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N A S A TECHNICAL NOTE

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WIND-TUNNEL INVESTIGATION

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NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON D. C. AUGUST 1977 1I

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1. Report No. 2. Government Accession No.

NASA TN D-8475 ' .I 4. Title and Subtitle WIND-TUNNEL INVESTIGATION OF A VARIABLE CAMBER AND TWIST WING

7. Author(s)

James C. Ferris

9. Performing Organization Name and Address NASA Langley Research Center Hampton, VA 23665

2. Sponsoring Agency Name and AddressNational Aeronautics and Space Administration Washington, DC 20546

5. Supplementary Notes

6. Abstract

TECH LIBRARY KAFB, NM

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5. Report Date August 19771I 6. Performing Organization Coda

8. Performing Organization Report Nc

L-11357

I10. Work Unit No.505-11-16-08

11. Contract or Grant No.

I

.13. Type of Report and Period CoversTechnical Note

j -~ 14. Sponsoring Agency Code

I

. .

A wind-tunnel investigation was made to determine the longitudinal aerodynamic characteristics of a.35O swept, variable camber and twist semispan wing in the pre ence of a body. The variable camber and twist were incorporated to allow a near optimum lift distribution over the wing for both the cruise condition and the highlift conditions for maneuverability. The wing incorporated movable leading-edge segments whose swept hinge lines provided maximum camber variations at the outboar leading edge and movable trailing-edge segments whose swept hinge lines providedmaximum camber variations near the inboard trailing edge. The model was investi­gated at Mach numbers of 0.60, 0.80,and 0.90 through an angle-of-attack range frc Oo to IOo or buffet onset.

18. Distribution Statement

Unclassified - Unlimited

Subject Category

19. Security Classif. (of this report) I 20. Security Classif. (of this page) 21. NO. of pager 22. Rice'

Unclassified Unclassified 79 $5 .OO

WIND-TUNNEL INVESTIGATION OF A VARIABLE CAMBER AND TWIST WING

James C. Ferris Langley Research Center

SUMMARY

A wind-tunnel investigation was made to determine the longitudinal aerody­namic characteristics of a 35O swept, variable camber and twist semispan wing in the presence of a body. The variable camber and twist were incorporated to allow a near optimum lift distribution over the wing for both the cruise condition and the high lift conditions for maneuverability. The wing incorporated movable leading-edge segments whose swept hinge lines provided maximum camber variations at the outboard leading edge and movable trailing-edge segments whose swepthinge lines provided maximum camber variations near the inboard trailing edge.The angle of the segments could be varied with an internal mechanism; thus, tf;e camber and twist could be optimized for various lift coefficients.

The model was investigated at Mach numbers of 0.60, 0.80, and 0.90 at Reynolds numbers based on the mean geometric chord of 6.2 x IO6, 7.4 x IO6, and 7.6 x IO6, respectively. The angle of attack was varied from 00 to approximately 100 or to buffet onset.

Leading-edge camber simulating elliptical camber, circular camber, and sim­plified camber of simple hinge and double hinge line configurations was investi­gated with uncambered trailing edges. The trailing-edge camber investigationincluded elliptical leading-edge camber with three variations of elliptical trailing-edge camber and one extensive trailing-edge camber configuration.

The results of the inve.stigation showed that, when properly incorporated, variable camber and twist can effectively reduce the drag of a thin low-aspect­ratio wing over a wide range of lift coefficients. Compared to the uncambered wing, the wing with leading-edge camber was effective in reducing the drag in the lift-coefficient range from 0 to 0.4 while the wing with trailing-edge cam­ber is required at lift coefficients greater than 0.5. The increase in buffet-free lift coefficient as a result of variable camber and twist relative to that of the basic uncambered and untwisted wing varied from 68 to 102 percent over the Mach number range of the investigation.

INTRODUCTION

The performance of most moderate aspect ratio, thin swept wing maneuveringaircraft is significantly degraded at high lift coefficients at high subsonic Mach numbers because of shock-induced boundary-layer separation and, at higherangles of attack, because of leading-edge separation and wing stall. The result­ing degradation in handling qualities significantly reduces the combat effective­ness of these airplanes. There are several approaches to the leading-edge stall problem, including leading-edge flaps, slats, and control of the boundary layer

by suction or blowing. These approaches, along with trailing-edge flaps, have been used effectively for low-speed landing and take-off and, at higher subsonic speeds, to increase the maximum usable lift coefficient.

Camber has always been used to provide lift with low drag at cruise (a more recent example is given in ref. 1 ) and is used to increase the buffet-free lift coefficient (ref. 2). Leading-edge and trailing-edge flaps have been used to increase the camber for maneuvering fighters at high subsonic speeds in order to maximize the usable lift coefficient. (See refs. 3 to 8.) Twist is used to pro­vide the desired span load distribution at cruise (ref. 9). While considerable data are available on varying the camber with flaps, very little data are avail­able on variable twist or washout.

Low-thickness-ratio wings incorporating variable camber and twist appear to offer higher performance for fixed-wing-planform fighter aircraft since the cam­ber can be reduced or reflexed for the supersonic mission and increased to pro­vide high lift coefficients required for transonic and subsonic maneuverability.

The objective of this study was to determine the aerodynamic characteris­tics of a low-thickness-ratio fighter wing. To accomplish the variable camber function, the wlng planform had four leading-edge segments and four trailing-edge segments, all with spanwise hinge lines. To accomplish the variable twist, the hinge lines of the leading-edge segments were parallel and swept more than the leading edge, and the trailing-edge segments were parallel and swept more than the trailing edge. For this hinge line layout, twist is increased as'camber is increased and, thus, the variable camber and twist concept. The purpose of this paper is to present force and moment data on this configuration at Mach numbers of 0.60, 0 .80 , and 0.90.

SYMBOLS

The longitudinal results are referred to the stability axis system. The origin of the stability axes is at the moment reference center, located at 25 percent of the reference length E and 147.57 cm (58.09 in.) from the fuse­lage apex.

Values are given in both SI and U.S. Customary Units. The measurements and calculations were made in U.S. Customary Units.

exposed semispan, 106.68 cm (42 in.)

Cacc wing-tip accelerometer output, g units

b/2

drag coefficient, Drag9s

lift coefficient, Lift 9s

pitching-moment coefficient, Pitching moment qSE

chord, cm (in.)

mean geometric chord, 59.59 cm (23.46 in.

incidence of airfoil section, deg

leading edge

lift-drag ratio

maximum lift-drag ratio

free-stream Mach number

free-stream dynamic pressure, Pa (lbf/ft2)

Reynolds number, based on -5

exposed wing area, 0.599 m2 (6.447 ft2)

trailing edge

longitudinal distance, cm (in.)

spanwise distance from root chord, cm (in.)

ordinate normal to airfoil reference line, cm (in.)

angle of attack, deg

incremental deflection angle of variable camber segments, deg (n denotes segment number, see fig. 4(b))

wing semispan station, 2y/(b/2)

MODEL DESCRIPTION

The general arrangement of the semispan model is shown in figure l(a). A spacer was installed between the fuselage and the wind-tunnel wall to place the entire semispan model in the free-stream flow of the tunnel. Details of the seg­ments of the wing planform are shown in figure l(b), and cross sections of'the fuselage are shown in figure l ( c ) . Photographs of the model installed in the

3

Langley 8-foot transonic pressure tunnel are presented as figure 2. The cross-sectional area distribution presented in figure 3 is for a three-dimensional full-span model having the same geometric characteristics as the semispan model. This area distribution was designed, according to the area-rule concept, for a Mach number of 1.0 at zero lift, and the fineness ratio was 7.52, which is repre­sentative of current fighter airplanes.

Only the wing was attached to the balance, and the nonmetric fuselage was separated from the wing by a small gap (approximately 0.32 cm (1/8in.) wide) around the airfoil near the wing root.

The wing consisted of an NACA 658005 airfoil at the root and an NACA 65A004 airfoil at the tip with a linear variation between them (no camber or twist in the basic wing). In addition, the airfoils were modified to have finite thick­ness at the trailing edge to improve the structural characteristics and to pro­vide increased thickness in the aft part of the airfoil for the variable camber and twist mechanism. This trailing-edge modification was small and varied lin­early from 0.00375~at the root to 0.00300~at the tip. Coordinates of the air­foil sections are given in table I.

The hinge lines of the leading-edge segments were swept 40.1O (5.10 greater than the leading edge) and were parallel, thus providing a cylindrical camber. The hinge lines of the trailing-edge segments were swept 25.4O (11.4O greater than the trailing edge), and they were also parallel, providing cylindrical cam­ber. As camber is applied to a wing with this arrangement, twist or washout also results.

The wing planform with the segments numbered is shown in figure 4(a); the semispan locations selected for the camber and twist computations and the cross section near the root and tip of the wing planform are also shown.

The incremental deflection angles of the variable camber segments are shown in the schematic drawing in figure 4(b). As shown in the small drawing at the top, all of the leading-edge segment deflection angles are measured in a plane normal to the hinge lines of the leading-edge segments, and the trailing-edge segment deflection angles are measured in a plane normal to the hinge lines of the trailing-edge segments. Each of the segments could be deflected through an angle range from 2O up to 1 2 O down. The incremental deflection angles (in deg) of the leading-edge segments are referenced to their aft adjacent segments (that is, for segment 1 86, is referenced to segment 2 and for segment 2 A62 is referenced to segment 3, etc.). The incremental deflection angles of the trailing-edge segments are referenced to their forward adjacent segments (that is, for segment 5 A65 is referenced to the main wing box and for segment 6 A66 is referenced to segment 5, etc.1. All of the angles from leading edge to trailingedge at any wing station represent the mean line of the airfoil at that station. The incremental angles, however, must be converted to their streamwise orienta­tion to obtain the mean line of the streamwise section.

A typical cambered configuration (L6T15) is shown in figure 4(c). The schematic diagram shows the shape of the mean camber line at inboard and out­board stations when camber and twist are applied to the wing panel by deflect­ing the leading-edge and trailing-edge segments. The trailing-edge segments

i n c r e a s e the inc idence o f the inboard p a r t o f t h e wing pane l , and the leading-edge segments reduce the i n c i d e n c e of the outboard p a n e l s ; t h u s , t h e o r i g i n a l uncambered wing can have an e f f e c t i v e t w i s t o f approximate ly 80 f o r t h e conf ig­u r a t i o n shown. The new chord l i n e s and t h e p o s i t i o n of maximum camber are a l s o shown for the two semispan s t a t i o n s .

F i v e t y p e s o f camber were i n v e s t i g a t e d w i t h t h e p r e s e n t v a r i a b l e camber model. The streamwise v a r i a t i o n of maximum camber and s e c t i o n inc idence is as a f u n c t i o n of t h e semispan are shown i n f i g u r e s 5(a) t o 5 ( d ) . Three o f these t y p e s involved the leading-edge v a r i a b l e segments only . They i n c l u d e an e l l i p t i ca l -type camber ( t h e mean l i n e o f the forward p o r t i o n o f t he a i r f o i l w a s shaped t o approximate an e l l i p s e , t he c o n f i g u r a t i o n s so cambered be ing des igna ted L5To and L ~ T o ) ,a c i r c u l a r - t y p e camber ( t h e mean l i n e of t he a i r f o i l w a s shaped t o approx­imate a c i rc le , t he c o n f i g u r a t i o n s so cambered be ing des igna ted L24To and LgTO), and a t h i r d type i n which t h e number o f movable segments w a s reduced t o s i m p l i f y the camber a c t u a t i o n system (these were des igna ted L25To and L28TO f o r two-segment c o n f i g u r a t i o n s and L29To f o r the one-segment c o n f i g u r a t i o n ) . The L6TO c o n f i g u r a t i o n w i t h an e l l i p t i c a l l e a d i n g edge was used w i t h t he t r a i l i n g - e d g e camber v a r i a t i o n s . The t r a i l i n g - e d g e camber was of two t y p e s , a sys t ema t i c v a r i ­a t i o n o f i n c r e a s i n g e l l i p t i c a l camber (L6T11, L6T1, and L6T10) and a n a t t empt wi th a s i n g l e c o n f i g u r a t i o n t o i n c r e a s e t h e l i f t c o e f f i c i e n t t o 1.0 a t M = 0.90 by a s e l e c t i v e i n c r e a s e i n t he t r a i l i n g - e d g e camber (L6T15).

The mechanism f o r changing the camber was l o c a t e d w i t h i n the wing, and the ad jus tmen t s were made through t h e h inge l i n e s ' o n the lower surface. The p r o t r a c ­t o r s f o r measuring and s e t t i n g the segment a n g l e s had magnetic bases designed t o f i t the upper s u r f a c e o f the wing; t h e p r o t r a c t o r s f o r leading-edge segments 1 and 3 and t r a i l i n g - e d g e segments 6 and 8 are shown i n p o s i t i o n i n f i g u r e 2 ( b ) . The h inge l i n e s were f i l l e d and smoothed t o contour w i t h a s i l i c o n e material t o p reven t a i r flow from the lower s u r f a c e t o t h e upper s u r f a c e . Aluminum t a p e w a s a l s o a p p l i e d t o the lower-surface h inge l i n e s af ter each camber adjus tment as a f u r t h e r p recau t ion a g a i n s t f low through the h inge l i n e s .

APPARATUS AND PROCEDURES

T e s t F a c i l i t y

The i n v e s t i g a t i o n w a s conducted i n t he Langley 8- foot t r a n s o n i c p r e s s u r e t u n n e l . This f a c i l i t y is a continuous-flow s i n g l e - r e t u r n r e c t a n g u l a r s l o t t e d -t h r o a t t u n n e l having c o n t r o l s t h a t a l low f o r independent v a r i a t i o n of Mach num­b e r , d e n s i t y , s t a g n a t i o n t empera tu re , and dewpoint tempera ture . The test sec­t i o n is approximately 2.2 m (7 .1 f t ) square . The c r o s s - s e c t i o n a l area is the same as t h a t of a c i rc le w i t h a 2.4-m (8-f t ) diameter. The upper and lower walls are a x i a l l y s l o t t e d t o permi t the te . s t - sec t ion Mach number t o be changed con t inuous ly throughout t h e t r a n s o n i c speed range . The s l o t t e d t o p and bot­tom w a l l s each have an ave rage open r a t i o o f approximate ly 0.06. The s tagna­t i o n p r e s s u r e i n t h e t u n n e l can be v a r i e d from a minimum o f about 0.25 atm ( 1 a t m = 0.101 ma) a t a l l Mach numbers t o a maximum o f approximate ly 2.00 atm a t Mach numbers l e s s than 0.40. A t t r a n s o n i c Mach numbers, however, the maxi­mum s t a g n a t i o n p r e s s u r e t h a t can be obta ined is about 1.5 atm.

5

Boundary-Layer Transition

Boundary-layer transition was'fixed at 5 percent of the local chord on the wing by the addition of a 0.25-cm (0.64-in.) wide (streamwise) strip of No. 100 carborundum grains and on the fuselage by a strip of No. 80 carborundum grains located 6.4 cm (16.3 in.) aft of the nose apex.

Instrumentation

Five-component static aerodynamic force and moment data were obtained for the wing using a wall-mounted strain-gage balance to which the model wing was attached. The fuselage, which was nonmetric, was mounted on a turntable that moved with the balance and wing and therefore remained at the same angle of attack as the wing. The angle of attack of the wing was measured with an accel­erometer located in the wing attachment block.

Strain gages were mounted inboard in the wing upper and lower surface at rl = 0.15, and an accelerometer was installed at rl = 0.93 as shown in fig­ure l(b). The root-mean-square output from these instruments was integrated for 45 sec, and coefficients were computed to determine the buffet characteristics of the wing.

Test Conditions and Data Reduction

Tests were made at angles of attack from Oo to loo or buffet onset at Mach numbers of 0'.60,0.80,and 0.90 for Reynolds numbers based on the mean geometric chord of 6.2 x lo6, 7.4 x lo6, and 7.6 x lo6, respectively. The basic uncambered configuration was also investigated at RE of 5.9 x 1Q6 and 9.8 x lo6 at M = 0.80. The maximum camber configurat5on L6T15 of the investigation was also run at RE = 5.9 x 106.

The constants for reducing the data were based on the exposed area and lin­ear dimensions of the semispan wing panel.

PRESENTATION OF RESULTS

The results of the investigation of the longitudinal aerodynamic and buffet characteristics of the variable camber and twist wing are presented in the fol­lowing figures:

FigureLongitudinal aerodynamic characteristics: Effect of Reynolds number . . . . . . . . . . . . . . . . . . . . . . . 6 Effect of elliptical leading-edge camber . . . . . . . . . . . . . . . 7 Effect of circular leading-edge camber . . . . . . . . . . . . . . . . 8 Effect of simplified leading-edge camber . . . . . . . . . . . . . . . 9 Effect of elliptical trailing-edge camber . . . . . . . . . . . . . . . 10 Effect of a selective increase in trailing-edge camber . . . . . . . . 1 1 Optimum-camber aerodynamic characteristics . . . . . . . . . . . . . . 12

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Figure Buffet characteristics: Effect of Reynolds number . . . . . . . . . . . . . . . . . . . . . . . 13 Effect of elliptical leading-edge camber . . . . . . . . . . . . . . . 14 Effect of circular leading-edge camber . . . . . . . . . . . . . . . . 15 Effect of simplified leading-edge camber . . . . . . . . . . . . . . . 16 Effect of elliptical trailing-edge camber . . . . . . . . . . . . . . . 17 Effect of a selective increase in trailing-edge camber . . . . . . . . 18

DISCUSSION

The following discussion pertains to force and moment-dataobtained on the wing only since the fuselage of the model was nonmetric. The data from the results of this investigation are presented and analyzed in regard to increments in the forces and moments obtained from camber and twist of the wing alone and are not directly comparable with three-dimensional wing and body configurations.

Basic Longitudinal Data

As shown in figure 6, increasing the Reynolds number caused a slightdecrease in drag resulting in a 6-percent increase in (L/D)max.

Leading-Edge Camber

The camber and twist distributions across the semispan for the various leading-edge configurations are presented in figures 5(a) and 5(b), respectively.The effect of elliptical leading-edge camber (L5To and L6TO) is shown in fig­ure 7 for the model without trailing-edge camber. This type of camber reduced the drag over much of the lift-coefficient range of the investigation and, as a result, the maximum lift-drag ratio (L/D)max and the lift coefficient at (L/D)aax are increased at all Mach numbers for these configurations. The L6To leading-edge configuration had incremental deflection angles of 12.8O, 4.7O, 2.8O, and 1.2O for leading-edge segments 1 to 4, respectively, and is the basic leading edge used in this investigation. It is evident (fig. 7(c)) that this camber is excessive for M = 0.90 at lift coefficients up to 0.3 as the L5To configuration with less camber has higher lift-drag ratios for these conditions.

The effect of circular leading-edge camber (L24To and L~To)is shown com­pared to that of the elliptical leading-edge camber (LgTo) in figure 8. The circular-camber configuration (L24To) shows the best improvement in drag of the three leading-edge configurations; it also shows, at Mach 0.80,an increase in (L/D)max and CL at (L/D)max although at Mach 0.90, it shows a loss in (LID),, -

In an effort to determine if the camber-changing mechanism could be sim­plified by reducing the number of movable segments, one-segment and two-segment Configurations of the leading edge were investigated. The one-segment configu­ration L29To was obtained by a 7.70 deflection of segment 3 while the other’seg­ments were left at zero angle. As shown in figures 5(a) and 5(b), this configu­

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ration had the least amount of camber and twist at the wing-tip station, but the distribution of camber and twist over the semispan was similar to the L6To elliptical leading edge. The two-segment configurations,L25To and L28T0, were obtained by deflecting the second and third segments and leaving segments 1 and 4 at zero angle. The two-segment configuration L25To had a low camber and twist similar to the one-segment L29To and L6To configurations. The high camber and twist two-segment configuration L28To had a camber distribution similar to the circular configuration L24To with somewhat more camber in the outboard part of the semispan. The results are presented in figure 9, and the single-segment configuration appears to compare favorably with the elliptical-camber configura­tions at all Mach numbers of the investigation. The two-segment configuration L28TO had significantly better aerodynamic characteristics at M = 0.60 than the other leading-edge camber configurations at lift coefficients greater than 0.4.

In general, the leading-edge camber reduced the values of lift coefficient by small amounts over the angle-of-attack range, the reduction being greater at M = 0.90 where the increment was approximately 0.05. (Compare the LOTO curve of fig. 7(c) with the L8To circular-camber curve of fig. 8(c) and the L28To two-segment curve of fig. 9(c).) It should be noted that the leading-edge cam­ber configurations cause a washout in the wing panel (see fig. 4(c)) because of the hinge line sweep, and the reductions in lift are at the outboard region of the wing panel. The camber and washout improve the span loading on the wing panel and cause a reduction in drag of approximately 28 to 40 percent compared to that of the uncambered wing at Mach numbers of 0.60 and 0.80 and 20 percent at a Mach number of 0.90, in the moderate lift-coefficient range from 0.35 to 0.45. The favorable influence of the leading-edge camber on the drag provides increases in (L/D)max of approximately 19 percent with the L25To configura­tion at Mach numbers of 0.60 and 0.80,and the increase was the same with the L5To configuration at M = 0.90.

The leading-edge camber (see figs. 7, 8, and 9) generally caused a reduc­tion in the pitching-moment coefficients. The elliptical-type camber also had a small destabilizing trend at Mach numbers of 0.60 and 0.80,whereas the circu­lar type had a stabilizing trend at these same Mach numbers. At M = 0.90 the circular camber LgTo and the two-segment L28TO increase the stability over the lift-coefficient range from 0 to 0.35 and decreased the stability at CL greaterthan 0.35.

Trailing-Edge Camber

The camber and twist distribution across the semispan for the trailing-edge camber configurations is shown in figures 5(c) and 5(d), respectively. The L6TO leading-edge camber is used with all of the trailing-edge configurations. The LgTll, L6T1, and L6T10 configurations show systematic increases in an elliptical-type trailing-edge camber. For the L6T15 configuration, the camber in the inboard 60 percent of the wing was increased substantially in an attempt to increase the wing lift coefficient to 1.0. The effect of the elliptical trailing-edge camber is shown in figure IO. As would be expected, increases in the trailing-edge camber caused large increases in the lift coefficient CL over the angle-of-attack range at all Mach numbers of the investigation.

The drag is reduced an additional amount by the trailing-edge camber. Comparing the L6To configuration to the L6Tlo configuration in figure IO, this reduction amounts to 34 percent at CL = 0.65 (highest CL obtained for L6TO at M = 0.80) at Mach numbers of 0.60 and 0.80 and 27 percent at a Mach number of 0.90. If higher lift coefficients are selected, another 5-percent reduction in drag is evident. For example, at M = 0.60 and CL = 0.7 the reduction is 39 percent, and at M = 0.90 and CL = 0.75 (the values for L6To are extrap­olated from L29To in fig. 9(c)) the reduction is 32 percent.

More camber was added to the trailing-edge segments LfjT15 in an effort to increase CL at buffet onset. The results are shown compared to the zero cam­ber configuration and the elliptical trailing-edge camber configuration L6T10 in figure 11. The lift coefficient was substantially increased at all Mach num­'bers of the investigation by the increased trailing-edge camber. While the lift coefficient increased to a value greater than 1.0 at M = 0.90 and near 1.0 at the other Mach numbers, it is not intended to suggest that this camber and twist combination is most efficient for the high Mach number and lift-coefficient range.

Drag polars, pitching-moment coefficient, and angle of attack for the basic uncambered wing compared to a best-camber polar derived from the various cam­bered configurations of the investigation are presented in figure 12 for all Mach numbers of the investigation. These polars are not necessarily the best camber and twist combinations for this aspect ratio and wing planform but rather the best combination obtained in this-investigation. Values of Cm and CL against c1 are shown for the data points on the drag polar only, with no attempt to arrive at a continuous curve for CL, Cm, and CY.

The leading-edge camber lowers the drag substantially in the low lift-coefficient range, while the trailing-edge camber is necessary for the very large improvements in the high lift-coefficient range. The trailing-edge camber causes very large increments in CL at constant angle of attack with substan­tial negative shifts in pitching-moment coefficients.

Buffet Characteristics

The buffet indicators of axial-force coefficient CA, wing-tip accelerometer Cacc in g Units, and wing root-mean-square bending-moment coefficient Cb,rms are presented in figures 13 to 18 as a function of the lift coefficient. Most of the buffet analysis is based on data from the wing-root bending gages, and the axial-force coefficient and wing-tip accelerometer are used for comparison pur­poses only. The small Reynolds number variation at M = 0.80 (fig. 13) had lit­tle effect on the buffet-onset characteristics of the model; however, there are some increases in intensity level as a result of the increased dynamic pressure.

The elliptical and circular leading-edge camber configurations (figs. 14 and 15, respectively) tend to extend the buffet onset to somewhat higher lift coefficients and to reduce the intensity levels.

9

As would be expected, the trailing-edge camber configurations provide the largest increase in buffet-free lift coefficient for the Mach numbers of this investigation. The buffet data for 'the elliptical trailing-edge camber varia­tions are presented in figure 17. Buffet onset is indicated by the marks on the curves of Cb,rms against CL and is established at the tangent point to the curve of a line drawn at a 45O angle to the axes. The increment in lift coeffi­cient compared to the L6To configuration at buffet onset for L6To is 60 percent and 40 percent at Mach numbers of 0.60 and 0.80, respectively. At M = 0.90 (fig. 17(c)) the lift-coefficient increase at buffet onset is 46 percent. '

The configuration with the selective increase in trailing-edge camber LgT15 is compared to the uncambered wing configuration LOTO and the maximum elliptical leading-edge and trailing-edge camber configuration L6T10 in fig­ure 18. The increase in buffet-free lift coefficient for this configuration L6T15 is 102 percent and 68 percent at Mach numbers of 0.60 and 0.80,respec­tively. Buffet onset at M 0.90 for this configuration is not indicated in these data.

CONCLUDING REMARKS

The results of an investigation t o determine the effects of camber and twist applied to a semispan wing in the presence of body by deflecting multiple leading-edge and trailing-edge segments indicated that all of the leading-edge camber configurations investigated were effective in reducing the drag at lift coefficients up to 0.4, the elliptical and circular camber were the most effective, and the simplified single hinge configuration was similar in performance to the more complex four-segment elliptical camber and twist that it simulated. The value of the maximum lift-drag ratio was increased approximately 18 percent over the Mach number range from 0.60 to 0.90 by the leading-edge camber. At the higher lift coefficients (0.50 and above) the combination of camber and twist applied by both leading-edge and trailing-edge camber was effective in reducing the drag and the trailing-edge camber gave large increases in the lift coefficient. The combination of leading-edge and trailing-edge camber and twist increased the lift coefficient at buffet onset from 68 percent to more than 100 percent over the Mach number range of the investigation.

Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 May 19, 1977

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REFERENCES

1. Dollyhigh, Samuel M.: Subsonic and Supersonic Longitudinal Stability and Control Characteristics of an Aft Tail Fighter Configuration With Cambered and Uncambered Wings and Uncambered Fuselage. NASA TM X-3078, 1974.

2. Ray, Edward J.; and Taylor, Robert T.: Buffet and Static Aerodynamic Charac­teristics of a Systematic Series of Wings Determined From a Subsonic Wind-Tunnel Study. NASA TN D-5805, 1970.

3. Ray, Edward J.; McKinney, Linwood W.; and Carmichael, Julian G.: Maneuver and Buffet Characteristics of Fighter Aircraft. NASA TN D-7131, 1973.

4. Monaghan, Richard C.; and Friend, Edward L.: Effects of Flaps on Buffet Char­acteristics and Wing-Rock Onset of an F-8C Airplane at Subsonic and Tran­sonic Speeds. NASA TM X-2873, 1973.

5. Friend, Edward L.; and Sefic, Walter J.: Flight Measurements of Buffet Char­acteristics of the F-104 Airplane for Selected Wing-Flap Deflections. NASA TN D-6943, 1972.

6. Fischel, Jack; and Friend, Edward L.: Preliminary Assessment of Effects of Wing Flaps on High Subsonic Flight Buffet Characteristics of Three Air­planes. NASA Tf.I X-2011, 1970.

7. Sisk, Thomas R.; Friend, Edward L.; Carr, Peter C.; and Sakamoto, Glenn M.: Use of Maneuver Flaps To Enhance the Transonic Maneuverability of Fighter Aircraft. NASA TM X-2844, 1973.

8. Sisk, Thomas R.: A Preliminary Assessment of the Transonic Maneuvering Char­acteristics of Two Advanced Technology Fighter Aircraft. NASA TM X-3439, 1976.

9. Pendergraft, Odis C., Jr.: Effect of Camber and Twist on Longitudinal Aerody­namic Characteristics of a Wing-Fuselage Model at Mach Numbers up to 1.30. NASA TM X-1610, 1968.

TABLE I.- WING AIRFOIL COORDINATES ALONG STREAMWISE CHORDS

Root NACA 658005 a i r f o i l (modif ied) Tip NACA 65A004 a i r f o i l (modif ied)Leading-edge r a d i u s = 0.00150 Leading-edge r a d i u s = 0.00102

Local chord Local chord

x / c z /c x/c Z / C

0.0000 0.00000 0.0000 0.00000 ,0050 .00385 .0050 .00304 .0075 .00467 .0075 .00368 ,0125 .00595 .0125 .00469 .0250 .00815 .0250 .00647

,0500 .01094 ,0500 .00875 .0750 .0 1326 - 0750 .0 1059 .IO00 .01519 . loo0 .01213 .1500 .0 1826 .1500 .o 1459 .2000 ,02059 .2000 .01645

.2500 .02237 .2500 .0 1788

.3000 .02367 .3000 .O 1892

.3500 .02453 .3500 .01962

.4000 .02496 .4000 .o 1997 ,4500 .02494 .4500 - 01996

.5000 .02440 .5000 .O1954

.5500 .0233 1 .5500 .O1868

.6000 .02 173 .6000 .01743

.6500 .0 1976 .6500 .0 1586

.7000 .01746 .7000 .0 1402

.7500 .0 1490 .7500 .01197

.8000 .01229 .8000 .00987

.8500 .00969 .8500 . .00778 .goo0 .00708 .goo0 .00569

.9500 .00448 .9500 .00359 1 .oooo .OO 188 1 .oooo .OO 150

12

Wind-tunnel woll

1- 6.03 I -- --

Fuselage plane of symmetryFus

1W ind-tunnel wollL--T/I I

~ I

1 0k- 100.663 0 . 6 6 3106.68 1 0 6 . 6 6I038 ~ I

142574 3 -Moment cenler-18.469- ,

- _ _ _ - -I---_ _ _ -I

(a) General arrangement of model.

Figure 1 .- Drawings of the wind-tunnel model. (Dimensions are in cm.)

---

3.54 -

I 80.21

I

i

!

24.03

81.85 L- - 55.71 - a

~t-29.24'-

I-

* 15.47 -/ \\\\

_ - _.- Yti.17 ­

- . ~ - 106.68

99.21

(b) Details of t h e wing segments and planform.

F igure 1 . - Continued.

14

i i ! I ! II I I II I

I

( c ) Fuselage c ros s s e c t i o n s .

Figure 1 . - Concluded.

I

Figure 2.- Photographs of model i n s t a l l a t i o n .

16

(b) Rear.

Figure 2.- Concluded.

17

0

u) ln

200- 1300­

1200­180­

1 1 0 0 ­-I60

DO0 ­

140- 900­

(u (u

.‘I 20-:E800­0 0

02 6700­0 ­=100-0 .-c .-3500­c c 0 0

0)? 80-r500­ln ln

2 2 0

60 -O400­

-40

-20

/- \ 0- I I I I I I I \ I I I I

0 25 50 75 100 125 150 175 200 225 250 275 300

Figure 3.- Model c r o s s - s e c t i o n a l area d i s t r i b u t i o n .

18

Wing stations selected for camber and twist

I

(a) Geometric arrangement of the va r i ab le camber segments of the wing panel.

Figure 4.- Details of the va r i ab le camber segments.

A

A i r f o i l m e a n l i n e

S e c t i o n A - A

Air foi l m e a n l i n e

S e c t i o n B - B

(b) Schematic drawing of t h e v a r i a b l e camber segment o r i e n t a t i o n .

F igure 4.- Continued.

20

i = -3.8'S /-Mean camber line

i + 4. 1'S

/ / GMean camber line

\ T L o r d line !Maximum camber .

( c > Camber and t w i s t i l l u s t r a t i o n .

F igu re 4.- Concluded.

21

a

--

-------

-----

--

-- --------

Iu Iu .07.

.O 6

L

a, 0 5 E U 0 ---aL5 To Ellipticol 9.1 3.2 1.8 0.8­

5 ,04 -.a L6To Elliptical 12.8 4.7 2.8 1.2 .-E 9L2J0 Circular 4.4 4.5 5.2 2.1 x U

E .03 a, v,

3 E g .02 L t v,

-01 -*-

0 0 .I .2 .3 4 .5 .6 .7 .8 ,9 1.0

Semispon, b/2 ( a ) Spanwise camber v a r i a t i o n of the leading-edge conf igura t ions .

Figure 5.- Spanwise camber and t w i s t d i s t r i b u t i o n .

.-

a-aU -0 c -2 l o

-80

N W

'Code ' Type AB, A82 L5 To Elliptical 9.1 3,2 L, To Elliptical 12.8 4.7 L ; ~ T ~Circular 4.4 4.5 L8 To Circular 7.2 7.8 k 9 T o I-segment 0 0 h5T0 2-segment 0 8.4 k 8 T 0 2-segment 0 12.1

.I .2 . .3

I.8 2.8 5.2 ;T8.2 7.7 0 4.5 0 12.1 0

4 .5 .6 .8 Semispan, b/2

(b) Spanwise inc idenc,e v a r i a t i o n of the leading-edge conf igura t ions .

F igure 5.- Continued.

.9

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

--- ------ ----------- 0 0 .I .2 .3 .4 .5 .6 .7 .8 9 1.0

Semispan, b/2 ( c > Spanwise camber v a r i a t i o n of t he t r a i l i ng -edge conf igu ra t ions .

F igure 5.- Continued.

( d ) Spanwise incidence v a r i a t i o n of t h e t r a i l i ng -edge conf igura t ions .

F igure 5.- Concluded.

.6 .7 .8 1.1

Figure 6.- Effect of Reynolds number on t h e l o n g i t u d i n a l aerodynamic characteristics. M = 0.80.

26

24

20

16

12

L/D

8

. 4

0

- 4

.I c

.OE

.Ot

CD

.O4

.O2

C- 0 . I .2 .3 .4

Figure

.6 .7

6.- Concluded.

.a .9 1.0 1.1

27

.o 4

0

-.o 4

a

I 1

10

9

8

7

6

5

4

3

2

I

0

_ I -.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 ID 1.1

CL

( a ) M = 0.60; R,- = 6.2 x lo6.

Figure 7.- Effect of e l l i p t i c a l leading-edge camber on the longitudinalaerodynamic characterist ics.

28

24

20

16

12

8

4

0

-4

-0 . I .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1

CL

(a> Concluded.

F i g u r e 7.- Continued.

I

Cm

a ,de

(b) M = 0.80; RE = 7.4 x lo6.

F igu re 7.- Continued.

28

24

20

16

12

8

4

0

- 4

.I 0

.o8

.06

.04

.o2

% (b) Concluded.

I

F i g u r e 7.- Continued.

31

1

a

CL

( c > M = 0.90; Re = 7.6 x lo6.

Figure 7. - Continued.

32

L

-.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 ID 1.1 C L

( c Concluded . Figure 7.- Concluded.

33

CL

( a ) M = 0.60; R,- = 6.2 x IO6.

Figure 8.- Effect of c i r c u l a r leading-edge camber 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 .

34

28

24

20

16

12 L/D

8

4

0

- 4

. I 4

. I 2

,I c

.OE

CD .OE

.O4

.O2

C - 0 . I CL

(a ) Concluded.

F igu re 8.- Continued.

35

cm

Wing-segment deflection,deg Code A81 A82 A83 A84 .A85 A86 A87 A88 Lh To 0 12.8 4.7 2.8 1.2 0 0 0 0 $dTo 4.4 4.5 5.2 2.5 0 0 0 0

0 7.2 7.8 8.2 4.2 0 0 0 0

a,de

- . I 0 .I .2 .3 .4 .5 .6 .7 .8, .9 ID 1.1

CL

(b) M 0.80; RE = 7.4 x lo6.

Figure 8.- Continued.

36

I

28

24

i

20

16

12

8

4

0

-4

I

(b) Concluded.

F igu re 8.- Continued.

37

(c) M = 0.90; RE = 7 . 6 x IO6.

Figure 8.- Continued.

38

28

24

20

i 16 i

12 L

8

4

0

-4

.I c

.oa

.0e

CQ . 0 4

.Oi

C-

( c > Concluded.

Figure 8.- Concluded.

39

( a ) M = 0.60; RE = 6.2 x lo6.

F igu re 9.- Effect of s i m p l i f i e d leading-edge camber on the l o n g i t u d i n a l aerodynamic characteristics.

40

'1

Wing-segment deflection,de9 t A82 A83 A84 A85 A86 A87 A&[,, 4.7 2.8 1.2 0 0 0 0

L

.I 4

.I 2

,I c

.O E

CD .O E

.OL

.Oi

C -

0 7.7 0 8.4 4.5.~ - . I

12.1 12.1 0

-2 . 3 .4 .5 .6 .7

CL

. ( a ) Concluded.

F igu re 9 .- Continued.

0 0 0 0 1 0 0 0 o o o 124

20

16

12

8

4

0

-4

.9 1.0 1.1

41

I I., I . . ..-- .-

a

0 .I -2 -3 .4 .5 .6 .7 .8 .9 ID 1.1

cL

(b) M = 0.80; RE = 7.4 x lo6 .

Figure 9 .- Continued.

42

28

24

20

16

12 'D

8

4

0

-4

.OE

.OE

'D .04

.Oi

C-

(b) Concluded.

F igure 9.- Continued.

43

L

cL

( c ) M = 0.90; Re = 7.6 x lo6 .

Figure 9.- Continued.

44

24

20

16

12 L/D

8

4

0

-4

.I 2

.I 0

.oa

CD .06

-04

.o 2

0- I

( c > Concluded.

F igure 9 . - Concluded.

45

cm

a ,dt

( a ) M = 0.60; RE = 6.2 x I O 6 .

F igu re 10.- Effect o f e l l i p t i c a l t r a i l i n g - e d g e camber 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 .

46

I

Wing-segment deflection ,deg

12.8 4.7 2.8 1.2 0 12.7 4.6 3.1 1.2 .7 1.5 2.2 2.1 12.9 4.7 3.0 1.3 1 . 1 2.4 28 6.3

.I 6

20

16

1 2L/D

8

4

0

- 4

.I 4

.I 2

, I 0

cD .08

.06

.O4

.02

0 lllllll.6 - 0 .I .2 .3 .4 .5 .7 .9 1.0 1.1

CL

(a) Concluded.

F igu re 10 .- Continued.

47

(b) M = 0.80; RE = 7.4 x IO6.

Figure 10.- Continued.

40

28

24

20

16

12 L/D

8

4

0

- 4

.I 2

.I c

.O8

CD .06

.Od

.O2

C- 0 .I *2 -3 .4 .5 $6 -7 .8 .9 ID 1.1 CL

(b) Concluded.

F igu re 10.- Continued.

49

CL

( c ) M = 0.90; RE = 7.6 x lo6.

Figure 10 .- Continued.

28

24

20

16

12 L

8

4

0

-4

C'D

-.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 ID I;i CL

(c> Concluded.

Figure 10.- Concluded. ,

51

I 111111111 11111111111m11111111I

CmI -,

-.

-.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 ID 1.1

cL

( a ) M = 0.60; R,- = 6.2 x lo6.

Figure 11.- Effect of a selective increase i n trailing-edge camber on the . longitudinal aerodynamic characterist ics.

52

... ..

24

20

16

12 L

8

4

0

. le -4

.I4

.I 2

.I c

CD .oe

.Of

.O4

.o 2

. c - 0 .I -2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1

CL

(a> Concluded.

Figure 11.- Continued.

53

CL

(b) M = 0.80; RE = 7.4 x lo6 .

Figure 11 .- Continued.

54

24

20

16

12

8

4

0

- 4

CL

(b) Concluded.

Figure 11 .- Continued.

55

.O4

0

-.04

-.OB

-.I 2

-.I 6

-.20

9

8

7

6

5

'9 4

3

2

I

0

Figure 11 .- Continued.

56

24

20

16

12

8 L

4

0

- 4

- 8

-.I .I .2 .3 .4 .5 .6 .7 .8 .9 ID 1.1

( c > Concluded.

Figure 11 .- Concluded.

57

, I, I . m . I , ......__ .

.I5 Wlng-sqmen! deflection. d q

Cod0

. I4 LOT0 O 0 0 ? O

L5 To a 8.8 3.4 2.1 . 7

$4T0 4.4 4.5 5.2 2.5 12.7 4.6 3.1 1.2

. I3 ‘6 L6 TI ’ 12.9 4.7 3.0 1.3

‘6 TtO a 12.8 4.6 3.1 1.2

‘6 12.8 4.6 3.0 1.2

. I I

. I O

.09

.O8

CD .07

.06

.05

.04

.03

.02

.o I

0-. I 0 .2 -4 .5 .6

( a > M = 0.60.

0 0

0 ’ 0 o h 0 0 O f 1.7 2.2 2.1

4 2.8 6.3

0 7.0 9.0 1

3 4.3 9.3 1

. I7 .8 .9 I.o

Figure 12 .- Optimum-camber aerodynamics.

58

II I I I 1 1 I ill Wlng-logmen! dcfledlcn. dog

C d C Ab, A% A 4 Ab4 Ab5 Abb Ab, Ab8

I O % T O O 0 0 0 0 0 0 0 0

T %To 0 4.4 4.5 5.2 2.5 0 0 0 0

L b T l l 12.1 4.6 3.1 1.2 .I 1.1 2.2 2.1 9 1.1,IlllHE: 0

E 12.9 12.8 12.8

4.1 4.b

4.6

3.0 3.1 3.0

1.3 1.2 1.2

1.1 1.5 4.0

2.4 3.3 7.0

2.8 4.3 7.0

6.3 9.3 9.0

L ~ a ~ 8.8 3.4 2.1 .I o o o o

8

7

6

5

4

3

2

I

0

- I -.I 0 .I .2 .3 .4 .5 .6 .7 .8 ,9 1.0

CL

( a ) Concluded.

F igu re 12.- Continued.

59

A 4 To O 0

i To O 8.8

!4 To O 4.4

1 '0 1.2

i '11 'S 12.7

i 'I 12.9

i TI0 O 12.8

, TI5 O 12.8

t

-.I 0 .2

. . . . . . . .. . Wlng-segment dcfledlon. d q

A%- A 4 A 4 A 4 0 0 0 0 0 0 0

3.4 2.I .7 0 0 0 0

4.5 5.2 2.5 0 0 0 0

7.9 8.2 1.I 0 0 0 0

4.6 3.I 1.2 . I I. 5 2.2 2.I

4.1 3.0 1.3 I.I 2.4 2.8 6.3 4.b 3.I 1.2 I . 5 3.3 4.3 9.3 4.6 3.I 1.2 4.0 7.0 7.0 9.0

.3 .4 .6 .7 .a .9 1.0'

M = 0.80.

Figure 12.- Continued.

60

3

(b) Concluded.

F igu re 12.- Continued.

61

. IJ

A 9 A$ A 4 A64 AB5 0 0 0 0 0

. I 4 9. I 3.2 1.8 .8 0 iza 4.7 2.8 1.2 0 4.4 4.5 5.2 2.5 0

. I3 12.7 4.6 3.1 1.2 .7 12.9 4.7 3.0 1.3 1.1 p a 4.6 3. I 1.2 1.5 12.8 4.6 3. I 1.2 4.0

.I2

.I I

.I O

.09

.08

CD .07

.06

.05

.04

.03

.02

.o I

0 -.I 0 . I .2 .3 .4

( C > M =

Figure 12.-

A 5 A% 0 0 0 0 0 0 0 0 0 0 0 0 1.5 2.2 2. I 2.4 2. I 6.3 3.3 4.3 9.3 7.0 7.0 9.0

-5 .6 I .o I . I CL

0 .90 .

Continued.

62

c,

.04

0

-.04

-.08

-. I 2

-. I 6

-.2c

9

8

7

6

5

4

3

2

I

0

-I-

63

I I 1111

-02

cA

-.02

.016

,012

Cb,rnns . -008

.004

0- I 0 .I .2 .3 .5 .6 .. I .o 1.1

CL

Figure 13 .- Effect o f Reynolds number on the b u f f e t c h a r a c t e r i s t i c s . M = 0.80.

64

wing-segment &flection,de A61 A62 A63 A64 ASS A66

0 0 0 0 0 0 0 8.8 3.4 2.1 .7 0 0

0 12.8 4.7 2.8 1.2 0 0 D2

0

CA

-.02

- .04

6

4

2 cacc

0

-2

-020

.016

.o 12 Cb,rms

.008

.a34

0-.I .I .2 .3 .4 .5 .6 .8 .9 IO I.!

CL

( a > M = 0.60; R,- = 6.2 x IO6 .

Figure 14.- Effect of e l l i p t i c a l leading-edge camber on the buffet character is t ics .

65

C

6

4

2 ell,

0

- 2

,rms

LL

(b) M = 0.80; R,- = 7.4 x IO6.

Figure 14.- Continued.

66

8

6

4

Cam

2

0

Cb,r

-.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 ,.

( C > M = 0.90; RE = 7.6 x lo6. Figure 14.- Concluded.

U

67

( a ) M = 0.60; RE = 6.2 x IO6.

Figure 15.- Effect of c i rcular leading-edge camber on the buff e t characterist ics .

68

-0:

(

CA -.0:

-.O'

.o li

.ooz 'b,rms

. 0 0 4

0-. I 0 . I .2 .3 .4 -5 .6 .7 .8 .9 1.0 1.1

C L

( b ) M = 0.80; RE = 7.4 x IO6.

Figure 15.- Continued.

69

,,...-.-,.. ..

Wing- segment deflection,deg A81 A62 A63 A84 A65 A86 A87 A88

Lb To 0 12.8 4.7 2.8 . 1.2 0 0 0 0 $4To 0 4.4 4.5 5.2 2.5 0 0 ’ 0 0

9 T~ 0 7.6 7.9 8.2 4.1 0 0 0 0

-2

-.I .I .2 -3 :.4 .5 .6 .7 .8 .9 ID 1.1

c L

( c ) M = 0 . 9 0 ; R,- = 7.6 x IO6

Figure 15 .- Concluded.

70

0 2

0

CA -.02

-.04

4

2

0

.2

.020

.O16

rms.0 I z

.OW

.004

0 -.I 0 .I .2 -3 .4 .5 -6 .7 .8 .9 ID 1.1

cL

(a) M = 0.60; R,- = 6.2 x lo6.

Figure 16.- Effect of simplified leading-edge camber on the buffet characteristics.

71

I

4

2

cocc

0

.2

" -.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 .o I

cL

(b) M = 0.80; R-, = 7.4 x IO6.

Figure 16.- Continued.

72

6

4

2

0

-2

1 . 1 0 .I .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1

cL

( c ) M = 0.90; RE = 7 .6 x IO6.

F igu re 16.- Concluded.

73

6

4

2 cacc

0

.2

CL

( a ) M = 0.60; RE = 6.2 x 106 . Figure 17.- Effect of e l l i p t i a a l trailing-edge camber on the

buffet character is t ics .

74

6

4

2

0

.2

-.I 0 .I .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1

cL

(b) M = 0 .80 ; Rg = 7.4 x IO6.

Figure 17.- Continued.

75

.O4

-02

CA 0

-.02

6

4

.o I 2

CL

( c ) M = 0.90; Re = 7 .6 x 106.

Figure 17. - Concluded.

76

.O 4

0 2

cA 0

-.02

- -04

.024

.02c

.O 16

‘b,rrrs .o I 2

.OOE

.004

Figure 18.- Effect of a selective increase i n trailing-edge camber on the buffet character is t ics .

77

6

4

2

0

- 2

CL

(b) M = 0.80; R,- = 7.4 x IO6.

Figure 18.- Continued.

78

.06

.O4

0

-.02

8

6

4

Gocc

2

0 '

.024 -2

.020

.016

C 'ms.012

.008

.004

-0 I

CL

( e > M = 0.90; RE = 7.6 x 106.

F i g u r e 18.- Concluded.

NASA-Lanqley, 1977 L-I 1357 79

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NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS TECHNICAL REPORTS: Scientific and ,

technical information considered important, complete, and a lasting contribution to existing knowledge. TECHNICAL NOTES: Information less broad in scope but nevertheless of importance as a contribution to existing knowledge.

TECHNICAL MEMORANDUMS: Information receiving limited distribution because of preliminary data, security classifica­tion, or other reasons. Also includes conference proceedings with either limited or unlimited distribution. -CONTRACTOR REPORTS: Scientific and technical information generated under a NASA contract or grant and considered an important contribution to existing knowledge.

TECHNICAL TRANSLATIONS: Information published in a foreign language considered to merit NASA distribution in English.

PUBLICAT1oNS: Information derived from or of value to NASA activities. Publications include final reports of major projects, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.

TECHNOLOGY UTILIZATION PUBLICATIONS: Information on technology used by NASA that may be of particular interest in commercial and other-non-aerospace applications. Publications include Tech Briefs, Technology Utilization Reports and Technology Surveys.

Details on the availability of these publications may be obtained from: ’

SCIENTIFIC AND TECHNICAL INFORMATION OFFICE

N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N Washington, D.C. 20546