Low-speed wind-tunnel test of a STOL supersonic-cruise fighter concept

NASA Technical Memorandum 4050

Low-Speed Wind-Tunnel Test

of a STOL Supersonic-Cruise

Fighter Concept

Paul L. Coe, Jr., and Donald R. Riley

JULY 1988

NASA Technical Memorandum 4050

Low-Speed Wind-Tunnel Test

of a STOL Supersonic-Cruise

Fighter Concept

Paul L. Coe, Jr., and Donald R. Riley

Langley Research Center Hampton, Virginia National Aeronautics and Space Administration Scientific and Technical Information Division Summary near the aircraft center of gravity. This arrange- ment is intended to allow large thrust-vector angles A wind-tunnel investigation was conducted to to be used to obtain significant values of powered lift examine the low-speed static stability and control while pitching-moment trim changes are minimized.

characteristics of a 0.10-scale model of a STOL A complete description of the concept and design supersonic-cruise fighter concept. The concept, re- rationale is provided in reference 1.

ferred to as a twin-boom fighter, was designed as a The purpose of the present investigation was to STOL aircraft capable of efficient long-range super- obtain low-speed stability and control information sonic cruise. The configuration name is derived from over an angle-of-attack range including the stall. The the long twin booms extending aft of the engine to investigation included a study of jet-induced power the twin vertical tails, which support a high center effects.

horizontal tail. The configuration propulsion system features a two-dimensional thrust-vectoring exhaust Symbols nozzle which is located so that the nozzle hinge line Longitudinal forces and moments are referred to is near the aircraft center of gravity. This arrange- the wind-axis system and lateral-directional forces ment is intended to allow large thrust-vector angles and moments are referred to the body-axis system.

to be used to obtain significant values of powered lift (See fig. 1.) The moment reference center is on the while pitching-moment trim changes are minimized.

fuselage centerline and, unless otherwise indicated, is The results of the investigation indicated that the located longitudinally at 25 percent of the mean aero- configuration exhibited significant nonlinear aero- dynamic chord of the reference wing. This location dynamic characteristics. Specifically, the results is 0.1181 forward of the nozzle flap hinge line. The showed that the configuration exhibited a pitch-up coefficients and symbols used are defined as follows.

tendency at an angle of attack of 8'. Deflection of the forward portion of the inboard wing panels (includ- b wing span, ft ing the engine inlets) delayed the onset of pitch-up CD drag coefficient, Drag/qS to an angle of attack of 10'.

CL lift coefficient, Lift/qS Downward deflection of the jet exhaust nozzle (for thrust vectoring) produced additional circulation c l rolling-moment coefficient, lift. Furthermore, with the horizontal tail off, thrust Rolling moment/qSb vectoring resulted in minimal changes in pitching Crn pitching-moment coefficient, moment.

Pitching moment/qS~ Because of high levels of thrust-induced down- wash, the horizontal tail was ineffective for providing c n yawing-moment coefficient, longitudinal stability. Similar results were found for Yawing moment/qSb conditions with a revised horizontal tail repositioned CT thrust coefficient, to locations outboard of the aft tail booms. The hori- Wind-off thrust force/qS zontal tail remained effective for longitudinal control but was deficient in pitch trim capability. Finally, CY side-force coefficient, Side force/qS the configuration experienced a marked reduction in mean aerodynamic chord of refer- lateral and directional stability for angles of attack ence wing, ft (see fig. 2) above 20'.

it horizontal-tail incidence angle, deg Introduction leading-edge segment of outboard L 1 wing panel (see fig. 2(b)) A wind-tunnel investigation was conducted to examine the low-speed static stability and control forward portion of inboard wing L2 characteristics of a 0.10-scale model of a STOL panel (see fig. 2(b)) supersonic-cruise fighter concept. The concept, re- free-stream dynamic pressure, lb/ft ferred to as a twin-boom fighter, was designed as S a STOL aircraft capable of efficient long-range su- planform area of reference wing, ft2 personic cruise. The configuration name is derived (see fig. 2(b)) from the long twin booms extending aft of the engine t17t2,t3 trailing-edge-flap segments of wing to the twin vertical tails, which support a high cen- (inboard, mid, and outboard, ter horizontal tail. The propulsion system features respectively) a two-dimensional thrust-vectoring exhaust nozzle body reference axes (see fig. 1) which is located so that the nozzle hinge line is x, y, a angle of attack of aircraft reference vertically above the centerline of the tunnel and was line, deg moved along an arc as angle of attack was varied.

(See fig. 3.) Because the model was relatively small P angle of sideslip, deg compared with the test section, no corrections have aileron (segment t 3 ) deflection, deg been applied to the data.

6 , 6, elevator deflection, deg Tests 6f streamwise deflection angle of The tests were conducted at a free-stream velocity trailing-edge flap, deg of 50.23 ft/sec (q = 3 psf), which corresponds to a Reynolds number of 0.58 x lo6 based on the mean 6~~ deflection angle of leading-edge segment of outboard wing panel, aerodynamic chord of the reference wing. Static- force tests were conducted over angles of attack from deg -8' to 32' and angles of sideslip of -5' and 5'.

deflection angle of forward portion ' ~ 2 Principal configuration variables included control- of inboard wing panel, deg surface deflections and model thrust settings.

deflection angle of centerline nozzle, 6n Presentation of Results deg The results and discussion are presented as rudder deflection normal to hinge 6, follows: line, deg Figure E downwash angle, deg Longitudinal aerodynamic characteristics:

Baseline configuration . . . . . . . . . . . 4

Abbreviation: Effect of deflection of wing-panel c.g. center of gravity

leading edge . . . . . . . . . . . 5 and 6

Derivatives: Effect of trailing-edge

deflection . . . . . . . . . . . . . . . 7

Effect of thrust . . . . . . . . . . . 8 to 13 Elevator effectiveness . . . . . . . 14 and 15 Model Longitudinal trim . . . . . . . . . . . 16 Effectiveness of revised The investigation was conducted with the 0.10-

horizontal tail . . . . . . . . . . 17 to 20

scale model of a STOL supersonic-cruise fighter con- cept shown in figure 2(a). In addition to the base- Lateral-directional aerodynamic line configuration, limited tests were also conducted characteristics: to determine the effectiveness of the revised horizon- Static lateral-directional tal tail shown in figure 2(d). Jet power effects were

stability . . . . . . . . . . . . . . . 21

simulated through use of model ejectors which were Rudder effectiveness . . . . . . . . . . . 22 supplied by compressed air. Aerodynamic forces and

Aileron effectiveness . . . . . . . . . . . 23

moments were measured with a conventional six- component strain-gage balance that was mounted Results and Discussion internal to the model. The model movable sur- faces included the exhaust nozzle, wing leading- and Longitudinal Aerodynamic Characteristics trailing-edge flaps, twin rudders, and a single eleva- tor. Surface deflections were obtained with preset Baseline configuration. Longitudinal aerodynamic angle brackets. All gaps were sealed for the test data characteristics for the baseline configuration with un- presented.

deflected leading- and trailing-edge flaps and unde- flected exhaust nozzle are presented in figure 4 for Tunnel and Apparatus the horizontal tail off and on. As shown, for the The investigation was conducted in the Langley tail-off configuration a mild pitch-up is observed near cr = 8'. This pitch-up typically results from lift gen- 12-Foot Low-Speed Tunnel. The model and inter- erated on the forward portion of the wing by leading- nal strain-gage balance were mounted on a sting at- tached to a curved strut system so angle of attack edge vortices. If sufficient pitch control is available, artificial stability methods such as alpha feedback could be varied. Sideslip angle was varied by rota- may be a means of stabilizing the configuration. In tion of the curved strut about a vertical axis. To en- the present study, configuration modifications were sure sufficient tail clearance, the model was located considered in an attempt to minimize or forestall this produced by trailing-edge flap deflection indicates pitch-up tendency. that for the conditions considered, the wing trailing- edge flaps begin to stall at angles of attack of about In figure 4 it is shown that for the 16'. Unlike the exhaust nozzle, the wing trailing- horizontal-tail-on configuration the pitch-up char- edge flap hinge line is relatively far rearward of the acteristic is somewhat more pronounced than that moment reference center. Wing trailing-edge deflec- for the corresponding tail-off configuration. As dis- tion results in an appreciable negative increment in cussed in a subsequent section, for the conditions pitching moment, and the stall of the trailing-edge under consideration, this loss of horizontal-tail effec- flaps aggravates the previously discussed pitch-up tiveness for angles of attack greater than 8' is re- tendency.

lated to a significant nonlinear increase in the down- wash angle E , and thus an increase in d e l a a , at the Effect of thrust. Wind-off static turning data horizontal-tail location. Similar variations of down- for the configuration with an undeflected exhaust wash angle with increasing angle of attack have been nozzle (6, = 0') and with a deflected exhaust nozzle noted for T-tail aircraft configurations in which the (6, = 42.15') are presented in figure 8. Analysis high horizontal tail moves from a region of relatively of the data indicates that the thrust-vector angle is slight downwash influence to a region of strong down- aligned with the exhaust nozzle deflection 6,.

wash influence with increasing angle of attack. The Figure 9 shows the effect of thrust on the static data of figure 4 also suggest that the basic airframe longitudinal aerodynamic characteristics of the con- may experience a deep-stall stable trim point near figuration with 6% = 0'. Also presented in figure 9(a) CY = 36'.

are the computed values of CL based on the relation- ship Effect of deflection of wing leading edge. The results I of deflecting the outboard-wing-panel leading-edge segment L1 (see fig. 2(b)) of the present configuration are presented in figure 5. As shown, deflecting which assumes only a direct thrust contribution to these particular leading-edge segments is ineffective lift. As shown, no appreciable additional circula- in altering the nonlinearity of Cm. Inasmuch as the tion lift is produced and the lift and drag data show separated leading-edge vortex system is considered essentially the anticipated results. The data of fig- to originate well forward of the outboard-wing-panel ure 9(a) show very little effect of thrust on pitch- leading edges, this result is expected.

ing moment, as would be expected because of the In a further attempt to forestall the formation of thrust vector passing through the moment reference separated wing leading-edge vortices, the entire for- center. However, for the configuration with.the hor- ward portion of the inboard-wing-panel (segment L2, izontal tail on (fig. 9(b)), the pitching-moment data see fig. 2(b)), including the engine inlet, was deflected show an unexpected positive increment associated downward. Figure 6 presents data for the baseline with thrust. This result suggests the possibility of a cruise wing with and without this forward inboard thrust-induced loading brought about by downwash wing panel deflected 20'. As shown, deflection of acting on the horizontal tail. To more clearly il- the forward portion of the inboard wing panel delays lustrate this thrust-induced downwash, the pitching- the occurrence of pitch-up to CY z 10' and delays moment data of figure 9 are replotted in figure 10.

wing stall. Figure 6 also shows that increasing the As shown, thrust results in increasingly positive deflection to 30' results in no further improvement increments in the contribution of the horizontal tail in the pitching-moment characteristics and results in to Cm.

degraded aerodynamic performance.

Figure 11 shows the effect of thrust on the static longitudinal aerodynamic characteristics for Trailing-edge effectiveness. Data for the effect the high-lift configuration with 6, = 42.15' and of deflecting the exhaust nozzle from 6 , = 0' to bf = 20'. Also presented in figure l l ( a ) are 6, = 42.15' (for the unpowered condition) are pre- computed values of CL based on the relationship sented in figure 7. Also presented in figure 7 arc data showing the effect of deflecting the wing trailing-edge flap in conjunction with the unpowered deflected ex- haust nozzle. The data show that in this unpowered condition, deflecting the exhaust nozzle results in a which, as previously mentioned, assumes only a di- slight positive increment in Cm. This result indi- rect thrust contribution t o lift. As shown, significant cates that the center of pressure of the additional lift levels of additional circulation lift are produced for due to nozzle deflection is forward of the moment the configuration with 6, = 42.15' and bf = 20'.

reference center. The incremental pitching moment Based on the theory presented in reference 2, the level Longitudinal trim. Examination of the data of of additional circulation lift achieved in the present figure 15 shows that the configuration is only slightly tests is about 75 percent of the theoretical maximum. stable over angles of attack of -8' < a < 8' for In figure l l ( a ) it is shown that, for the horizontal 6, = 0'. The data of figure 15 also show that tail off, increasing thrust results in increasing nega- for a > 8' or for Se # 0°, the configuration is tive increments in pitching-moment coefficient. For longitudinally unstable. It is of course recognized the horizontal tail on, the data of figure l l ( b ) indi- that active controls in the form of alpha feedback cate that a t low angles of attack the magnitude of to the horizontal tail may be a possible means for this thrust-induced negative increment in Cm is re- providing longitudinal stability, provided sufficient duced, while at higher angles of attack it appears control power is available.

that thrust results in a positive increment in Cm.

Figure 16 presents the regions of a vs CT In order to isolate the direct-thrust effects from the and CL vs CT for which (based on the data of preceding pitching-moment results, the data of fig- fig. 15) sufficient longitudinal control exists to achieve ure 11 have been recomputed for a moment reference Cm = 0. It is of course recognized that a compre- center corresponding to the exhaust nozzle hinge line hensive analysis including other pitch control require- (0.368~) and are presented in figure 12. As shown in ments and static margin requirements would be nec- figure 12(a) (horizontal tail off), for angles of attack essary to determine if, in fact, unstable trim can be less than about a = 16' thrust has only a minimal achieved for the regions presented in figure 16. How- effect on Cm, as expected. However, for the hor- ever, the results summarized in figure 16 do serve to izontal tail on, the data of figure 12(b) show that illustrate that the configuration is deficient in pitch thrust produces a positive increment in Cm. In or- trim capability.

der to illustrate this point more clearly, the pitching- moment data of figure 12 are replotted in figure 13. Revised horizontal tail. As noted previously, As shown in figure 13, thrust results in a positive for the high-lift configuration with S j = 20' and increment in the horizontal-tail contribution to Cm. 6, = 42.15', the high center tail is subject to a sub- This result is undoubtedly associated with a thrust- stantial thrust-induced downwash which results in induced downwash, which produces a downward load the stall of the horizontal tail for 6, = -20' and on the horizontal tail. It should be further noted that in a reduction in the horizontal-tail contribution to for a 5 go, thrust reduces the horizontal-tail contri- longitudinal stability. In an attempt to alleviate this bution to static longitudinal stability. Although no condition, an alternate low outboard tail was con- downwash measurements were made, the reduction sidered. (See fig. 2(d) for geometric details.) The in horizontal-tail contribution to longitudinal sta- effects of thrust for this high-lift configuration with bility is thought to be related to a thrust-induced the modified horizontal tail off and on are presented in figure 17 for a design center-of-gravity location of increase in the downwash factor d ~ l d a .

0.25E and in figure 18 for a center of gravity coin- cident with the exhaust nozzle hinge line located at Elevator eflectiveness. Data showing elevator 0.368E. As discussed previously, this latter center of effectiveness for 6, = 0' and S j = 0' and for gravity was chosen in order to eliminate the direct- 6, = 42.15' and S j = 20' are presented in figures 14 thrust-vector component from the pitching-moment and 15, respectively. As shown in figure 14, for the coefficient. The pitching-moment data of figure 17 undeflected exhaust nozzle (6, = 0') the elevator ef- show increasing thrust results in significant negative fectiveness remains fairly constant over the test range pitching moments for the revised outboard tail con- of angles of attack and is not significantly affected figuration with a moment reference center of gravity by thrust. The data demonstrate that for the con- located at 0.25c. However, the data of figure 18 in- ditions under consideration, the horizontal tail and dicate that the low outboard tail location is subject elevator are not stalled except for elevator deflec- to a downward load as was the high center tail, and tion angles S, of magnitudes greater than about 20'.

hence a high level of thrust-induced downwash also In the high-lift configuration with Sn = 42.15' and exists for the outboard tail location.

S f = 20' (fig. 15), the data show that for the unpow- ered condition (CT = 0) the elevator effectiveness is Control effectiveness for the all-moving revised horizontal tail is presented in figure 19. Analysis similar to the data for the cruise configuration (see fig. 14). For the power-on conditions (CT = 0.97 and of the data of figure 19 shows that for the unpow- CT = 2.02), however, it appears that the horizontal ered condition, the horizontal-tail downwash factor tail stalls for 6, = -20'. This result is probably re- d s l d a increases from about 0.5 for a < 4' to about lated to the high levels of downwash which, as previ- 1.0 for 4 < a < 16. Therefore, as shown, the revised ously discussed, are associated with thrust vectoring. horizontal tail does not contribute to longitudinal stability for a > 4'. Furthermore, the revised hor- of the wing or the.occurrence of adverse sidewash or izontal tail does not alleviate the previously men- both. As expected, based on the previous discussion tioned deficiency in pitch-trim capability. of horizontal-tail effectiveness, power effects appear From the preceding results it is apparent that to increase the flow over the vertical-tail surfaces.

to utilize the high-lift capability afforded by the For the power-on conditions investigated, the level of thrust-vectoring system, an increase in horizontal- directional stability remains constant for -8" < a < tail control power is required for pitch trim. One 20'. Above a = 20°, a marked reduction in Cnp is possible configuration would include both the high noted, and instability occurs at a = 24'.

center tail (to maintain the structural advantage of Rudder effectiveness. Rudder effectiveness is de- the center tail section) in combination with the low termined from data for b , = f 23' and is presented outboard tail panels. Analysis of the data, with in figure 22. Based on the preceding discussion of the assumption that the incremental results for the directional stability, the variation of Cn6T with a is high center horizontal tail are linearly additive to expected. For the unpowered configuration there is those for the low outboard horizontal tail, results a degradation in rudder effectiveness with increasing in the longitudinal control effectiveness depicted in o; again, this degradation is thought to be due to figure 20. As shown in figure 20, this assumed the progressive immersion of the rudder in the wake configuration is unstable; however, the region for of the wing. Power effects apparently provide an in- which unstable trim could be achieved by horizontal- crease in flow over the tail surfaces which, in addition tail alpha feedback is greatly expanded. It is of to increasing CnD, results in increased levels of Cn6, course recognized that other alternatives exist. One up to a % 20'.

such alternative is to develop a high-lift horizontal tail mounted atop vertical tails of increased height.

Aileron effectiveness. Aileron effectiveness is de- The increase in vertical-tail height could position the termined from data for ba = f 20' and is presented horizontal tail in a more favorable downwash field in figure 23. These data were obtained for the cruise and thereby achieve both longitudinal stability and configuration (i.e., flaps and nozzle undeflected). The control.

data show that for the conditions investigated the Lateral-Directional Aerodynamic aileron effectiveness is maximum for a = 0' and the Characteristics aileron effectiveness is significantly reduced as the angle of attack is increased to the point at which Static lateral-directional stability characteris- leading-edge separation occurs.

tics. Static lateral-directional stability characteris- tics were determined for P = f 5' for the complete

configuration with bj = 20' and 6 , = 42.15' and are Summary of Results

presented in figure 21. The effective dihedral deriva- A wind-tunnel investigation was conducted in the tive C10 typically provides an accurate appraisal of Langley 12-Foot Low-Speed Tunnel to examine the the wing flow conditions. For the unpowered con- low-speed static stability and control characteristics figuration, the occurrence of vortex separation at of a 0.10-scale model of a STOL supersonic-cruise o % 4' is characterized by the nonlinear break in fighter concept. The results of the investigation indi- C10 data; the breakdown of the vortex core on the cate that the configuration exhibits significant non- advancing side of the sideslipped wing apparently be- linear aerodynamic characteristics. Specific results gins at a x 17' and is characterized by the break are summarized as follows: and subsequent positive slope of CLO; finally, the breakdown of both vortex cores apparently occurs at 1. The configuration exhibited a pitch-up o F Z 25' and is characterized by the second reversal tendency at an angle of attack a of 8'.

in ClO. Of course, the vertical tail also contributes 2. Deflection of the forward portion of the to the effective dihedral. For the conditions consid- inboard wing panels (including the engine in- ered, the reduction in C y with increasing angles of lets) delayed the onset of the pitch-up to attack indicates that the vertical-tail contribution to a = 10'.

C1, decreases with increasing angles of attack.

3. Downward deflection of the jet exhaust noz- The directional stability derivative Cn8 is stable zle for thrust vectoring produced additional up to a = 22' but shows a marked reduction as a circulation lift.

increases from 6 ' to 20' for CT = 0. This condition is typically related to the progressive immersion of the 4. For the horizontal tail off, thrust vectoring re- vertical tails in the reduced-dynamic-pressure wake sulted in minimal changes in pitching moment.

5. Significant levels of thrust-induced downwash NASA Langley Research Center Hampton, Virginia 23665-5225 acted on the horizontal tail.

May 27, 1988 6. The horizontal tail remained effective for lon- References gitudinal control but was ineffective for pro- 1. Dollyhigh, Samuel M.; Foss, Willard E., Jr.; Morris, viding longitudinal stability and was deficient Shelby J., Jr.; Walkley, Kenneth B.; Swanson, E. E.; in pitch-trim capability.

and Robins, A. Warner: Development and Analysis of a STOL Supersonic Cruise Fighter Concept. NASA 7. The configuration experienced a marked re- TM-85777, 1984.

duction in lateral and directional stability for 2. McCormick, Barnes W., Jr.: Aerodynamics of V/STOL a > 20°.

Flight. Academic Press, Inc., 1967.

Figure 1. System of axes and angular notation.

0.25 C Airplane ref. line-, \

-

(a) Three-view sketch of baseline model.

Figure 2. Geometric characteristics of model. Linear dimensions are in inches.

Area

Wing (ref. area), in2. . 975.24

Flaps (per side): t,, i n , . . . . . . . . . 28.50 Z t2, in . . . . . . . . . 22.1 3 t3, in . . . . . . . . . 23.00 L,, in,. . . . . . . . . 13.80 (b) Details of baseline wing.

Figure 2. Continued.

Horizontal-tailarea = 1 16.28 in Elevator area = 79.1 6 in

- - - - - - - -

( c ) Details of baseline horizontal tail.

Figure 2. Continued.

(d) Details of revised horizontal tail.

Figure 2. Concluded.

11- 40' wedge

Tunnel G Figure 3. Model and support system as mounted in wind tunnel.

Horizontal tail 0 Off On; 1 5 ~ = 0" Figure 4. Longitudinal aerodynamic characteristics of baseline configuration. b1 = 0'; f i L 2 = 0'; 9 = 0'; 6 , = 0'; CT = 0; c.g. at 0.25E.

Figure 5. Effect of deflection of outboard-wing-panel leading edge. Horizontal tail off; c.g. a t 0.25c; SL, = 0'; Sf = o O ; 6 , = o O ; cT = o .

Figure 6. Effect of deflection of forward portion of inboard wing panel. Horizontal tail off; c.g. a t 0.25E; bL1 = 0'; Sf = 0'; 6 , = 0'; CT = 0.

12 L2 LO LO . 8 .8 . 6 . 6 CL .4 CL .4 . 2 . 2 0 0 - . 2 - . 2 -.4 -.4 -12 -8 -4 0 4 8 12 16 20 24 28 32 3 6 . 3 . 2 . I 0 - . 1 - . 2 0 . 2 4 . 6 . 8

a, deg c , CD

Figure 7. Effect of trailing-edge deflection. Horizontal tail off; c.g. at 0.25c; SL, = 0'; SL, = 0'; CT = 0.

I I I 1 I I

0 10 20 3 0 40 50 -Axial force, Ib Figure 8. Results from static turning tests.

a, deg Cm (a) Horizontal tail off.

Figure 9. Effect of thrust on longitudinal aerodynamic characteristics with 6, = 0'. bL, = 0'; sL, = OO; Sf = 0'; c.g. at 0.25E.

-12 -8 - 4 0 4 8 12 16 20 24 28 32 36 a, deg (b) Horizontal tail on; 6 , = 0'.

Figure 9. Concluded.

C T = o

Horizontal tail

0 Off

On; = 0"

CT =0.9

Horizontal tail

0 Off

On; = 0"

a , deg

Figure 10. Effect of thrust on horizontal-tail contribution to C, with 6 , = 0'. bL, = 0'; 6L, = 0'; bf = 0'; c.g. at 0.25E.

(a) Horizontal tail off.

Figure 11. Effect of thrust on longitudinal aerodynamic characteristics for c.g. at 0.25E. liL, = 0'; liL2 = 0'; l i j = 20'; 6, = 42.15'.

(b) Horizontal tail on; 6 , = 0'.

Figure 11. Concluded.

(a) Horizontal tail off.

Figure 12. Effect of thrust on longitudinal aerodynamic characteristics for c.g. a t 0.368E. SL, = 0'; SL, = 0'; Sf = 20'; 6, = 42.15'.

(b) Horizontal tail on.

Figure 12. Concluded.

CT = o Horizontal tail 0 Off On CT = 0.47 Horizontal tail 0 Off On CT = 0.97 Horizontal tail 0 Off On CT =2.02 Horizontal tail 0 Off On Figure 13.

Effect of thrust on horizontal-tail contribution to C, with 6 , = 42.15'. SL, = 0'; bL2 = 0'; Sj = 20'; c.g. at 0.368E.

(a) CT = 0.

Figure 14. Elevator effectiveness with S f = 0 ' and Sn = 0'. SL1 = 0'; 6 L 2 = 0'; c.g. a t 0.25E.

(b) CT = 0.90.

Figure 14. Concluded.

(a) CT = 0.

Figure 15. Elevator effectiveness with Sf = 20' and 6, = 42.15'. SLl = 0'; SL, = 0'; c.g. a t 0.25E.

(b) CT = 0.97.

Figure 15. Continued.

Figure 15. Concluded.

(a) a vs CT.

(b) CL VS CT.

Figure 16. Trim region for configuration with SL, = O0 and 6L2 = 0'. 6f = 20°; 6, = 42.1S0; c.g. at 0.25~.

(a) Horizontal tail off.

Figure 17. Effect of thrust on longitudinal aerodynamic characteristics with SL, = 2 0 ' for c.g. at 0.25E.

SL1 = 0 ' ; Sf = 2 0 ' ; Sn = 4 2 . 1 5 ' .

(b) Revised outboard horizontal tail on.

Figure 17. Concluded.

(a) Horizontal tail off.

Figure 18. Effect of thrust on longitudinal aerodynamic characteristics with SL2 = 20' for c.g. at 0.368E.

SL1 = 0'; Sf = 20'; 6 , = 42.15'.

(b) Revised outboard horizontal tail on.

Figure 18. Concluded.

i t , deg 0 Off 0 -20 A 20 (a) CT = 0.

Figure 19. Effectiveness of revised outboard horizontal tail. c.g. at 0.25C; SL, = 0'; SL, = 20°; Sf = 20°; 6, = 42.15'.

Bap ' w 96 ZE 8Z fZ O Z 91 Z T 8 t 0 t- 8- ZT- 2 ' f' 9 ' 8' OT 2 ' 1 t l OZ z.z tZ OZ v oz- 0 3 3 0 0 Sap '?!

i t , deg 0 Off 0 -20 A 20 (c) cT = 2.02.

Figure 19. Concluded.

Figure 20. Calculated control effectiveness for high center (baseline) horizontal tail in combination with revised outboard horizontal tail (data of fig. 15 linearly added to data of fig. 19). SL, = 0'; SL, = 20'; 6, = 42.15'; Sf = 20'.

Figure 21. Lateral-directional stability derivatives versus angle of attack. Complete configuration; SL, = 0'; SL2 = 0'; Sf = 20'; Sn = 42.15'.

Figure 22. Rudder effectiveness for complete configuration. SLl = 0 ' ; SL2 = 0 ' ; Sf = 2 0 ' ; 6 , = 4 2 . 1 5 ' ; c.g. at 0.25E.

Figure 23. Aileron effectiveness for complete configuration. SLl = 0 ' ; SL, = 0 ' ; Sf = 2 0 ' ; Sn = 4 2 . 1 5 ' ; c.g. at 0.25E.

Report Documentation Page S ~ a c e Adm~n~slral~on

1. Report No. 1 2. Government Accession No. 1 3. Recipient's Catalog No. I

NASA TM-4050 5. Report Date 4. Title and Subtitle Low-Speed Wind-Tunnel Test of a STOL Supersonic-Cruise Fighter Concept 6. Performing Organization Code 7. Author(s) 8. Performing Organization Report No.

Paul L. Coe, Jr., and Donald R. Riley 10. Work Unit No.

9. Performing Organization Name and Address 505-61-71-07 NASA Langley Research Center 11. Contract or Grant No.

Hampton, VA 23665-5225 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546-0001 I 15. Supplementary Notes 16. Abstract A wind-tunnel investigation was conducted to examine the low-speed static stability and control characteristics of a 0.10-scale model of a STOL fighter concept capable of efficient long-range supersonic cruise. The configuration has long twin booms extending aft of the engine to the twin vertical tails, which support a high center horizontal tail. The propulsion system features a two- dimensional thrust-vectoring exhaust nozzle which is located so that the nozzle hinge line is near the aircraft center of gravity. This arrangement is intended to allow large thrust-vector angles to be used to obtain significant values of powered lift while pitching-moment trim changes are minimized.

The investigation obtained low-speed stability and control information over an angle-of-attack range including the stall. The investigation also included a study of jet-induced power effects.

18. Distribution Statement 17. Key Words (Suggested by Authors(s)) STOL Unclassified-Unlimited Supersonic cruise Vectored thrust Subject Category 08 22. Price 21. No. of Pages 20. Security Classif.(of this page) 19. Security Classif.(of this report) A03 42 Unclassified Unclassified NASA-Langley, 1988 NASA FORM 1626 OCT 86 For sale by t h e National Technical Information Service, Springfield, Virginia 22161-2171 National Aeronautics and Space Administration BULK RATE Code NTT-4 POSTAGE & FEES PAID NASA Washington, D.C.

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