NASA-TN-D-6826-1972 Wind-tunnel investigation at low speeds of a model of the Kestrel (XV-6A) vectored-trust V STOL airplane《Kestrel(XV-6A)矢量信托V STOL飞机模型的低速风洞研究》.pdf

《NASA-TN-D-6826-1972 Wind-tunnel investigation at low speeds of a model of the Kestrel (XV-6A) vectored-trust V STOL airplane《Kestrel(XV-6A)矢量信托V STOL飞机模型的低速风洞研究》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-6826-1972 Wind-tunnel investigation at low speeds of a model of the Kestrel (XV-6A) vectored-trust V STOL airplane《Kestrel(XV-6A)矢量信托V STOL飞机模型的低速风洞研究》.pdf（179页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICAL NOTE NASA TN 0-6826 -e cv Qo Y) n z c WIND-TUNNEL INVESTIGATION AT LOW SPEEDS OF A MODEL OF THE KESTREL (XV-6A) VECTORED-THRUST V/STOL AIRPLANE by Ricburd J. Murgason, Ruymond D, Vog-ley, und Mutthew M. Winston LuPzgley Reseurch Center Hdmpton, vu. 23365 NATIONAL AERONAUTICS AND SPACE

2、 ADMINISTRATION * WASHINGTON, D. C. 0 JULY 1972 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WIND-TUNNEL INVESTIGATION AT LOW SPEEDS OF A MODEL OF THE KESTREL (XV-6A) VECTORED-THRUST V/STOL AIRPLANE By Richard J. Margason, Raymond D. Vogler, and M

3、atthew M. Winston Langley Research Center SUMMARY An investigation was made in the 5.18-meter (17-foot) test section of the Langley 300-MPH 7- by 10-foot tunnel to determine the longitudinal and lateral characteristics of a 1/6-scale model of the Kestrel (XV-6A) vectored-thrust V/STOL airplane at lo

4、w-speed cruise and transition conditions. Data were obtained out of and in ground effect over a moving ground plane for a range of model angles of attack and sideslip at various thrust coefficients by using compressed air ejecting from nozzles in the fuselage. In the cruise configuration, the model

5、is longitudinally stable, but in transition, model instability is increased as power is increased, and the flaps and horizontal tail have little effect at high thrust coefficients (low forward speeds). At high thrust condi- tions in transition, the model is unstable in roll and yaw at small sideslip

6、 angles. Jet- free-stream interference generally results in an increase in drag and noseup pitching moments and a reduction in lift. Ground proximity generally reduces the interference effects on lift and pitching moment but has little effect on drag. Deflected thrust at speeds above transition prod

7、uces increments of lift and large deceleration forces useful in maneuvering flight, but with a reduction in stability. INTRODUCTION The National Aeronautics and Space Administration has provided a large body of detailed information on various vertical take-off and landing (VTOL) concepts and has rec

8、ently emphasized the importance of continued development in this area of aeronautics. - Detailed aerodynamic data have been obtained for several configurations such as the ducted propeller (ref. l), tilt wing (refs. 2 and 3), lift jet (ref. 4), and lift fan (ref. 5), some of which have evolved into

9、flight vehicles. One V/STOL airplane now operational is the Harrier (AV-8A) vectored-thrust fighter. The present paper supplements previous investigations on the prototype Hawker Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(P-1127) version of thi

10、s airplane, first flown in the early 1960s. The previous works include the results of flight tests of a 1/6-scale model (ref. 6) and the results of static- force tests at high subsonic speeds (refs. 7 and 8). The present investigation utilizes a 1/6-scale model of the Kestrel (XV-6A) version of this

11、 airplane. This version was used by the United Kingdom (U.K.), Federal Republic of Germany (F.R.G.), and United States (U.S.) in tripartite evaluations conducted during the mid- 1960s. Since those evaluations, two of these Kestrels were loaned to NASA Langley Research Center for flight tests (ref. 9

12、). Completing the evolution of this vectored-thrust V/STOL fighter to operational status is the Harrier (AV-8A), the third version of this airplane which is currently being flown by both the U.K. and U.S. military forces. . Both the Kestrel and Harrier were developed largely through modifications of

13、 earlier versions of this airplane. As a result, very few wind-tunnel data are available on either airplane. The present investigation of the Kestrel (XV-6A) was undertaken to provide static longitudinal and lateral characteristics with a wind-tunnel model to com- plement the flight-test results obt

14、ained by both the NASA and the USAF (ref. 10). The principal purpose of these flight tests has been the determination of the V/STOL transition-flight characteristics between hover and wingborne flight. Correspondingly, most data herein pertain to this speed regime. These tests were made through angl

15、e-of-attack and angle-of-sideslip ranges at sev- eral thrust conditions, flap deflections, and nozzle deflections. Also included are results of tests made for several heights above a moving ground plane to determine the effects of ground proximity in transition flight. Interference effects between t

16、he jets and the free stream are also presented. Recently considerable interest (ref. 11) has been expressed regarding the use of deflected thrust for situations not related to take-off or landing. One area of interest is its use to increase airplane maneuverability. Some of the recent flight tests c

17、onducted at Langley with the Kestrel (XV-6A) airplane have used deflected thrust at an altitude of approximately 4572 meters (15 000 feet). To complement this flight experience, some wind-tunnel data are included herein to describe the static aerodynamic characteristics of the plane in these maneuve

18、rs. SYMBOLS The longitudinal force and moment data are referred to the stability-axis system. - The lateral-directional data are referred to the body-axis system. The coefficients include jet and inlet momentum effects. center shown in figure 1. The units of measurement used in this paper are given

19、in both the International System of Units (SI) (ref. 12) and U.S. Customary Units. The measure- ments and calculations were made in U.S. Customary Units. For all data, the origin is located at the moment 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-

20、,-,-b cD cL Cm Cn CT CY - C D De h it L MY M, wing span, m (ft) D drag coefficient, - qms L lift coefficient, - qoos Rolling moment rolling-moment coefficient, q,Sb MY q,sz;. pitching-moment coefficient, - Yawing moment yawing-moment coefficient, q,Sb rn 1 thrust coefficient, - q,s Side force side-f

21、orce coefficient, %as mean aerodynamic chord, m (ft) drag, N (lbf) equivalent diameter (diameter of nozzle equivalent in area to total area of four model nozzles), 0.193 m (0.633 ft) height of moment center above ground plane, m (ft) tail incidence angle, positive when trailing edge is down, deg lif

22、t, N (lbf) pitching moment, m-N (ft-lbf) free-stream Mach number total pressure in ejector plenum, N/mm2 (lbf/in2) free-stream dynamic pressure, N/m2 (lbf/ft2) 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-S wing area, m2 (ft2) T thrust at static

23、 conditions, N (lbf) Ve Free - s tream dynamic pressure Nozzle-exit dynamic pressure effective -velocity ratio, v, free-stream velocity (true airspeed), knots wi ejector-inlet weight flow, N/sec (lbf/sec) a P angle of sideslip, deg AL,AD ,AMY fuselage angle of attack (wing incidence, 1.75O), deg inc

24、rements of lift, drag, and pitching moment, respectively, due to interference aileron-deflection angle (sum of right (up) and left (down) deflections), deg flap-deflection angle, deg nozzle-deflection angle, measured downward from plane containing fuselage reference line and perpendicular to plane o

25、f symmetry, deg rudder deflections, positive when trailing edge is deflected toward left, deg effective-dihedral parameter directional - stability parameter side-force parameter rolling moment per degree of aileron deflection pitching moment per degree of tail deflection yawing moment per degree of

26、rudder deflection Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MODEL AND APPARATUS A three-view drawing of the model with dimensions is given in figure 1, and photo- graphs of the model are given in figure 2. The model was a 1/6-scale representati

27、on of the Kestrel (XV-6A) vectored-thrust jet V/STOL airplane. It was constructed of alumi- num and wood covered with fiber glass. The flaps and ailerons could be fixed at various deflection angles. The empennage includes a variable-incidence horizontal tail and adjustable rudder. Landing-gear provi

28、sion was made to simulate both the landing and the cruise conditions. The model was mounted on a six-component strain-gage balance attached to a sup- port sting which enclosed the air line to the model powerplant (fig. l(b). The sting was fixed to a vertical strut which had provisions for varying th

29、e angle of attack, angle of- sideslip, and height above a movable ground plane. The movable ground plane was a fabric belt over two spanwise rollers driven by an electric motor (ref. 13). The model was powered by four cold-air ejectors (ref. 14) each exhausting through swiveling nozzles along the fu

30、selage and supplied by a compressed-air line enclosed in the support sting. The nozzles were individually adjustable to give jet deflection angles between Oo and 95O with respect to the model horizontal plane. Power variations were obtained by varying the ejector air-supply pressure. Ejector operati

31、ng variables were determined from calibrations based on the reference pressures measured in the ejector plenum chambers by electrical pressure transducers. Angles of attack were measured by an electronic inclinometer mounted in the fuselage, and sideslip angles were deter- mined from a calibrated ge

32、aring arrangement on the model support strut. The foregoing measurements together with tunnel operating variables and forces and moments were recorded on magnetic tape. TESTS AND CORRECTIONS The investigation was conducted in the 5.18-meter (17-foot) test section of the Langley 300-MPH 7- by 10-foot

33、 tunnel. The maximum free-stream dynamic pressure was 527 N/m2 (11.0 lbf/ft2) or a maximum free-stream velocity of 29.3 m/sec (96.1 ft/sec) resulting in a Reynolds number based on E of about 0.88 X 106. Data were obtained through a range of angles of attack from -4O to 24O and a range of sideslip an

34、gles generally from -12O to 12O (or -6O to 24 in some cases) at various thrust coefficients or effective free-stream-to- jet velocity ratios. Model variations investigated at fixed thrust coefficients and nozzle deflections were tail incidence, rudder deflection, flap deflection, aileron deflection,

35、 and height above the ground plane. In addi- tion to the study of the transition configuration (bf = 600, 6n = 65O, 85O, and 95O, and land- ing gear down), the normal-cruise configuration (6f = Oo, 6, = Oo, and landing gear 5 Provided by IHSNot for ResaleNo reproduction or networking permitted witho

36、ut license from IHS-,-,-retracted) was investigated with and without power to determine horizontal-tail, rudder, and aileron effectiveness in cruise. Attachment of the air-supply line to the model affects the sensitivity of the strain- gage balance. Variations in air-line pressure also produce force

37、s and moments on the balance. Both of these nonaerodynamic effects are consistently repeatable. Calibrations were made to determine these effects for which corrections were made on the balance readings. No blockage or wall corrections have been applied as they are believed to be small for a model of

38、 the present size in the 5.18-meter (17-foot) test section (ref. 15). ENGINE SIMULATION Simulator Calibration The Pegasus 5 turbofan engine of the Kestrel (XV-6A) was simulated by using four cold-air-powered ejectors (ref. 14) exhausting through swiveling nozzles. Figure 3 pre- sents the variation o

39、f inlet weight flow and thrust for each ejector as functions of the ref- erence pressure measured in the ejector plenum chambers. The inlet weight flow was determined by using a calibrated bellmouthed entrance. The thrust was the total resultant force measured statically by the strain-gage balance;

40、as a result, the thrust in this paper corresponds to the airplane gross thrust. The model thrust used to compute thrust coef- ficient during this investigation was determined from these calibration curves as a func- tion of the reference pressure. The swiveling nozzles were modeled to represent the

41、Pegasus 5 turbofan engine. In this investigation, there were some variations in individual nozzle deflections for a given nominal total-thrust deflection angle. For example, for a nominal setting of 85O, individual nozzle deflections were 89O, 82O, 88O, and 88O, or a total-thrust deflection of 860 b

42、ased on measured axial and normal forces. This investigation also used nominal nozzle deflections of Oo, 45O, 65O, and 95. For these nominal deflections, the actual deflections, based on measured axial and normal forces, were -3O, 45O, 66O, and 96O. Effective-Velocity-Ratio Simulation The relationsh

43、ip between effective velocity ratio Ve and thrust coefficient CT for this investigation is given in figure 4. The velocity ratios were obtained by varying bath the jet thrust T and the tunnel dynamic pressure q,. A series of runs were made to determine whether the force- and moment-thrust ratios for

44、 a given velocity ratio Ve were independent of the magnitude of the dynamic pressures involved. Runs were made at several tunnel dynamic pressures while varying the jet thrust. This procedure gave force- and moment-thrust ratios at some equivalent velocity ratios but at different dynamic pressures.

45、The data are presented in figure 5(a) for the Oo nozzle deflection 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-and in figure 5(b) for the 85O nozzle deflection. The curves do not always coalesce and this indicates that force and moment ratios f

46、or a given velocity ratio are not entirely inde- pendent of the dynamic pressures used in obtaining that velocity ratio. However, exami- nation of these data generally indicates that the major differences in the curves occur at conditions where both the lower values of tunnel dynamic pressure q, 191

47、.5 N/m2 (4 lbf/ft2) and the lower values of thrust T 133.4 N (30 lbf) are involved. In these cases, small errors in force measurements and in calculated thrust may noticeably affect ratios of small values. From these results, a combination of q, and T was estab- lished for variation of the effective

48、-velocity ratio through the transition-speed range. The maximum thrust was used for Ve 0.20 while q, was varied from 0 to 527 N/m2 (11 lbf/ft2). For Ve between 0.20 and 0.50, the maximum q, was used while thrust was varied from the maximum value 645 N (145 lbf) to approximately 146 N (33 lbf). - Unf

49、ortunately, it was necessary to use very low values of T at the maximum tunnel dynamic pressure (527 N/m2 (11 lbf/ft2) to obtain values of Ve greater than 0.50; as a result, the accuracy of these data may not be as good as the transition (Ve 0.50) data. 6f, deg Data description PRESENTATION OF RESULTS CT or ve Figurt deg I The data figures are presented in the following table. The out-of