrected t o account f o r t h e e f f e c t s of ground proximity on l i f t , drag, and pitching moment. Although t h e r e a r e no methods a v a i l a b l e for making accurate cor- r e c t i o n s f o r t h e ground e f f e c t i n t h i s case, a general i n d i c a t i o n of t h e magnitude of t h e e f f e c t can be obtained from previous i n v e s t i g a t i o n s with delta-wing models i n and out of t h e presence of t h e ground. (See ref. 7. ) Based on t h i s a v a i l a b l e information, ground-effect corrections have been made t o t h e b a s i c d a t a i n a num- b e r of cases and a r e presented f o r reference purposes.

RESULTS AND DISCUSSION Before t h e f o r c e - t e s t r e s u l t s a r e discussed, it appears desi.rable t o first point out some of t h e more pronounced wing i r r e g u l a r i t i e s noted i n t h e investiga- t i o n s i n c e t h i s information may be u s e f u l i n 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 .

Visual observations and camera records were made t o obtain some i n d i c a t i o n of t h e changes i n t h e wing f a b r i c and support members as t h e tunnel test, conditions were varied. Representative photographs obtained during some of t h e t e s t s a r e pre- sented i n f i g u r e 5 .

It w a s observed i n t h e wind-tunnel t e s t s of t h e b a s i c configuration t h a t a t k e e l angles of a t t a c k below about 20° t h e aft portion of t h e wing (and i n p a r t i c - u l a r t h e inboard s e c t i o n ) f l u t t e r e d badly. It appeared t h a t t r a v e l i n g waves moved rearward along t h e wing with amplitudes t h a t increased as t h e angle of a t t a c k w a s reduced. In addition, high-frequency t r a i l i n g - e d g e f l u t t e r w a s very pronounced at k e e l angles of a t t a c k below about 20°. A s t h e angle of a t t a c k w a s increased above 20°, t h e t r a i l i n g - e d g e f l u t t e r and wave motion of t h e fabr.ic became l e s s apparent and appeared t o s t o p completely nea:r an angle of a t t a c k of about 2T0, but near t h i s angle of attack, a l a r g e depression formed i n t h e a f t s e c t i o n of t h e wing, j u s t ahead of t h e battens, and became more pronounced with increasing angle of a t t a c k . It appeared t h a t t h e wing t r a i l i n g edge had considerably more down- ward d e f l e c t i o n a t than a t t h e lower angles of a t t a c k probably because ak = 27O t h e boltrope r e s t r i c t e d t h e t r a i l i n g edge while t h e f a b r i c forward of t h e t r a i l i n g edge s t r e t c h e d as a r e s u l t of t h e increased loading a t t h e higher angles of a t t a c k .

The wing contour changes noted i n t h e wind-tunnel t e s t s a r e generally s i m i l a r t o those observed i n t h e preliminary f l i g h t evaluation t e s t s reported i n r e f e r - ence 5 . It i s believed t h e r e f o r e t h a t t h e f o r c e - t e s t r e s u l t s a r e applicable f o r use i n i n t e r p r e t i n g t h e f l i g h t - t e s t r e s u l t s . It should be pointed out, however, t h a t t h e a p p l i c a t i o n of these r e s u l t s t o o t h e r parawing arrangements having d i f - f e r e n t wing c h a r a c t e r i s t i c s (such as m a t e r i a l , d i r e c t i o n of f a b r i c weave and seams, and leading- and trailing-edge shapes) may be d e f i n i t e l y l i m i t e d .

It should be pointed out t h a t i n t h e t u n n e l t e s t s and at times i n t h e f l i g h t t e s t s t h e r e were l a r g e f l u c t u a t i o n s i n both t h e p i t c h and r o l l c o n t r o l f o r c e s . A sample of d a t a from t h e control-force measurements made i n t h e tunnel t e s t s i s shown i n f i g u r e 6. Fluctuations i n s t i c k f o r c e of as much as f20 pounds were obtained, and t h e r e w a s a l s o a s h i f t i n t h e general l e v e l of t h e readings from one time t o another. The data represented by t h e s o l i d l i n e i n f i g u r e 6(a) were obtained at the beginning of a test run at a keel angle of attack of 220, whereas the data for the dashed line were obtained several minutes later under presumably identical test conditions after runs had been made at higher anglcbs of attack.

The fluctuations in the data and the shift in level of readings from one time to another are believed to be related to such factors as trailing-edge flutter, flex- ibility, and fabric stretch.

I Longitudinal Stability and Control Aerodynamic data for basic configuration.- The basic longitudinal data for

the airplane configuration are presented in figures 7 and 8 for the power-off

and power-on conditioos for wing incidences of 22.5O, 2 5 O , and 2 8 . 5 O . The data

of figure 7 are plotted against the angle of attack of the wing keel whereas the

data for figure 8 are plotted against angle of attack of the platf3rm. The data of these figures were obtained with the dynamic pressure held constant during the test run. In order to represent a lg flight condition (lift e'qualto air- craft weight), these data require certain corrections which can be made by using of figure 9.

! , h e data

Figure 9 presents lift, drag, and pitching-moment data for the power-off

obtained in test runs at different values of constant i , = 2 5 ' condition at d;inamic pressure, ranging from 1.60 to 5.60 pounds per square foot which corre- sponds to airspeeds of about 22 to 41 knots. The data of figure 9 show a con- sistently greater negative pitching moment with increasing dynamic pressure.

The dashed curve intersecting the pitching-moment curves of figure 9 represents the pitching moments for a l g flight condition. This curfe was obtained from

w/s

the basic relationship CL = - ( o r qCL = W/S) by using a value of W/S of

3.32 lb/sq ft for the airplane. The dashed curve intersects each of the other curves at the lift coefficient where the product of CL and the measured q is equal to the airplane wing loading. The curve representing the lg flight con- dition has a flatter slope (and therefore, less static longitudinal stability) than the curves obtained at the higher values of constant dynamic pressure, par- ticularly in the low and moderate lift-coefficient range. A l g curve for lift and drag data was not presented since the effect of dynamic pressure was gener- I ally small and inconsistent in these cases.

For ease of comparison, the pitching-moment curve for l g flight and the curve obtained at the constant dynamic pressure of 3.07 lb/sq ft used in most of the tunnel tests are replotted in figure 10. Since the effects of dynamic pressure were determined only for the power-off condition at iw = 2>O, the pitching- moment data for other test conditions were corrected to l g conditions by using the increments between the two pitching-moment curves of figure 10. The data cor- rected in this manner for the various wing-incidence and power conditions of fig- ures 7 and 8 are presented in figure 11.

The lift curves of figure 11 appear to be normal with a lift-curve slope slightly greater than 0.05 per degree with power on and slightly less than 0.05 with power off. The maximum lift coefficient is obtained at a keel angle of attack of about 42O and is 1.24 with power off and 1.35 with power on. Although d a t a are presented f o r a combination of platform angle of a t t a c k and wing- incidence angle corresponding t o angles of a t t a c k o f t h e k e e l as low as 1 4 O , t h e wing t r a i l i n g - e d g e f l u t t e r which occurred a t k e e l angles of a t t a c k below about 20° probably makes it undesirable t o operate t h e a i r c r a f t a t angles below t h i s value.

Figure 12 shows f a i r l y good agreement between l i f t c o e f f i c i e n t s measured i n f l i g h t The l i f t curve i n f i g u r e 1 2 for t h e t u n n e l t e s t s i s an and i n t h e t u n n e l t e s t s .

average of t h e power-on l i f t curves of f i g u r e 11.

Figure 1 1 shows t h a t t h e m a x i m u m of t h e a i r p l a n e i s about 5.5 and i s L/D

'

obtained at a k e e l angle of a t t a c k of about :17O o r 280. An estimate of t h e L/D f o r t h e wing alone of about 7 was made from -,he d a t a of f i g u r e 1 1 1 t o g e t h e r with t h e d a t a f o r t h e platform alone shown i n f i g u r e 8 ( b ) . It should be pointed o u t

I

l t h a t t h e L/D of t h e wing alone cannot be determined d i r e c t l y b y s u b t r a c t i n g t h e d a t a of t h e platform alone from t h e d a t a f o r t h e complete configuration This because of some favorable i n t e r f e r e n c e e f f e c t of t h e platform on t h e wing.

4) of a small-scale configuration s i m - e f f e c t became apparent i n t e s t s ( s e e r e f .

ilar t o t h e vehicle of t h e present i n v e s t i g a t i o n i n which d a t a were obtained f o r t h e wing alone, platform alone, and wing-platform combination.

The pitching-moment d a t a of f i g u r e 1 1 show n e u t r a l s t a t i c l o n g i t u d i n a l sta- b i l i t y at k e e l angles of a t t a c k below about 20°, a moderate degree of s t a b i l i t y 350, and 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, from about 3 5 O from about 200 t o t o 42O. A t 42O t h e d a t a of f i g u r e l l ( c ) ind.icate a s t a b i l i z i n g break i n t h e pitching-moment curve. The pitch-up noted f o r t h i s configuration a t high angles of a t t a c k i s unusual f o r parawing configuratAons based on tests of small-scale models which showed s t a b l e , pitch-down moments a t t h e stall. One p o s s i b l e expla- n a t i o n f o r t h e pitch-up of t h e f u l l - s c a l e v e h i c l e might be t h e p a r t i c u l a r changes i n t h e t r a i l i n g - e d g e contour, which appeared t o be more severe than previously noted i n o t h e r parawing s t u d i e s .

Presented i n f i g u r e 13 are 1ongitudina:L d a t a f o r t h r e e wing p i v o t p o s i t i o n s which cover t h e wing forward and a f t p i v o t : L i m i t s a v a i l a b l e on t h e a i r p l a n e f o r l o n g i t u d i n a l t r i m . The incremental changes i n p i t c h i n g moment :indicated by t h e data closely approximate t h e changes t h a t would be expected from consideration of t h e center-of-gravity s h i f t s corresponding t o these wing p o s i t i o n changes. For example, a &-inch s h i f t i n wing p i v o t (which corresponds t o about a 1-percent a 1-percent change i n s t a t i c change i n center-of-gravity p o s i t i o n ) produced about l i f t and drag with changes i n t h e wing p i v o t margin. Incremental changes i n t h e l o c a t i o n a r e probably i n d i c a t i v e of v a r i a t i o n s i n t h e i n t e r f e r e n c e e f f e c t s between t h e wing and fuselage.

As pointed o u t previously, t h e a i r p l a n e probably experienced ground e f f e c t on l i f t , drag, and p i t c h i n g moment i n t h e t u n n e l t e s t s . Although t h e r e a r e no methods a v a i l a b l e f o r making accurate c o r r e c t i o n s t o t h e d a t a f o r t h i s ground a general i n d i c a t i o n of t h e magnitude of t h e e f f e c t can be obtained from e f f e c t , previous i n v e s t i g a t i o n s with delta-wing models i n and o u t of t h e presence of t h e ground. (For example, see r e f . 7 . ) The r e s u l t s of t h e s e s t u d i e s would i n d i c a t e t h a t t h e a i r p l a n e i n t h e f u l l - s c a l e t u n n e l t e s t s experienced s l i g h t l y higher values o f l i f t - c u r v e slope and L/D, and SI-ightly more negative values of C, than it would experience o u t of ground e f f e c t .

It appears t h a t any c o r r e c t i o n s f o r t h e e f f e c t of t h e ground on l i f t - c u r v e slope and L/D would be very s m a l l .

The effect of the ground on Cm, however, may be more significant since it may involve corrections as large as Nm = O.OICL or 0.02CL, and such corrections could greatly affect the longitudinal trim characteristics of the airplane as will be discussed subsequently.

Hinge-moment data for basic configuration.- The hinge-moment coefficients of the wing in pitch measured about the pivot (O.5Ock) as determined in test runs at constant dynamic pressure are presented in figure 14. Figure 15 presents hinge-moment data for the power-off condition at measured at various iw = 250 dynamic pressures. As in figure 9, a dashed curve has been superimposed on the The incremental other curves of figure 15 to represent the lgflight condition.

hinge moments between the condition of lgand that of q = 3.07 in figure 15 were

used to correct the basic data of figure 14 to lgconditions, and the corrected

data are presented in figure 16. Presented in figure 17 are the hinge-moment data for three wing pivot positions corrected in this same manner. Since the hinge-moment data in this case are equivalent to the pitching moment of the wing about the pivot point, the stick-free, static longitudinal stability of the air- plane can be determined from the slope of these curves. Although the data show that the airplane is untrimmed for conditions which should be approximately trimmed according to the flight data, stick-free stability is indicated in the moderate lift-coefficient range and instability above and below this range.

F’resented in figure 1 8 are the hinge-moment data for the power-on conditions

of figure 1 6 and the stick forces corresponding to these hinge moments. The data

show no consistent effect of wing incidence. It is believed that the differences in the shape of these curves can probably be attributed to normal scatter of data and that an average curve representative of the measured stick forces for all three wing incidences should be used rather than the individual curves. The large stick forces required for trim are believed to be associated with ground effect and will be discussed in more detail in the following section.

Interpretation of longitudinal data for basic configuration.- In figure 19 and stick forces determined in the tunnel tests are the static margin aCm/aC, compared with values measured in flight. The left-hand plots show tunnel data uncorrected for ground effect whereas the right-hand plots indicate the effect of two assumed values of ground-effect correction: Em = 0 . 0 1 C ~ and A ( & = 0 . 0 2 C ~ .

Force-test data on delta wings in and out of ground effect have indicated that corrections of this order of magnitude may apply in the present case. (See

ref. 7.)

Two plots at the top of figure 19 show the stick-fixed static margin of the airplane when trimmed at various airspeeds as determined from the tunnel data of

figure 11 and from flight data of reference 5. The tunnel data show a slightly

higher value of static margin than the flight-test data and indicate stick-fixed stability over a speed range from about 30 to 4 8 knots. No effect of the ground as a on static margin is shown because the type of ground effect assumed (Em function of CL) changes longitudinal trim but does not change the static margin for trimmed conditions at a given lift coefficient or airspeed. Including the ground-effect correction lowers the trim airspeed for any given flight condition as indicated by the vertical lines in the upper right-hand plot.

I n t h e lower p l o t s of f i g u r e 19, an average s t i c k - f o r c e curve taken from t h e t u n n e l d a t a of f i g u r e 18 i s compared with f l i g h t d a t a taken from reference 5 .

Although t h e f l i g h t - t e s t data shown i n d i c a t e s l i g h t l y s t a b l e s t i c k forces, t h e r e were i n d i c a t i o n s i n t h e f l i g h t t e s t s t h a t t h e a i r p l a n e i n t h e s t i c k - f r e e condition may have been s l i g h t l y unstable o r n e u t r a l l y s t a b l e . The p i l o t reported t h a t t h e t r i m s e t t i n g s and t h a t a i r p l a n e had a tendency t o d r i f t o f f speed at. various e s s e n t i a l l y zero s t i c k f o r c e w a s required t o move t h e s t i c k from full-forward t o f u l l - a f t p o s i t i o n . The f o r c e - t e s t data of f i g u r e 19 i n d i c a t e a s m a l l m o u n t of s t i c k - f r e e s t a b i l i t y from about 35 t o 41 knots and i n d i c a t e s t i c k - f r e e i n s t a b i l - The tunnel data i n d i c a t e very l a r g e p u l l i t y above and below t h i s speed range.

f o r c e s f o r conditions which should be approximately trimmed according t o t h e f l i g h t data. The p u l l f o r c e s a r e made even g r e a t e r when t h e mass unbalance present on t h e a i r p l a n e (wing c e n t e r of g r a v i t y ahead of p i v o t ) i s taken i n t o account. The p l o t a t t h e lower r i g h t shows ;he l a r g e e f f e c t t h a t a c o r r e c t i o n of has on t h e s t i c k - f o r c e c h a r a c t e r i s t i c s . From t h e s e aCm = 0.01CL o r 0 . 0 2 C ~ r e s u l t s it is seen t h a t a c o r r e c t i o n of AC, = 0 . 0 2 C ~ o r g r e a t e r t o t h e tunnel d a t a i s required t o provide t r i m i n t h e speed range from 30 t o 43 knots.

An i n d i c a t i o n of t h e l o n g i t u d i n a l t r i m ' z a p a b i l i t y of t h e a i r p l a n e with v a r i - ous wing incidences and f o r e and aft wing p i v o t p o s i t i o n s i s presented i n f i g - u r e 20. The tunnel d a t a are shown f o r no gmund-effect c o r r e c t i o n and f o r t h e two amounts of c o r r e c t i o n i l l u s t r a t e d i n f i g u r e 19. For t h e wing-incidence range and wing-position t r a v e l a v a i l a b l e , t h e r e appears t o be ample c a p a b i l i t y f o r a t t h e higher speeds b u t only l i m i t e d c a p a b i l i t y f o r trimming i n t h e l o w trimming speed range, unless t h e ground-effect c o r r e c t i o n t u r n s o u t t o be f a i r l y l a r g e .

i f AC, = 0 . 0 2 C ~ proves t o be t h e proper ground-effect c o r r e c t i o n fac- However, t o r , t h e lower p l o t of f i g u r e 20 i n d i c a t e s t h a t t h e a i r p l a n e would have more than enough c o n t r o l power t o t r i m t o t h e s t a l l .

E f f e c t of b o l t r o p e and b a t t e n s . - The r e s u l t s of t u n n e l t e s t s t o evaluate t h e e f f e c t of t r a i l i n g - e d g e b o l t r o p e and b a t t e n s on t h e longitudinal- c h a r a c t e r i s t i c s of t h e a i r p l a n e a r e presented i n f i g u r e s 2 l ( a ) and 21(b).

The d a t a of f i g - u r e 21(a) show t h a t changing t h e b o l t r o p e geometry from t h e s l a c k condition t o the basic condition or t o conditions of reduced boltrope length (up t o 1 inch from t h e b a s i c condition) produced r e l a t i v e l y s m a l l changes i n t h e l i f t , drag, and pitching-moment c h a r a c t e r i s t i c s of t h e airpl.ane. Reducing t h e boltrope length by 4.5 inches produced a l a r g e incremental change i n t h e l o n g i t u d i n a l character- i s t i c s and a l s o a reduction i n s t a t i c l o n g i t u d i n a l s t a b i l i t y .

Presented i n f i g u r e 21(b) are t h e r e s u l t s of t e s t s t o evaluate t h e e f f e c t of t r a i l i n g - e d g e b a t t e n s . These d a t a i n d i c a t e that, t h e b a t t e n arrangements i n v e s t i - gated had l i t t l e e f f e c t on t h e l i f t , drag, and pitching-moment c h a r a c t e r i s t i c s of t h e a i r p l a n e .

I n t h e t u n n e l t e s t s , an attempt w a s made t o eliminate o r t o minimize t h e l a r g e depression i n t h e a f t portion of t h e .wing by doubling t h e l e n g t h of t h e o r i g i n a l b a t t e n s and a l s o by rearranging t h e double-length b a t t e n s . These changes d i d not appear t o improve t h e wing contour c h a r a c t e r i s t i c s appreciably, and t h e depression simply moved forward on t h e wing remaining j u s t ahead of t h e b a t t e n s .

Effect of keel trailing-edge deflection for pitch control.- As part of a general study to explore other methods of providing control for parawing config- - urations, tests were conducted in which the trailing edge of the keel was hinged to deflect upward and downward for pitch control.

The results of these tests (presented in figs. 22(a) and 22(b)) show that a downward deflection of >o from neutral produced relatively large incremental changes in lift, drag, and pitching- moment characteristics but that the effectiveness decreased rapidly for higher deflections. A n upward deflection from neutral produced the desired changes in longitudinal characteristics but caused excessive flutter in the fabric. Tests for this deflection were therefore limited to only high angles of attack where the trailing-edge flutter was less critical. The hinge-moment data of figure 22(b) show that incremental stick forces of about 50 pounds were required for 5 ° of deflection over the angle-of-attack range considered practical (above keel angles of attack of about 20O).

Although the results of figures 22(a) and 22(b) do not indicate a great deal of promise for this particular control system, it is felt that some type of trailing-edge control (such as boltrope, wing-tip, or keel device) could be made to operate efficiently for pitch control. Before such devices can be made prac- tical, however, some means of providing positive Cm,o is necessary to take advantage of an initial downward deflection as a neutral condition since upward deflections Srom the normal wing contour tend to produce excessive trailing-edge flutter. In connection with flutter problems of this type, it was observed in the tests with the present keel control system that a downward deflection of the rear part of the keel eliminated the trailing-edge flutter and furthermore elim- inated the large depression in the aft portion of the wing which had existed throughout the test program. Photographs of the wing with the keel deflected in a downward position are shown in figure 4(a).

Lateral Stability and Control Lateral stability.- The lateral stability characteristics of the airplane are presented in figures 23, 24(a), and 24(b) in terms of the lateral coefficients measured at sideslip angles of 5 ° and -50. Most of the data are plotted against platform angle of attack except for those presented in figure 23 where it was more convenient to compare the effects of wing incidence by plotting the data against lift coefficient. For these tests, the airplane was in the basic config- uration as defined previously in this report. The results indicate that for the power-on, rudder-on case (fig. 24(a)) the airplane had fairly good lateral trim at the lower angles of attack, that is, the rolling and yawing moments are approx- imately symmetrical at sideslip angles of 50 and - 5 O . It should be pointed out that this condition incorporates the 1-inch asymmetric boltrope setting for good lateral trim in powered flight.

It is apparent from the remainder of the

data (figs. 23 and 24) that for all of the power-off conditions and for the power-

on condition with the rudder off, the airplane is out of trim in yaw to the left.

These out-of-trim characteristics are apparently directly attributable to the asymmetric boltrope setting, which was necessary for lateral trim in the power-on, rudder-on condition. It appears, therefore, that the asymmetrical boltrope set- ting was necessary in the basic condition to offset a lateral trim change caused by some induced effect of power on the rudder.

The basic lateral stability data are summarized in figures 25 and 2 6 as the variation with angle of attack of the static lateral stability derivatives Cyp (the direct,ional-stabilityparameter), and (the side-force parameter),

czP

CnP

(the effective-dihedral parameter). 250 (fig. 25), For a wing-incidence angle of the airplane was directionally stable throughout the angle-of-attack range tested except for the rudder-off conditions at the highest angles of attack. Power is shown to increase the directional stability when the rudder is installed, but there is little effect of power on the stabi:Lity with the rudder off and little effect of the rudder on stability with power off. The values of the effective- dihedral parameter are rather large and generally increase with increasing angle of attack. Wing incidence has essentially no effect on the values of the deriva- tives for keel angles.of attack up to about 35O. (See fig. 26. ) The inconsist- encies at higher angles of attack are probably related to stall effects and interferenck effects between the wing and platform.

Wing-bank control.- The variation of the lateral coefficients produced by

banking the wing 5 O and -50 is presented in figures 27 and 28. All of the wing-

bank control tests were made with the rudder installed but undeflected. These data show the same general effects of power on lateral trim noted previously in

the variation of the lateral coefficients at sideslip angles of 50 and -50. The

trim changes due to power are larger for the 28.50 wing-incidence angle than for 25O angle, and this effect is probably related mainly to the higher value of the at the higher wing incidence.

T , The data of figures 27 and 2 8 are presented in figure 2 9 in the form of incremental lateral-force and lateral-moment, coefficients due to banking the wing 5'. In general, the data indicate low rolling effectiveness and favorable yawing moments at the lower lift coefficients. The r o l l effectiveness decreases and becomes negative and the yawing moments become adverse at the higher lift coefficients. With power on (iw = 25O, fig. 2 9 ( b ) ) , the rolling effectiveness is somewhat improved, and the yawing moments are more favorable than with power off.

There appear to be no consistent effects of wing incidence on the magnitude of the control moments, but positive roll control is indicated to higher values of lift coefficient for the wing-incidence angle of 2 5 O . (See fig. 2 9 ( a ) . ) The reason for the low rolling effectiveness of the wing-bank control system is explained in figure 30, which shows f o r both the power-on and power-off condi- tions how the forces and moments from two different sources coEbine to produce the resultant control moments. The bottom plots of figure 30 show a comparison of the measured rolling moments (solid curves, taken from fig. 29(b)) with the moments calculated from the data of previously presented figures (short-dash curves).

The short-dash curve in each case is the sum of the two long-dash curves, which represent the independent and opposite contributions of CL and c z P . The upper long-dash curve (CL sin @ 2) represents the rolling moment produced about the cen- b ter of gravity by banking the wing lift vector over 5 O with the wing pivot at a height z/b above the center of gravity. When the wing banks about an axis par- allel to the wing keel as in the present case, an angle of sideslip of the wing

is produced (sin p = sin % sin @) and this sideslip is adverse, that is, a nose

left sideslip with a right wing bank. This adverse sideslip angle introduces an as indicated adverse rolling moment through the effective-dihedral parameter C Q-3 by the lower long-dash curves. Inasmuch as this adverse rolling moment is almost as large as the favorable rolling moment produced by banking the lift vector, the resultant rolling effectiveness is very small.

The middle plots of figure 50 show a comparison of the measured and calcu- lated yawing moments produced by banking the wing. In this case CL produces adverse yawing-moment increments and p acts through C to produce favorable yawing moments. The wing lift vector produces an adverse yawing moment when it is x/b is negative).

tilted because it acts behind the center of gravity (that is, The favorable yawing moment produced by p results from the fact that the wing is directionally stable positive Cnp) and therefore produces a positive yawing

(

moment when the wing is banked to the right because of the accompanying adverse sideslip angle. A relatively small favorable yawing moment is produced by the displaced drag vector of the wing, and its contribution has been combined with the contribution of Cn in order to simplify this comparison. It is interesting to P note that the loss of wing directional stability at the higher angles of attack accounts for the adverse yawing moments produced by banking the wing as shown in figure 29.

Presented in figure 31 are the incremental hinge-moment coefficients for a wing-bank angle of 5O as measured directly from the wheel force. In addition, hinge-moment data derived from the rolling moment about the wing pivot axis are presented for purposes of comparison. As implied in the preceding discussion, the only rolling moment produced about the wing pivot axis when the wing is banked is that moment due to C of the wing in combination with the adverse sideslip angle resulting from banking the wing; therefore, calculation of this rolling moment provides one means of evaluating the hinge moment due to banking the wing, and it should be equal and opposite to the hinge moment. A second means of obtaining the rolling moment about the wing pivot axis is to transfer the meas- ured roll-control data from the center of gravity to the wing pivot axis. The agreement in the hinge-moment data obtained directly from the wheel force and those calculated from the measured rolling moments is relatively good. The stick forces shown on the right-hand side of figure 31 were computed from the average hinge moments for lg flight conditions over the angle-of-attack range shown. The

roll stick force of about 75 pounds (& = 2 5 O ; Tc = 0.135; % = 2 5 ' ) is in general

agreement with values measured in flight tests.

The roll-control data presented in figures 27 to 29 were obtained in tunnel

tests in which the platform of the airplane was mounted on the tunnel support struts and remained fixed when the wing was banked. As pointed out previously, when the wing banks about an axis parallel to the wing keel as in the present case, an adverse angle of sideslip of the wing is produced (sin p = sin a l . , sin @).

This test condition does not exactly represent what happens in flight when the wing is banked. Actually, when a roll control is applied in flight, the wing and platform momentarily roll and sideslip in opposite directions; the amount each moves is determined by the relative inertia and the aerodynamic moments of the two.

Thus, the true flight condition following the abrupt deflection of the wing-bank control system can be represented by a case somewhere between the two extreme cases of platform fixed at zero bank and sideslip (as in the present tun- nel tests) and wing fixed at zero bank and sideslip (with the platform being deflected in bank and sideslip to provide roll control).

Figure 32 shows how the rolling-, yawing-, and hinge-moment coefficients vary between the two extreme cases. In the plots of the moments against bank angle, the wing-bank angle of 5 O (and platfom-bank angle of 0') represents the test condition used in the tunnel, whereas the wing-bank angle of Oo (and platform-bank angle of - 5 O ) represents the wing-fixed case. For the test con- dition illustrated (iw = 250; a = 00; and 5 ' right wing-bank con-trol), there is a difference of about 2 O in sideslip angle between the wing and platform, with the wing being sideslipped 2 O more nose left than the platform. In the aCz and ACn plots, the horizontal long-dash lines represent the effect of tilting the lift vector, and the short-dash lines represent the moments produced by the wing and platform when they sideslip. The heavy solid lines are the resultant values obtained by adding the long- and short-dash lines. The tunnel-test data point is shown by the symbols at a wing-bank angle of 5 O . The & h plot at the right in figure 32 was constructed in a similar manner, with the long-dash line representing the hinge moments about the pivot produced by the weight of the plat- form and with the short-dash lines representing the aerodynamic moments about the wing pivot produced by the wing and platform when they sideslip. The agreement appears to be satisfactory between the tunnel-test data points and the resultant curves in all three plots of figure 32.

Large effects of sideslip on all the moments are indicated in figure 32 so that the results for the wing-fixed and plat,form-fixed conditions appear to be For example, if the tunnel. tests had been run with the wing quite different.

fixed at zero bank and sideslip, the results would have shown much higher rolling moments but would also have shown adverse yawing moments for the wing-bank control system. Actually, the overall control effectiveness should be about the same for the two cases inasmuch as the yawing moments produce sideslip of the airplane (either favorable or adverse), and this sideslip, acting through the effective- dihedral parameter C z p , produces rolling moments that tend to equalize the net rolling moment acting in the two cases. Perhaps the best indication of the net roll-control effectiveness shown in the plots of figure 32 is the rolling moment for the case where the yawing moment is zero. This condition occurs at the point where the wing is banked 3 . 5 O right and the platform l.5O left. This proportion of initial wing bank to platform bank also appears generally reasonable on the basis of estimated relative inertias and the aerodynamic moments of the wing and platform as indicated by the results of one-degree-of-freedom, initial-response calculations. Inasmuch as the net rolling moment (rolling moment for the case when yawing moment is zero) appears to be of the most significance in evaluating control effectiveness in the wing-bank control system, an equation has been devel- oped in the appendix to facilitate calculation of the net rolling moment when only the most fundamental aerodynamic characteristics are known.

Hinged wing-tip controls.- In view of the inadequacies of the existing wing- bank control system (low roll-control effectiveness and high hinge moments), tests were conducted to evaluate the effectiveness of two alternate roll-control systems.

These alternate control systems were described previously as wing-tip controls A

and B and are illustrated in figure 3 and shown in photographs in figure 4. In

both tests the wing-bank system was rigidly locked at zero deflection.

The results of the tests made to evaluate the effectiveness of these two alternate control systems are presented in figure 33, and a comparison of these data shows much better rolling moments for control A than for control B. In fact, no consistent roll effectiveness was obtained with control B. Both systems appear to have adverse yawing moments when the control is deflected in a direction to produce positive roll control, and these adverse yawing moments generally become larger with increasing angle of attack. The results presented in this comparison are for deflection of the control devices on the left wing tip only. In order to see what could be done to minimize the adverse yaw characteristics, the data of figure 33(a) have been used to prepare figure 34 for the case of differentially operated controls on both wing tips. Data are presented only for control A inasmuch as control B was found lacking in roll effectiveness. Data are shown for several neutral settings of the hinged wing tip. It is apparent that an inward neutral setting tends to reduce the adverse yawing moments due to control deflection at the higher angles of attack such that with both tips initially set 5 O inward, the adverse yawing moments due to control deflection are essentially at reduced to zero throughout the angle-of-attack range. For this case, the r o l l i n g moments with a wing-tip deflection of only 50 or - 5 O (nC1 = 0.011 to 0.015) are appreciably larger than values obtained with the wing-bank control system, and also, roll-control effectiveness is maintained at the higher angles of attack.

The hinge moments are also appreciably lower than those obtained with the wing- bank system; and, for the zero-angle-of-attackcondition previously cited, they would result in wheel forces about half as great as those experienced in flight tests.

Comparison of roll-control systems.- The wing-bank-control characteristics of the airplane are summarized in figure 35. Also presented in this figure for comparison are data for hinged wing-tip control system A. In addition, estimated control characteristics are presented for this wing-bank control system with nega- tive geometric dihedral of the wing added.

In the left plot of figure 35, incremental rolling-moment coefficient is plotted against incremental roll hinge-moment coefficient & ! h . The horizontal dashed line represents the value of &!I required to produce a value of pb/2V

of 0.09, based on the relationship - pb = - c 2 and an estimated value of the

2v c, LP damplng-in-roll parameter C of -0.15. The value of pb/2V of 0.09 is the IP minimum value specified in the handling-qualities requirements for a light liaison airplane. This criterion is presented here merely to establish a reference for purposes of comparison and is not intended to imply that a value of pb/2V of 0.09 is a valid requirement for parawing applications. For recovery-system appli- cations, a much smaller value may well prove to be acceptable; whereas, for utility-airplane applications (which may involve flight at very low speeds in confined areas), an even larger value than 0.09 may be required. In any event, considerably more research and flight experience will be required to establish 1 6 t h e proper c r i t e r i a f o r t h e various applicati.ons envisioned f o r t h e parawing.

~ s c a l e a r e t h e hinge-moment c o e f f i c i e n t s t h a t Also indicated along t h e ACh I correspond t o s t i c k f o r c e s of 50 t o 100 pouncls. The s o l i d c i r c l e a t t h e lower r i g h t , representing t h e wing-bank c o n t r o l system i n s t a l l e d on t h e a i r p l a n e ( n e t r o l l c o n t r o l as represented by t h e nCn = 0 c a s e ) , shows t h a t banking t h e wing 5 ' r e q u i r e s about 70 pounds of s t i c k f o r c e and only produces about one-third of t h e r o l l i n g e f f e c t i v e n e s s required by t h e pb/2V = 0.09 c r i t e r i o n . Calculations by using a negative geometric d i h e d r a l i n d i c a t e (open symbol) t h a t reducing angle of t h e wing of 180 would decrease t h e s t i c k f o r c e t o about 45 pounds and increase t h e e f f e c t i v e n e s s t o about one-half of t h e c r i t e r i o n value. The wing- t i p c o n t r o l system appears t o be q u i t e e f f e c t i v e i n t h a t a wing-'tip-control d e f l e c t i o n of approximately 7 O produces a value of pb/2V of 0.09 with appreci- ably lower s t i c k f o r c e than t h a t of t h e wing-bank system ( 3 0 t o 45 pounds, depending on n e u t r a l s e t t i n g of t h e t i p s ) .

The right-hand p l o t of f i g u r e 35 shows t h e incremental yawing moments pro- , duced by t h e various r o l l - c o n t r o l arrangements. The yawing moment i s zero f o r ~ t h e wing-bank c o n t r o l system because t h i s condition w a s s p e c i f i c a l l y s e l e c t e d from f i g u r e 32 t o give t h e b e s t i n d i c a t i o n of n e t r o l l e f f e c t i v e n e s s .

Although t h e yawing moments with t h e hinged wing-tip c o n t r o l appear t o be q u i t e small f o r t h e angle-of-attack conditicn represented i n t h i s f i g u r e (% = Oo; iw = 25'), adverse yawing moments of considerable magnitude would be encountered a t higher angles of a t t a c k unless an i n i t i a l . inward n e u t r a l s e t t i n g of t h e t i p s w a s used. An inward n e u t r a l s e t t i n g of about 5 O i n combination with about k7O

i

d i f f e r e n t i a l d e f l e c t i o n of t h e wing t i p s should provide roll c o n t r o l of s u f f i c i e n t ' magnitude t o meet t h e pb/2V = 0.09 criterj.on with no adverse yaw throughout a k e e l angle-of-attack range from 20° t o 3 5 O .

Rudder c o n t r o l . - The rudder-effectiveness d a t a a r e presented i n f i g u r e 36 i n t h e form of side-force, yawing-moment, and rolling-moment c o e f f i c i e n t s . The sons t h e p i l o t made extensive use of t h e r.udder i n f l y i n g t h e a i r p l a n e . It should be pointed out, however, t h a t t h e r o l l response obtained through t h i s i n d i r e c t c o n t r o l i s s u b j e c t t o appreciable time l a g and o t h e r dynamic e f f e c t s , and t h e r e - SUMMARY OF RESULTS The results of the full-scale tunnel investigation of the performance and static stability and control characteristics of a flexible-wing manned test vehicle are summarized as follows: 1. The tunnel tests showed that the maximum lift coefficient of the airplane occurred at a keel angle of attack of 42O and was 1.24 with power off and 1.33 with power on. The maximum lift-drag ratio was about 5.5.

2. With stick fixed, the airplane had about neutral static longitudinal sta- bility at keel angles,of attack below 2 0 ° , a moderate degree of stability from 2 0 ' to 3507 and longitudinal instability, or pitch-up, from 3 5 O to 42O. At an angle of attack of the keel of 420 the airplane again became stable. With the stick free, the longitudinal stability was generally worse with the airplane being unstable at the lower angles of attack, about neutrally stable in the interme- diate range, and unstable at the higher angles of attack.

3. The airplane, in general, was directionally stable and had positive effective dihedral throughout the angle-of-attack range investigated.

4 . The lateral control provided by banking the wing did not appear to be

satisfactory because of inadequate rolling moments and excessively high stick forces. This result is in agreement with flight-test results.

5. Analysis of the factors contributing to the low rolling effectiveness obtained by banking the wing indicated that the use of negative geometric dihe- dral of the wing to reduce the high values of positive effective dihedral may be a relatively simple means of improving the effectiveness of this type of roll- control system.

6. Analysis of the tunnel data indicated that the rudder was generally a better roll-control device with power on (since the rudder is in the slipstream The rudder provides of the pusher propeller) than the wing-bank control system.

roll control in an indirect manner by sideslipping the airplane and making use of the large value of effective dihedral (rolling moment due to sideslip).

7. The hinged wing-tip control device tested on the airplane appeared

promising in that it provided higher rolling effectiveness with lower stick forces than that of the wing-bank control system provided on the airplane.

Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., June 4, 1963.

1 8 APPENDIX DERIVATION O F AN EXPFBSSION FOR CALCULATING THE NET ROLLING M O M E N T PRODUCED BY THE W I N G B A N K OR CENTER- OF- GRAVITY- SHIliT CONTROL SYSTEM Sketch 1 shows t h e r e l a t i o n s h i p of t h e :Lift, drag, and r e s u l t a n t - f o r c e vec- t o r s of a parawing configuration i n trimmed, l e v e l f l i g h t . For such a condition, t h e r e s u l t a n t - f o r c e vector must pass through t h e c e n t e r of g r a v i t y and, therefore, from t h e geom- z/b - L

e t r y i n t h i s case it can be sho-m t h a t - - -

x/b D' When t h e wing i s banked f o r r o l l c o n t r o l , t h e l i f t vector i s t i l t e d and has a l a t e r a l component as shown i n sketch 2. For t h e condi- CL s i n $8

x

t i o n shown, it i s seen t h a t t h e l a t e r a l compo- nent of t h e l i f t vector produces a r o l l i n g T + @ moment about t h e c e n t e r of g r a v i t y through t h e I z

-++ x/b I-

arm z/b (thus, AC, = - CL s i n fl . Also, s i n c e

) \ b

w

t h i s vector component i s behind t h e c e n t e r of gravity, i t . produces an adverse yawing moment Sketch 1 about t h e c e n t e r of g r a v i t y through t h e arm

x/b (thus, A C , = g X CL s i n #). I n order t o

deternine t h e n e t r o l l i n g moment i n t h i s case ( t h e r o l l i n g moment for zero yawing moment ), it i s necessary t o t a k e i n t o account t h e equi- librium s i d e s l i p condition where: L

I

From t h i s e s t a b l i s h e d value of p, t h e incremental r o l l i n g moment introduced through t h e e f f e c t i v e - d i h e d r a l parameter C z p can then

0-i-

be determined as Sketch 2 For normal conditions, the directional-stability and effective-dihedral parameters of parawings are positive (+Cnp, -Clp) and, therefore, in a right determined from equation (A2) would be adverse, o r wing bank the value of ACZ negative, and would subtract from the favorable rolling moment produced by the lift vector. The net incremental rolling moment produced in this case can there- fore be written as or z/b

Factoring out C sin # and substituting L/D for - gives

5 L x /b

czB 1

is called the rolling-effectiveness factor and is The term 1 + --

CnP L/D convenient for estimating very readily the percentage of roiling effectiveness that is actually available for a configuration employing the wing-bank or center- of-gravity-shift control system. When this factor approaches 1.0 the loss of roll-control effectiveness is minimized, whereas, when this factor approaches the net roll-control effectiveness also approaches 0. For configurations having

high negative values of - cZp and low values of L/D, the rolling-effectiveness

CnB term becomes small and therefore the net rolling moment produced in such cases

cZP

is reduced. Configuration changes which would tend to reduce the ratio -

CnP and increase are obviously desirable from the standpoint of net roll- L/D control effectiveness. One of the most obvious improvements would be to reduce the derivative C by introducing negative geometric dihedral of the wing.

(Reduction of C i p by reducing z/b would defeat the purpose because the pri- mary roll-control term (E CL sin 6 ) would also be reduced.

For purposes of comparison, values of calculated from equation ( A 5 ) AC1,net

are presented in figure 37. These calculations were made for the rudder-on

configuration, with power off and on, by using the measured force-test data cor- responding to these conditions. The calculated data show the loss in rolling effectiveness at the higher angles of attack indicated previously by the meas- ured data of figure 29 although, as expected, the calculations generally show higher rolling effectiveness than the measured data in the lower angle-of-attack range. Good agreement is shown between values of AC,,net calculated from equa-

(A5) and those determined graphically in figure 32 for (iw = 25') % = 0 '

tion since in both cases the rolling moments were calcylated on the basis of zero yawing moment.

REFERENCES 1. Rogallo, Francis M., Lowry, John G., Croom, Delwin R., and Taylor, Robert T.: Preliminary Investigation of a Paraglider. NASA TN D-443, 1960.

2. Naeseth, Rodger L.: An Exploratory Study of a Parawing as a High-Lift Device for Aircraft. NASA TN e-629, 1960.

3. Hewes, Donald E.: Free-Flight Investigation of Radio-Controlled Models With Parawings. NASA TN D-927, 1961.

4. Johnson, Joseph L., Jr. : Low-Speed Wind-Tunnel Investigation to Determine

the Flight Characteristics of a Model of a Parawing Utility Vehicle. NASA I TN D-1255, 1962.

I 5. Landgraf, F., Everett, W. L., Burich, J. H. : Flexible-Wing Manned Test Vehicle.

TCREC Tech. Rep. 62-25 (Ryan Rep. 61~131~), U.S. Army Transportation Res.

Command (Fort Eustis, Va.), June 25, 1962.

6. &France, Smith J.: The N.A.C.A. Full-scale Wind Tunnel. NACA Rep. 459, 1933.

7 . Scallion, William I.: The Effect of Ground on the Low-Speed Aerodynamic, Control, and Control Hinge-Moment Characteristics of a Delta-Wing-Fuselage Model With Trailing-Edge Controls. NACA RM L54H03, 1954.

TABIX I . . CRARACTERISTICS OF THE AIRPLANE

Airplane weight. l b . . . . . . . . . . . . . . . . . .

. . . . . . . . . . 1 . 840

I Keel and leading-edge length. f t . . . . . . . . . . .

. . . . . . . . . . 2 8

Leading-edge sweep angle ( f l a t plan geometry). deg . . . . . . . . . . . . 45

Span (based on 4 5 O leading-edge sweep). f t . . . . . . . . . . . . . . . . 39.6

Wing area ( f l a t plan geometry. 4 5 O leading-edge sweep). sq f t . . . . . . . 555

Leading-edge sweep angle ( f l i g h t condition). deg . . . . . . . . . . . . . 50

I Wing aspect r a t i o . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . 2.82

Eugine power. hp . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . 180

Propeller diameter. f t . . . . . . . . . . . . . . . .

. . . . . . . . . . 6

Rudder dimensions: Area. s q f t . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . 13.4

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

. . . . . . . . . . 4.75

span. f t . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . 3.75

Chord. ft

Aspect r a t i o . . . . . . . . . . . . . . . . . . . . I

. . . . . . . . . . 1.70

C

A’\

(b) Flat plan geometry of wing showing battens and trailing-edge scallop.

Figure 1.- Concluded.

62- 63l L- .

tunnel scale - .

full details Langley in testing support-system force showing for view mounted front uarter airplane -q of Three (a ) Photographs .- Figure f\) 0\ t--- C\J \() I C\J \() ~

c

.....

+' () QJ en +' en QJ +' +' QJ @ <-i -d QJ

e

-g . ....

<-i al () s:!

0 0 ""' (.)

Q 0 I .....

C\J +' oj QJ <-i QJ h ~ . ....

bD Q r...

·S ..t:1 en :> QJ .....

> +' s:: t: ..t:1 ~ k h d v

i

I pc\

c

.d c r(

i

a .I see detail at left Side view Perspective view of wing showing control in outward position ( b ) Wing-tip c o n t r o l A.

- -r Wing trnillng edge

- - _jJL - - - * A - - - "A ..

- Control arm Plan view ,-Control arm - support members : l e e detail at left Sida view Perspective view of wing showing control in downward position ( c ) Wing-tip c o n t r o l B.

Figure 3.- Concluded.

( a ) K ee l trailing - edge control system .

L- 63 -3l62 F igure 4 .- Al ternative control systems used in the investigation .

<>: .,j .-I gJ M s:: .., 'M .., s:: s:: t> 0 (.)

P< I 'M .., .:j- ~ OJ a 'M ~ ~ ....

Ii.

~ _ ..

p::i -d (lJ rl '(j >< ::l ..., rl s:: tJ 0 C tJ 0 U P< 0'-; ..., ..;.

I t10 (lJ s:: 0'-; ;l: ~ 0'-; tJ r«

I

-1 (a ) Basic configuration ; ~ = 5° . L- 63 - 3l63 F igure 5 .- Ph otograph of airplane showing wing contour irregularities . Power off ; iw = 25°.

(b) Boltrope slack, batte Qs in; ~ = 0° .

L-63-3l64 F igure 5 .- Continued.

(c ) Boltrope slack, battens out ; ~ = 0° .

L- 63 -3l65 Figure 5 .- Concluded .

D d d d rEnd of t e s t I l O O r - a F - I \ /- 0 b- s o s Beginning of t e s t t : 2 100- a a , - 0 , 50’ 1oc- Boltrope slack, b a t t e n s i n 50- 5 0 - l O O L , Boltrcpe s l a c k , battens o u t 40- 20 - 0 - /-7 20- 40- 1 I ( b ) uk = 2 6 ’ .

Figure 6.- Time h i s t o r y of c o n t r o l forces measured i n tunnel t e s t s . i, = 25’; Tc = 0; 9 = 3.07 n/sq f t .

2 2 . 5 2 5 . 0 28.5 . 1

- . 1

1 . 4 1 . 2 1 .o .8 . 6 .4 Figure 7.- Static longitudinal characteristics of airplane.

= 0; lb/sq ft.

E U .02 .01 -.01 Crn -.02 -.03 -.04 -.os I .4 I .2 I .o .8 CL C D .6 .4 .2 -20 -15 -10 -5 0 5 I O 1 5 .02 .01 0 -.O I -.02 -.03 -.04 -.05 a p t d e g C r n Figure 9.- Effect of dynamic pressure on sttitic longitudinal characteristics of airplane.

i , = 250; T , = 0.

.o I -.o I Cm -.02 -.03 -.04 -.05 I .4 I .2 I .o .8 C L C D .6 .4 .2 -15 -10 -5 0 5 1 0 1 5 .OI 0 - . 0 1 -.02 -.03 -.04 -.05 a p t d e g Crn Figure 10.- Comparison of static longitudinal characteristics of airplane for constant loading and constant dynamic pressure.

i , = 2 5 O ; Tc = 0.

o c t- o o c P C E u

50 40 35 30 27 Velocity, kn

1 b 4

1 b 2

1 .o

. 8

CL

.6

b 4

b 2

Figure 12.- Comparison of lift characteristics as measured in wind-tunnel tests and flight tests of airplane. Airspeeds were estimated for 1,840 pounds.

-15 -10 -5 0 5

10 15 2 0 . 1 0 - . 1

Figure 13.- Effect of wing pivot p o s i t i o n on s t . a t i c l o n g i t u d i n a l charactei.istics of a i r p l a n e f o r 1 g condition.

Tc = 0.

1 , = 22.50 TC

--

. 0 1 C h

- 001

ch

- . 0 1

- . 0 2

. 0 1 ' h - . 0 1 -002 1 . 0 1 . 2 104 0 02 04 06 08 cL Figure 14.- Hinge-moment characteristics of wing in pitch measured about pivot (O.5Ock) at constant dynamic pressure.

r l cv

0 0 I I

c

u

--

.120

- . 0 1

%

- . 0 2 .01 ' h - . 0 1

- 002

. 0 1

- . 0 1

- . 0 2 16.- Hinge-moment characteristics of King In pitch corrected for l g condition.

Figure d d d IJ u <D d L ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! I ! ! ! ! ! ! ! ! ! ? ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! ! !!!-11"1' c 3 d 0 P I 0 0 I I I

s

Figure 18.- Hinge-moment characteristics of wing in pitch (data from fig. 16) and corresponding pitch stick forces for three different wine incidences. Power on.

e, a , c k

L L n LA-

0 ?I

3 rl rl

L c I I - Ground-effect -.

\ correctlon htiating ning- incidence range wing pivot P -3 posltlon li r( \ 2 5 0.512Ck - 7 7 -~ -- -- M i c (

‘&LA?

20 F

.02CL I

16 I

0 10 20 30 40 50 60 .35 Figure 20.- Estimated angle of incidence required f o r p i t c h t r i m with s e v e r a l d i f f e r e n t values of ground e f f e c t included. Parer on.

Boltrope length Slack Basic condi1;ion 0.50-inch docrease in length A 1.00-inch decrease in length h 4.50-inch decrease i n length .1

- . 1

1 . 2 1 .o .4 .2 -10 -5 0 5 10 15

01 0 - 0 1

ap, deg

cm

( a ) E f f e c t of boltrope length.

Figure 21.- E f f e c t of wing trailing-edge modifications on s t a t i c 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 a i r p l a n e f o r l g condition. Tc = 0.

Battens Original 1 2 3 4 5 c,

- . 1

1 . 4 1 . 2 1 .o . 8 CL CD . 6 .4 . 2

-

(b) Effect of batten modification.

Figure 21.- Concluded.

. 1 Cm 0 - . l 1.4 1 . 2 1 .o . 8 CL . 6 .4 . 2 -.1 -10 -5 0 5 10 15 L. del2

7' --- -m

(a) P i t c h e f f e c t i v e n e s s Figure 22.- S t a t i c 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 airp1,ane with k e e l t r a i l i n g edge d e f l e c t e d f o r p i t c h c o n t r o l f o r l g condition.

Tc = 0; iw = 25'.

Trailing-edge keel deflectlon, deg

- .005

- ,010

- 0015

- . O m

-e025 rl r( s 50

1, -10 -5 0 5 10 15

(b) Longitudinal hinge-moment and stick-force characteristics.

Figure 22.- Concluded.

CY Cn .2 . 4 .6

.0 I .o I .2 I .4

Figure 23.- Effect of wing incidence on variation of lateral coefficients with lift coefficient due Tc = 0 ; rudder on; 6 , = Oo; q = 3.07 to 3.77 lb/sq ft.

to sideslip angles of 5 ' and -5'.

CY -10 -5 0 5 10 15 QP, deg ( a ) Rudder on; 6 , = 0'.

Figure 24.- E f f e c t of power ' on v a r i a t i o n of l a t e r a l c o e f f i c i e n t s with platform angle o f a t t a c k due t o s i d e s l i p angles of 5 ' and -5'. i , = 25O; q = 5.07 lb/sq f t .

CY Cn ( b ) Rudder o f f .

Figure 24.- Concluded.

T , Rudder 0 on

---

.135 on

- - ---

0 Off --- .135 o f f .02

. 01

- . 0 1

. O 0 4 .002

- .002

- .002

czP - . O M

- -006

-10 -5 0 5 10 15 aPS deg Lane. iw = Fj Lgure Ef 'fect of power and rudder on static-lateral-stability parameters of airpl 9 = 3.07 to 3.77 ib/sq ft.

2 2 . 5

---

25.0

-

2 8 . 5 .02

. 01

cyP

- . 0 1

.002 C "P

- ,002

-.m

- .006

Figure 26.- Effect of wing incidence on static-lateral-stability parameters of airplane.

T , = 0; rudder on; q = 3.07 lb/sq ft.

C Y -.VL .8 1.0 1.2 1.4 0 .2 .4 .6 CL .ent due Figure 27.- E f f e c t of ' wing incidence on v a r i a t i o n O f l a t e r a l c o e f f i c i e n t s with l i f t c o e f f i c j t o wing-bank angles of 'jo and -5'. TC = 0; 9 = 3.07 lb/sq f t .

5 0 5 .135 -- - 5 0

$-- - - 5 .135

.2 . 1 CY -.l

- . 2

. 0 2 . 0 1 Cn

- . 0 1

. 0 1 CZ 0 - . 0 1 -10 -5 0 5 10 15 QP’ deg ( a ) iw = 250.

Figure 28.- E f f e c t of power on l a t e r a l c o e f f i c i e n t s aue t o wing-bank angles of 5 O a n d - 5 O .

9 = 3.07 lb/sq f t .

9, deg TO 0 5 0 0 5 .i60 d-- - 6 a-- - 5 . 1 6 0 .e . 1 -.l -.2 r 0 2 . 0 1

- . 0 1

.02 . 0 1 - . 0 1 -.02 -10 -5 0 5 10 15 QP, deg (b) i , = 28.5O.

Figure 28. - Concluded.

- . 0 1 =t -001 - . 0 2 C L (a) Effect of wing incidence; 0 .

T C Figure 29.- Incremental l a t e r a l force and mcments produced by a wing-bank angle o f 5'.

q = 3.07 lb/sq f t .

---

.135 . 2 . 1

. 01

. 0 1 . 0 1 . 0 1

.o -5 0 5 10 15 -1

( b ) Effect of power; i , = 25'.

Figure 29.- Concluded.

u) U a ti I n

t-

c__ l m m c +, - g - Y c s 5 : Lo U ti rl ?

i

4 Q Q 4 4

-!n --Y) - I n -!n

8 10

A 15 Ch - . 0 1 ch -.02 - .02 . 1 CY CY 0 -.l .01 Cn C n o -.01 -.01 - .02 -.02 .02 .02 . 0 1 .01 CZ cz 0 - . 0 1 - .01

I- t f;T lj y : t

- . 0 2 - ."I 15 -10 -5 0 5 10 -10 -5 0 5 10 ap, deg (a) Wing-tip control A.

( b ) Wing-tip control B.

F i p e 33.- Comparison of lateral control characteristics of two alternate roll-control devices on left wing tip only. q = 3.07 lb/sq ft.

N e u t r a l p o s i t l o n , Wing-tip d e f l e c t l o n , d e g r i g h t t l p l e f t t i p 5.0 inward 5.0 outward

~-

10.0 inward 10.0 outward - - - - - - 5.0 inward 5.0 outward 5.0 lnward _ _ _ - - 7.5 outward 2.5 lnward 7.5 inward .010 “ h .005

. 01

*Cn

- . 0 1

- .02

.03 .02 . 0 1 a - -5 0 5 10 ap’ deg re ..- Inc 1 remental l a t e r a l - c o n t r o l moments obtained with d i f f e r e n t i 3 4 . a1 d e f l e c t i o n of hinged wing t i p s f o r s e v e r a l d i f f e r e n t n e u t r a l s e t t i n g s .

i= .rl a

.$

.rl d-.

rd k k do o a l .rl 4 4 ffi h P v P- ?

N 0 In 0- 0 d L o r(

0 w 8 a

9 9 9 9 9 N V a . 1

- . 1

.02 0 01

. 01

, . 0 1 -I 20 -15 -10 -5 0 5 10 15 Right control Left control Figure 36.- Effect of power on rudder control..

% = 0'; iw = 25'; q = 15.07 lb/sq ft.

---

. 0 1

u

a

+.,

- . 0 1

- . 0 2

-10 -5 0 6 10 15 20

Figure 37.- Calculated incremental net rolling-moment coefficient, produced by 5 ' of wing bank.

i , = 2 5 O .

72 NASA-Langley, 1963 L- 3373