Wind-tunnel investigation of highly maneuverable supersonic V/STOL fighter

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NASA Technical Memorandum 78599 N79- 26017 WIND-PUt4NFL INVts'IIGATION (NASA-TM-78 9 )9") .:" VFf ANi. :AIYFPICNIC V/STOL OF HIGHLY M R F F1C10 F@ (N,%SA) 24 j) VC 402/M1F U111 CSCI. C 1 A rl Tic 1 as G3/02 27794

Wind-Tunnel Investigation of

Highly Maneuverable Supersonic

VIISTOL Fighter

Michael Falarski

June ,979 i C7j r&a`t`jii ^y1; N

RUSA

National Aeron ,_, hcs and Spare Administration

__._-

WIND-TUNNEL INVESTIGATION 01' HIGHLY MANEUVERABLE SUPERSONIC V/.,TOL FIGHTER Michael Falarski Ames Research Center, NASA Moffett Field, California 94035 SL` ZIARY Thin paper presents a brief summary of the results from the initial wind- tunnel test of a large-scale, highly maneuverable supersonic V/STOL fighter model in the Ames 40- by 80-Foot Wind Tunnel. The STOL configuration which was tested combines p pper surface blowing and spanwise blowing to improve the lift characteristics over a wide angle-of-attazk range. A close-coupled canard was added to this configuration to create a highly maneuverable STOL aircraft. The 7.28 m (24 ft) span model is power a d by two .I-97 turbojet engines, each producing 9340 N (2200 lb) thrust at a pressure ratio of 2.

With the nozzle flap and aileron set at 30°, the model produced lift coefficien_a greater than 4. The model was longitudinally unstable beca.se of the forward canard position and because of the large body area of fuselage, strike, and nacelles forward of the center of gravity. The canard had a posi- tive interference effect on )oth lift and drag but was limited in its control effectiveness by stall. Spanwise blowing delayed wing stall and increased the linear portion of the lift curve. It did not significantly increase maximum lift, however, because of body stall.

INTRODUCTION In recent y ears, interest in military application of V/STOL technology has revived, espec'ally regarding the development of a highly maneuverable, supersonic fighter aircraft. Recent aircraft design studies (ref. 1) have detailed areas where technological uncertainties still exist which will impede the development of a V/STOL aircraft. Because the takeoff and approach per- formance is one of these areas of uncertainty, Ames has designed and fahri- cated it large-scale wind-tunnel model of a V/STO1. fighter to stud y low-speed aircraft characteristics. This model can be adapted to different V/STOL propulsion concepts.

This paper presents a brief summary of the results from the initial wind- tunnel test of this model adapted to the upper surface blowing propulsion con- cept. This Is combined with spanwise blowing (SWB) and closed-coupled canards to augment the lift over a wide angle-of-attack range. Results of wind-tunnel tests of a similar small-scale. model (refs. 2-3) have shown that this combina- tion produces a highl y maneuverable STOI. concept.

The complete model description and analysis of all the wind-tunnel results will be published at a later date in a NASA Technical Memorandum.

SYMBOLS b wing span, m (ft) C d drag coefficient, drag lift C l lift coefficient , s q circul ation lift CL circulation lift coefficient, - - qs roll momen t C roll moment coefficient, Z qs b CZ rate of change of C Z with 6 f3 ment pitching-moment coefficient, pitching mo q Sc g moment yawing-moment coefficient, 'Lin C.11 qs b Cn rate of change of C n with B h total axial gross thrust C thrust coefficient, T qs

s side force CY side-force coefficient, rate of changes of C with 6 C t3 c local wind chord, m (ft) F mean aerodynamic chord, m (ft) 'I wind-tunnel free-stream dynamic pressure, n/m (lb/1 S wing reference area, m (ft) Q wing angle of attack, deg sideslip angle, deg 6 i1 aileron deflection angle ref. to aileron centerline 6 c canard deflection angle, deg nozzle flap deflection angle ref. to flap centerlii 6 f Note: All moments referenced to leading; edge of c. F4 moments are stabilit y axes.

MODEL AND TEST DESCRIPTION The V/STOI. fighter model is shown installed in the test section of the Ames 40- b y 80 -Fort Wind Tunnel in figure 1. A sketch of the overall model and propulsion system geometries are presented in figure 2. The model is approximately 0.7 scale incorporating a 40° swept wing with pitching moment control provided by a close-coupled canard and beaver tail. For this investi- gation the canard was mounted in the forward position of the three available longitudinal canard positions. Limited test time did not allow for the inves- tigation of the canard leading- and trailing-edge flaps nor the beaver tail.

The model is powered b y two J-97 turbojet engines mounted in nacelles at 0.33 b/2 to provide a strake between the fuselage and nacelles for future intogration of a VTOL ejector propulsion system. The engine exhaust is pre- turned to 25° by the aspect ratio = 8, two-dimensional wedge nozzle. It is blown over the upper surface of the nozzle flap. The nozzle flap is capable of vectoring the exhaust from -10° to +30°. When tests were made with SWB, the flap nozzle area is reduced 171 and the SWB nozzle uncovered. The SWB nozzle is mounted flush with the outboard nacelle wall at 23% of the wing root chord. Both a circular and an aspect ratio = 4 rectangular nozzle have been designed for the model. The rectangular nozzle was used for this investigation.

The model has been instrrimented to measure: canard, strake, wing, and beaver-tail surface pressures; wing- and beaver-tail surface temperatures; exhaust tott:l pressure distribution at nozzle exit aril flap trailing edge; and engine duce flow properties before and after the SWB nozzle station. The sur- face instrunWntation in located on the model's left side and can he seen in figure 1 is she unpainted strips.

The model was investigated through angle of attack and sideslip ranges of -8° to +;0° and -10° to +30°, respectively, and thrust coefficients of approx- imateiy 0 to 2.0. The engines were operated at exhaust total pressurt, ratios of 1.8 and 2.0. Most of the data were taker: at ^.0 which produces a thrust per engine of 9341 N (2200 lb) and an exhaust temperature of 1100° F.

► he model is now undergoing a static thrust calibrattor. to determine exact thrust coefficients.

RESULTS AND DISCUSSION The principal objective of this first wind-tunnel test war: to investigate the longitudinal aerodyi.amics of the basic v/SToi. fighter configuration. and to briefly sample the canard and SWB effects and lateral-directional charac- be discussed first in detall followed

teristics. The longitudinal results will

of the lateral-directional data. The modal data matrix b y a brief discussion will be completed with two additional wind-tunnel entries in 1979.

Basic Model Longitudinal Characteristics The longitudinal characteristics of the model with the nozzle flap and aileron set at 30° and the canard at (1 0 is presented in figure 3. The lift curve slope is linear tip to a = 15° where the wins; stalls. The lift continues to increase after the wing stalls because of the lift generated by the boil'.

consisting of the strake, fuselape, and nacelles. Power increased C I,MAK and The 4 were "STALL but had little• effect on tite lift curve slopo.

Ci,MAX'- a 2.

achieved with C 1 1 The model pitching moment has been referenced to the leading edge of the mean aerodynamic chord, c, which destr;n studies have shown to he a reasonable location of the aircraft center of gravity. With this moment reference the model is longitudinally unstable with a negative static margin of X0.40.

Power does not change the nnrgin but does produce large negative moment shifts.

This large instability results from the lift generated by the canard, strake, nacelle, and fuselage forward to of the reference. Moving the canard the aft position will relieve this instability but may not reduce it to the -0.15 to -0.20 static margin desirable for modern aircraft control systems.

Canard Effects The effect of canard presence and its deflection are presented in fig- tires 4 and 5, respectively. The canard has a favorable effect on both lift

and drag. Wing stall is delayed and increased by approximately 0.3.

CIMAy Drag is reduced over the entire lift range with the effect larger at the high lifts. The principal adverse effect is the increased longitudinal instability.

Moving the canard aft would reduce this instability. This configuration will be investigated during the next wind-tunnel test.

The canard was effective in controlling pitching moment up to a combined canard deflection and angle of attack of 24°. Above 24° the canard stalled and the pitching moment returned to the undeflected valve. The use of the canard leading edge to delay stall and extend the canard usefulness will be investigated during; the next test.

Nozzle Flap Effects The data discussed up to now have been for a nozzle flap deflection of 30°. A limited amount of data was also recorded at a deflection of 0°. in deflected to the Same angle as the nozzle flap. A both cases the aileron were comparison of the two flap deflections with the canard removed is presented in figure 6. Reducing the flap deflection reduced the lift at a constant angle of attack and delayed stall allowing almost the same maximum lift to be achieved. Drag was reduced at low lifts and increased at the high lift. This flap change also produced a large positive shift in the zero-lift pitching moment.

Spativise Blowing (SWR) Effects The upper surface blowing concept is designed to enhance aircraft lift over a wide range of a by combining the jet flap effect to induce circula- tion lift, and vortex augmentation to delay wins; stall. The jet flap effect is created by the two-dimensional nozzle/flap, while the vortex augmentation is provided by the SWB. The SWB delays wing leading-edge vortex breakdown delaying wing stall (ref. 4). The SWi1 for the fighter model is provided by diverting 17% of the J-97 exhaust to a rectangular nozzle mounted flush with the outboard nacelle wall at 23% of the wing root chord. As can be seen in figure 7, SWB did delay stall and increased lift the high a. The maximum at lift and model stall were not significantly altered because they are controlled b y body stall. SWB also reduced drag and pitching moment at the high Iifts.

All of these effects were much more pronouncea at the high thrust coefficients.

To assist in understanding of the SWB effects, the wing surface pressures and temperatures are being analyzed. A typical example of these data is pre- sented in figures S and (). These data show SWR to significantl y increase both the aerodynamic load and temperature at the wing tip. The temperatur e data were severely limited by premature failure of the thermocouple probes result- ing from the adverse flaw environment. These probes will be modified for future: tests, and several dynamic transducers will be installed in the y wing to measure pressure fluctuations. The SWB data are :still being analyzed, but initial results indicate that incorporation of SWB into an aircraft will necessitate change in the wing structure to cope with the pressure and temper- ature environment.

Circulation Lift As previously indicated the model is designed to provide flaw circulation to enhance the aircraft lift. The model circulation lift, CL r , at a • 0 with and without SWB is presented in figure 10. The up l ,er surface blowing does benefit from circulation lift and SWH increases it even further. The CL, can amount to 15 to 30% of the total lift depending on the thrust coeffi- cient and the SWB. Without SWB the C Lr CTS increases very slowly above 0.8, while with SWB operation C L continues to increase as a result of increased circulation around the wing.

Lateral-Direction Characteristics The limited lateral-directional data recorded during this first test indicate no unusual characteristics (figs. 11-12). The basic mcdel, with the flaps and aileron deflected to 30°, and the canard at 0°, shows positive side force and lateral stability. It also shows neutral to slightly unstable directional stability at low a and Ei. Power has a small effect on the lateral and directional stability. Angle of attack has a destabilizing effect oil lateral stability while increasing directional stability. SWB had only small effects on the lateral-directional characteristics.

CONCLUSIONS call The following general conclusions drawn from the results of the initial investigation of the fighter model: An upper surface blowing concept applied to a V/STOI. fighter config- 1.

uration can produce maximum lift coefficients greater than 4.

2. Lift generated by the canards, strake, and fuselage area forward of the c.g. contributes to static instability, and for the configuration tested, resulted in a large unstable static margin.

3. Canards delay wing stall and .educe drag, but their control effec- tiveness is limited by stall.

4. The spanwise blowing delayed wing stall but did not significantly increase maximum lift because maximum lift is controlled by body stall.

a 5. At Y 0°, the upper surface blowing induces circulation lift that is about 15 to 20% of the total lift.

REFERFNCES 1. Lummus, J. R.: Study of Aerodynamic Technology for a V/STOL Fighter/ Attack Aircraft. NASA CR-152128, 1978.

2. Whitten, Perry U.: An Experimental Investigation of a Vec • tored-Fngine- Over-Wing Powered-Lift Concept. AFFUL-TR-76-92, 1978.

3. Bradley, R. G.; Jeffries, R. R.; and Capone. F. J.: A Vectored-Engine- Over-Wing Propulsive-Lift Concept. AIAA Paper 97-917, Sept. 1976.

4. Bradley, R. G.; and Wray, W. 0.: A Conceptual Study of Leading-Edge Vortex Enhancement by Blowing. AiAA J. of Aircraft, vol. II, .Ian. 1974.

(a) 3/4 rear view.

Figure 1.- V/STOL fighter mode] installed in wind tunnel.

%I'

K

8 ^^ ^• .

(h) 3/4 front view.

Figure 1.- Concluded.

2 13 17 001 243(7991 273(8961 FS 1165 7 - FS 2b0 250 -6 MA(^ DL 83 65 --^ / MAC BL '0 01 FS 75.23 B0 -- 728 ,_BEAVERTAIL 11199 8 I 123 881--T ^- --T 1.52 15 U01 FS 141 980 Ht 480 ^^ FS 375.0 t 09766250 I — 1 45 W H 1 69 1 .

0 48 11 40 - HL 118 90 0 77 +{ —BL 132 8 12.531 - -^ -0.91 (2 98) Bt 14330 WINGS 1ft 2 REFERENCE AREA, m 2 ) 17 00 1183 01 ASPECT RATIO 3.12 TAPER RATIO 0.238 AIRFOIL SECTION 64A204 FS 266 88 FS 290 733 GEOMETRIC TWIST -4 t MF ^,N At RO CHORD. m Ift1 233 17 661 187 1815) CANARD m 2 (1t 2 ) AREA, 543(5851 - CANARD AREA%ING AREA 0.32 ASPECT RATIO 0.05 10.15) FS 403 239 2.4 FS 0 TAPER RATIO 0.345 NACELLE AIRFOIL SECTION 64AO04 1 51 (494) ALL DIMENSIONS IN m Ift) MEAN AERO CHORD, m (ft) (a) V/STOL fighter mode- overall geometry.

Figure 2.- V/STOL fighter model geometry.

FS 342.67 FS 34 7 3 83 J 92 71 cm B1 A 36 50 in.

BL 69 — SPANWISE BLOWING I I FS 276 00 I NACELLE SKIN

—^^

AILERON J97\ ,2 D NOZZLE FLAP Z47.85 cm 1R 84 in.

WING PROFILE Al NACELLE SPANWISE BLOWING (h) Nozzle geometry.

Figure 2.- Concluded.

APPROX J Cr CL O 0 q 0.3 p 1.0 A 2.0 -10 10 20 30 40 o. deg CL 0 ' 1 I -16 -1.2 -8 .4 0 .4 8 12 16 8 4 0 -.4 8 -1.2 16 Co CMLE a Figure 3.- Basic model longitudinal characteristics; 6 f 6 a = 30°, 6 c = 0°, SWB off.

APPROX C CAk!ARD C L 2 O 2.0 ON 1 OFF O O 0 ON Q 1 OFF C -1 -10 0 10 20 30 a, deg

11 F

C, 2 I!

0 -A -.8 -1.2 -1.6 0 .4 .8 1.2 1.6 .8 .4 -1.6 -1.2 -,8 -.4 CU CMLE Figure 4.- Effect of canard on longitudinal characteristics; 6 f = 6 a 30°, 6 c = 0°, 5WB off.

APPKOX

C T

Nc, day ct

2.0 0

O

0 70

0 05 0

0 20

0 10 20 30 JA) -10 ,%, (ley U CL

8 12 1.6 2.0 N 4 0 4 8 17 16

t 0 4 -1.6 -1.2 .8

Cn

EE Figurt. , 5.- Fffect of canard deflection on It , ngItudin.11 rharacterist i, s;

S a - 30

1S f - , SWIi c'ff.

I •',

APPROX C T 61 . h , , deg

O 2.0 I

CL n 4 0

O 0.3 I 34

0 1

0 10 20 30 40 -10 Lk, deg CL -1 1.2 4 0 4 -.8 -1.2 -1.6 -.4 0 4 .8 -20 -1.6 -1.2 -.8 C MLE CO Figure 6.- [affect of flap/aileron deflection on longitudinal characteristics; canard and SWR off.

APPROX C T SWR CL O 2.0 ON 2 4 i OFF O 0.3 ON OFF -10 0 20 30 <,. deg CL -1.0 H -Lz -A 0 .8 1.2 1.6 b a 0 -.4 -1.2 -1.6 CD C MLE % Figure 7.- Effect of spanwise blowing on longitudinal characteristics!

00.

6 f = 6 f = 30 0 , d o =

-e -6 cp .3 -2 -1 .5 6 J 8 .9 1.0 b/2 8.— Spanwise wing pressure distribution with and without SWB; Figure a = 20°, 0.50c.

Q = 6 a = 30°, 6 c = 0

dey

A

O 0 0 0 0 5 / 4 C ^ T ^ 20 / "A- 300 / NACELLE WALL % b12 = 0.48 ` 050 ^y o / `1 w 200 7 w ^' r W O"" U G K D 100 0 1 1 I l I 5 6 7 9 e SEMISNAN,% Figure 9.- Wing surface temperaturts with A f SWH; = b a Q 30°, S C = 0°, a = 0°.

'%

SWB ON CSI.

OFF CL t ON .4 CSI.

OFF 0 4 8 1.6 1.2 APPROX C1 Figure 10.- Fighter model circulation lift; 6 f = 6 a = 30 0 , d = 0',a 0°.

c I ()

Q APPHOX C u, deg T O 20 0 0 0 • 2.0 120 CY0 -.2 -0011 CY -.3 - -0010 -•4 -0.0185 1.2 -.5 - -0031 .8 -.6 4 - 7 1 Ck Ckif -.4 C, 00028 00028 -.8 C n 0.0003 -1.2 - - -0.0006 -0.009 -0.0008 10 20 30 -20 -10 0 40 -20 -10 0 10 40 20 30 ,+, deg P, deg 6 a - 30 lateral -direction characteristics; 6 f - Figure 11.- Basic model 5c = 0. SWB off.

I,

SWB O ON q OFF C Y -.1 -.2 .4 -.3 C^ -.4 .04 -.e 00013 -1.2 -.04 -1.6 -20 -10 30 40 -20 -10 20 -- 30

ll. deg

0, (leg Figure 12.— F,ffect of SWB on lateral—directional characteristics;