NASA-TM-X-2555-1972 Supersonic aerodynamic damping and oscillatory stability in pitch and yaw for a model of a variable-sweep fighter airplane with twin vertical tails《带有两个垂直尾翼的可变掠.pdf

《NASA-TM-X-2555-1972 Supersonic aerodynamic damping and oscillatory stability in pitch and yaw for a model of a variable-sweep fighter airplane with twin vertical tails《带有两个垂直尾翼的可变掠.pdf》由会员分享，可在线阅读，更多相关《NASA-TM-X-2555-1972 Supersonic aerodynamic damping and oscillatory stability in pitch and yaw for a model of a variable-sweep fighter airplane with twin vertical tails《带有两个垂直尾翼的可变掠.pdf（42页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICALM EMORANDUMNASA TM X-2555CSUPERSONIC AERODYNAMIC DAMPING ANDOSCILLATORY STABILITY IN PITCH AND YAWFOR A MODEL OF A VARIABLE-SWEEP FIGHTERAIRPLANE WITH TWIN VERTICAL TAILSby Robert A. Kilgore and Jerry B. Ad cockLangley Research CenterHampton, Va. 23365NATIONAL AERONAUTICS AND SPACE ADM

2、INISTRATION WASHINGTON, 0. C. MAY 1972Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. Report No.NASA TM X-25552. Government Accession No.4. Title and SubtitleSUPERSONIC AERODYNAMIC DAMPING AND OSCILLATORYSTABILITY IN PITCH AND YAW FOR A MODEL OF A

3、VARIABLE -SWEEP FIGHTER AIRPLANE WITH TWINVERTICAL TAILS7. Author(s)Robert A. Kilgore and Jerry B. Adcock9. Performing Organization Name and AddressNASA Langley Research CenterHampton, Va. 2336512. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D.C. 205463

4、. Recipients Catalog No.5. Report DateMay 19726. Performing Organization Code8. Performing Organization Report No.L-802510. Work Unit No.136-63-02-2811. Contract or Grant No.13. Type of Report and Period CoveredTechnical Memorandum14. Sponsoring Agency Code15. Supplementary Notes16. AbstractWind-tun

5、nel measurements of the aerodynamic damping and oscillatory stability inpitch and yaw for a 1/22-scale model of a proposed carrier-based variable-sweep fighterairplane have been made by using a small-amplitude forced-oscillation technique. Testswere made with a wing leading-edge sweep angle of 68 at

6、 angles of attack from about -1.5to 15.5 at a Mach number of 1.60 and at angles of attack from about -3 to 21 at Machnumbers of 2.02 and 2.36.The results of the investigation indicate that the basic configuration has positive damp-ing and positive oscillatory stability in pitch for all test conditio

7、ns. In yaw, the damping isgenerally positive except near an angle of attack of 0 at a Mach number of 1.60. The oscil-latory stability in yaw is positive except at angles of attack above 16 at Mach numbers of2.02 and 2.36. The addition of external stores generally causes increases in both pitch andya

8、w damping. The oscillatory stability in pitch is reduced throughout the angle-of-attackrange by the addition of the external stores. The effect of adding stores on the oscillatorystability in yaw is a function of angle of attack and Mach number. The effect of changinghorizontal-tail incidence on the

9、 pitch parameters is also very dependent on angle of attackand Mach number.17. Key Words (Suggested by Author(s)Dynamic stabilityVariable sweepSupersonic aerodynamic damping19. Security dassif. (of this report)Unclassified18. Distribution StatementUnclassified - Unlimited20. Security Classif. (of th

10、is page) 21. No. of Pages 22. Price*Unclassified 39 $3.00For sale by the National Technical Information Service, Springfield, Virginia 22151Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUPERSONIC AERODYNAMIC DAMPING AND OSCILLATORY STABILITYEST PI

11、TCH AND YAW FOR A MODEL OF A VARIABLE-SWEEPFIGHTER AIRPLANE WITH TWIN VERTICAL TAILSBy Robert A. Kilgore and Jerry B. AdcockLangley Research CenterSUMMARYWind-tunnel measurements of the aerodynamic damping and oscillatory stability inpitch and yaw for a 1/22-scale model of a proposed carrier-based v

12、ariable-sweep fighterairplane have been made by using a small-amplitude forced-oscillation technique. Testswere made with a wing leading-edge sweep angle of 68 at angles of attack from about-1.5 to 15.5 at a Mach number of 1.60 and at angles of attack from about -3 to 21 atMach numbers of 2.02 and 2

13、.36.The results of the investigation indicate that the basic configuration has positivedamping and positive oscillatory stability in pitch for all test conditions. In yaw, thedamping is generally positive except near an angle of attack of 0 at a Mach number of1.60. The oscillatory stability in yaw i

14、s positive except at angles of attack above 16at Mach numbers of 2.02 and 2.36. The addition of external stores generally causesincreases in both pitch and yaw damping. The oscillatory stability in pitch is reducedthroughout the angle-of-attack range by the addition of the external stores. The effec

15、tof adding stores on the oscillatory stability in yaw is a function of angle of attack andMach number. The effect of changing horizontal-tail incidence on the pitch parametersis also very dependent on angle of attack and Mach number.INTRODUCTIONStudies are being made by the National Aeronautics and

16、Space Administration todetermine the aerodynamic characteristics of a proposed carrier-based variable-sweepfighter airplane. As a part of these studies, wind-tunnel measurements of the aerody-namic damping and oscillatory stability characteristics are being made at subsonic, tran-sonic, and superson

17、ic speeds.This paper presents the damping and oscillatory stability results in pitch and yawfor a 1/22-scale model of the proposed airplane obtained at supersonic speeds in theLangley Unitary Plan wind tunnel. The tests were made at Mach numbers from 1.60 toProvided by IHSNot for ResaleNo reproducti

18、on or networking permitted without license from IHS-,-,-2.36 at angles of attack from approximately -3 to 21 by using a small-amplitude forced-oscillation technique.COEFFICIENTS AND SYMBOLSMeasurements were made and are presented herein in the International System ofUnits (SI). Details concerning th

19、e use of SI, together with physical constants and conver-sion factors, are given in reference 1.The aerodynamic parameters, which are referred to the body system of axes, areshown in figure 1. These axes originate at the oscillation center of the model as shownin the drawings which are presented in

20、figure 2(a). The longitudinal location of the oscil-lation center is coincident with the 16.5 percent mean geometric chord station with thewing in the 20 sweep position. The reference dimensions are based on the geometriccharacteristics of the wing of the model in the 20 sweep position.b span, 0.888

21、5 meterc mean geometric chord, 0.1358 meterf frequency of oscillation, hertzit tail incidence, degreesk reduced-frequency parameter (, in pitch; , in yaw, radiansM free-stream Mach numberq angular velocity about Y-axis, radians/secondq free-stream dynamic pressure, newtons/meter2R Reynolds number ba

22、sed on cr angular velocity about Z-axis, radians/secondS area, 0.1085 meter2V free-stream velocity, meters/secondProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a angle of attack, degrees or radians, or mean angle of attack, degreesj3 angle of sidesl

23、ip, radiansu angular velocity, 2id , radians /second ., , . , ,.,. . . Pitching momentCm pitching-moment coefficient, - s_ -9CmCm = . _t , per radian per radian2VCm + Cm . damping- in-pitch parameter, per radianO MlCm = . , per radiana 9a “= per radianCm - kCm. oscillatory longitudinal stability par

24、ameter, per radian_ . , . . , Yawing momentCn yawmg-moment coefficient, -oo9Cn, per radian9CnCnr = -7-7T Per radianCn - Cn. cos a damping-in-yaw parameter, per radianProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-9CnCn = , per radian Facnn = “ per r

25、adlanC cos a + kCn oscillatory directional stability parameter, per radiann0 “r. A dot over a quantity indicates a first derivative with respect to time. The expres-sion “cos a“ appears in the lateral parameters since these parameters are referred toithe body-system axes.MODEL AND APPARATUSModelThe

26、model which was tested is a 1/22-scale version of the proposed variable -sweepairplane and is geometrically similar to the proposed airplane except for aft fuselagemodifications necessary for sting clearance. Drawings of the model showing importantdesign dimensions are presented in figure 2 and phot

27、ographs of the model are presentedin figure 3. Additional information on the geometric characteristics of the wings, verti-cal tails, and horizontal tails is presented in tables I, n, and IE, respectively.The model was tested with the wings in the 68 sweep position and with the vanesextended from th

28、e fixed inboard wing panels as shown in figure 2(a). Details of the wingin the reference 20 sweep position are given in figure 2(b) and in table I. The flow-through ducts shown in the photographs of figures 3(d) and 3(e) have two-dimensionalinlets and are canted up 0 44 and out 0 51 at the exit plan

29、e.The twin vertical tails are mounted on the duct center lines with 5 outboard cant.Details of the vertical tails are given in figure 2(c) and in table n. Details of the hori-zontal tails, which can be set at 0, -10, or -20 incidence, are given in figure 2(d) andin table m. Details of the twin ventr

30、al fins are given in figure 2(e) .Store configurations consisted of two nacelle-mounted fuel tanks and two missilesmounted on the fixed inboard wing panels. Details of store geometry and location areshown in figures 3(c) and 3(d).Three-dimensional roughness in the form of sparsely distributed No. 60

31、 carborun-dum grains was applied to the wings, vertical and horizontal tails, and ventral fins inbands 0.16 cm wide located approximately 1 cm from the leading edges in the stream-wise direction. Similar bands were applied to the engine inlets and to the vanes extendingProvided by IHSNot for ResaleN

32、o reproduction or networking permitted without license from IHS-,-,-from the fixed inboard wing panels. Individual particles of No. 45 carborundum grainsspaced approximately 0.16 cm apart were applied to the fuselage in a ring approximately3 cm from the nose. The size and location of the roughness w

33、ere computed by using themethod of reference 2 to insure a turbulent boundary layer aft of the applied roughness.Oscillation-Balance MechanismA view of the forward section of the oscillation-balance mechanism which was usedfor these tests is presented in figure 4. Since the oscillation amplitude is

34、small, therotary motion of a variable-speed electric motor is used to provide essentially sinusoi-dal motion to the balance through the crank and crosshead mechanism. Amplitudes fromnear zero to about 2 can be obtained by using cranks having different throws. An ampli-tude of slightly less than 1 wa

35、s used for this investigation. The oscillatory motion isabout the pivot axis shown in figure 4 which was located at the model station identified as“oscillation center and center of moments“ in figures 1 and 2(a).The strain-gage bridge which measures the torque required to oscillate the modelis locat

36、ed between the model attachment surface and the pivot axis. This torque-bridgelocation eliminates the effects of pivot friction and the necessity to correct the data forthe changing pivot friction associated with changing aerodynamic loads. Although thetorque bridge is physically forward of the pivo

37、t axis, the electrical center of the bridgeis located at the pivot axis so that all torques are measured with respect to the pivot axis.A mechanical spring, which is an integral part of the fixed balance support, is con-nected to the oscillation balance at the point of model attachment by means of a

38、 flexureplate. After assembly of the oscillation balance and fixed-balance support, the mechani-cal spring and flexure plate were electron-beam welded in place in order to minimizemechanical friction. A strain-gage bridge, fastened to the mechanical spring, providesa signal proportional to the model

39、 angular displacement with respect to the sting.Although the forced-oscillation balance may be oscillated through a frequency rangefrom near zero to about 30 Hz, as noted in reference 3, the most accurate measurementsof the damping coefficient are obtained at the frequency of velocity resonance. For

40、 thesetests, the frequency of oscillation varied from 2.67 Hz to 8.50 Hz.Wind TunnelThe data presented herein were obtained in test section 1 of the Langley UnitaryPlan wind tunnel. This single-return tunnel has a test section about 1.2 meters squareand about 2.1 meters long. An asymmetric sliding b

41、lock is used to vary the area ratio inorder to change the Mach number from about 1.47 to 2.87. The angle-of-attack mechanismwhich was used for these tests had a total range of about 25 when used in conjunction withProvided by IHSNot for ResaleNo reproduction or networking permitted without license f

42、rom IHS-,-,-the oscillation-balance mechanism. A more complete description of the Langley UnitaryPlan wind tunnel is given in reference 4.MEASUREMENTS AND REDUCTION OF DATAFor the pitching tests, measurements were made of the amplitude of the torquerequired to oscillate the model in pitch Ty, the am

43、plitude of the angular displacementin pitch 0, the phase angle 77 between Ty and $, and the angular velocity of theforced oscillation u. Reference 5 gives the method of making these measurementsand the procedure to calculate the oscillatory pitch parameters, Cm + Cm . andn HO!Cm - kCm., from these m

44、easurements.For the yawing tests, measurements were made of the amplitude of the torquerequired to oscillate the model in yaw Tz, the amplitude of the angular displacementin yaw of the model i/, the phase angle A between Tz and i/, and the angular veloc-ity of the forced oscillation w. Again referen

45、ce 5 shows the procedure to calculate theoscillatory yaw parameters, Cnr - Cn- cos a and Cno cos a + k2Cn-, from theseP P rmeasurements.DATA CORRECTIONS AND PRECISIONEffects due to aft fuselage modifications necessary for sting clearance and effectsdue to model-support interference are assumed to be

46、 small and no corrections for theseeffects have been made to the data. The values of a. (mean angle of attack for the testsin pitch and angle of attack for the tests in yaw) have been corrected for flow angularityin the test section as follows:Mach number,M1.602.022.36Flow angularitycorrection,deg0.

47、451.251.35These corrections apply for a model at a given tunnel station at the vertical center of thetest section; however, since the model was never far from the tunnel center line becauseof the small angle-of-attack range, the corrections were applied to all the measured val-ues of a as a first-or

48、der correction.For the data presented herein, values of the probable error of the various quanti-ties are as follows:Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Mach number, MAngle of attack or mean angle of attack, a , degReynolds number, RDampi

49、ng- in-pitch parameter, Cm + Cm, per radianOscillatory longitudinal stability parameter, Cm -k2Cm.,per radian Reduced-frequency parameter in pitch, k, radians .Damping- in-y aw parameter, Cnr - Cn. cos a, per radian . . .Oscillatory directional stability parameter, Cn cos a + k2Cn. ,P rper radian Reduced-