NASA-TN-D-3845-1967 Rolling stability derivatives of a variable-sweep tactical fighter model at subsonic and transonic speeds《在亚音速和跨音速下可变掠翼战斗机模型的旋转稳定性导数》.pdf

《NASA-TN-D-3845-1967 Rolling stability derivatives of a variable-sweep tactical fighter model at subsonic and transonic speeds《在亚音速和跨音速下可变掠翼战斗机模型的旋转稳定性导数》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-3845-1967 Rolling stability derivatives of a variable-sweep tactical fighter model at subsonic and transonic speeds《在亚音速和跨音速下可变掠翼战斗机模型的旋转稳定性导数》.pdf（57页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICAL z c NOTE NASA TN D- c. I -_- - 3845 - ROLLING STABILITY DERIVATIVES FIGHTER MODEL AT SUBSONIC AND TRANSONIC SPEEDS OF A VARIABLE-SWEEP TACTICAL - . . .- . ;t, .; , by William P. Henderson, W. Pelbum Pbills, and Thomas G. Gainer Langley Research Center Lungley Station, Humpton, Vu. ,I

2、, 9 ; 8. -. !;: , NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. FEBRUARY 1967 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I111111 11111 11111 l1lI1 Ill11 Ill1 Ill1 1111 111 ROLLING STABILITY DERIVATIVES OF

3、A VARIABLE-SWEEP TACTICAL FIGHTER MODEL AT SUBSONIC AND TRANSONIC SPEEDS By William P. Henderson, W. Pelham Phillips, and Thomas G. Gainer Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghousefor Fed= Scientific and Technica

4、l Information Springfield, Virginia 22151 - Price $2.50 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ROLLING STABILITY DERIVATIVES OF A VARLABIX-SWEEP TACTICAL FIGHIER MODEL AT SUBSONIC AND TRANSONIC SPEEDS By William P. Henderson, W. Pelham Phi

5、llips, and Thomas G. Gainer Langley Research Center SUMMARY An investigation was made in the Langley high-speed 7- by 10-foot tunnel to determine the rolling stability derivatives of a variable-sweep tactical fighter model. This investigation included the effects of wing sweep, angle of attack, Mach

6、 number, and the addition of tail surfaces. The study was made at Mach numbers from 0.40 to 1.20 and at angles of attack from -5 to 20. test Reynolds number per foot (per 30.48 cm) varied from 2.45 X lo6 to 4.15 X lo6. system of axes and are nondimensionalized with respect to the wing in a 16O swept

7、back position. The The derivatives presented herein are referred to the stability The results indicate that at low angles of attack the wing-fuselage com- bination exhibited large reductions in the damping-in-roll derivative C and slight decreases in the yawing moment due to rolling velocity as the

8、wing sweep was increased from 20 to 72.5O. uration with the wing swept back 20 was increased, the damping in roll was considerably reduced. However, for the configuration with the wings swept back 72.5, the damping in roll increased for angles of attack from 0 up to about 8; above this angle-of -att

9、ack range, reductions occurred. without the wings, the horizontal and vertical tails provided an increment in the damping-in-roll derivative at low angles of attack of about -0.04. the addition of the wings at either 20 or 72.5O of sweep, this increment was reduced by more than one-half. contributed

10、 a small positive increment to Cnp at zero angle of attack. With increasing angle of attack, this positive contribution decreased and became a negative contribution. Cnp As the angle of attack for the config- For the configuration With For all wing sweep angles, the tail assembly Estimates of the ro

11、lling stability derivatives for the wing-fuselage com- bination were in good agreement with experimental results in the low to moderate angle-of-attack range. The contribution of the tail assembly to the rolling stability derivatives was not accurately predicted. Provided by IHSNot for ResaleNo repr

12、oduction or networking permitted without license from IHS-,-,-INTRODUCTION An extensive research program is being conducted by the National Aeronautics and Space Administration to provide aerodynamic information for airplane configurations employing variable-sweep wings. A nmber of investiga- tions

13、have indicated that the use of variable sweep offers a means of realizing efficient subsonic and supersonic flight characteristics in one airplane con- figuration. Recently the study of variable-sweep airplane configurations has been extended to include measurements of the rolling stability derivati

14、ves Cnp, and Cy , which are important to the calculation of the lateral motion of P the airplane. Clp, Reference 1 presents measurements of the rolling stability derivatives at subsonic and transonic speeds on a variable-sweep configuration at wing-leading- edge sweep angles of 25O, 75O, and 108. co

15、mparison of experimental and estimated data, made to determine the usefulness of some known methods of estimating these derivatives. Also presented in reference 1 is a The purpose of the present investigation was to measure the rolling sta- bility derivatives Czp, Cnp, and Cyp of a variable-sweep ta

16、ctical fighter model. Estimates of these derivatives were also made by using the procedures outlined in reference 1, and these estimates are compared with the experimental results. The investigation was made in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.40 to 1.20 and at angl

17、es of attack from -5 to 20. Configurations with wing-leading-edge sweep angles of 20, 50, and 72.5 were investigated. Static longitudinal and lateral aerodynamic characteristics of a similar model at subsonic and transonic speeds are pre- sented in references 2 and 3. SYMBOLS The results of this inv

18、estigation are referred to the stability system of axes shown in figure 1. The wind-tunnel data are nondimensionalized with respect to the geometric characteristics of the wing in a 16O sweptback posi- tion. These reference dimensions, given both in the U.S. Customary Units and in the International

19、System of Units (SI), are presented in table I. For com- parison purposes, the wing area and span for the 20 and 72.5 wings are also presented. The moment reference center was located at fuselage station 23.21 inches (58.95 em) for the 20 sweptback position and at fuselage station 23.70 inches (60.2

20、0 em) for the 50 and 72.5 sweptback position, as shown in figure 2. b reference wing span, feet (meters) - C mean aerodynamic chord of 16O sweptback wing of configuration A, feet (meters) 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Rolling mome

21、nt qsb cz rolling-moment coefficient, Yawing moment qSb Cn yawing-moment coefficient, Side force ss CY side-force coefficient, ac, , per radian horizontal-tail incidence angle (positive when trailing edge is down), degrees it M free-stream Mach number P angular velocity about X stability axis, radia

22、ns/second wing-tip helix angle, radians free- s tream dynamic pressure, Pb 2v 9 - $V2, pounds /f oot2 (newtons /meter2) S v f ree-s tream velocity, feet /second (meters /second) wing reference area, feet2 (meters2) x,y,z stability axes a angle of attack, angle of the wing chord relative to the relat

23、ive wind, degrees P angle of sideslip, degrees n increment in a derivative due to tail assembly A leading-edge sweep angle of outboard wing panel, degrees P air density, slugs/foot3 (kilograms/meter3) 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,

24、-CONFIGURATION DESIGNATIONS Two configurations, designated configurations A and B, were tested. Con- figuration A had a longer fuselage nose length but a shorter wing span than configuration B. (See fig. 2.) The following letter designations are used to represent component parts of the configuration

25、s: F fuselage w20 wing with leading edge of outer wing panel swept back 20 wing with leading edge of outer wing panel swept back 50 w50 wing with leading edge of outer wing panel swept back 72.5 w72. 5 v vertical tail and ventral fins H horizontal tails MODEL AND APPARATUS A two-view drawing of the

26、configurations tested is shown in figure 2, and photographs of the model are presented as figures 3 and 4. untwisted, employed NACA 64A-series airfoil sections with 0.20 camber defined parallel to the free stream for the wing in the 16 sweptback position. thickness for the 16O sweptback wing of conf

27、iguration A varied from about 11 percent chord at the wing pivot to about 10 percent chord at the wing tip. This wing was mounted at 1 positive incidence relative to the fuselage refer- ence line. The wing for configuration B differs only from the wing for con- figuration A in that the aspect ratio

28、was increased by extending the wing tips 1.91 inches (4.85 cm) along the span. because it is the reference wing used to nondimensionalize the data, even though no data were obtained in this investigation for the configurations with the wings in this sweep position. Unless otherwise stated, the inboa

29、rd glove shown in figure 2 was used on the wing. The wing, which was The The 16 sweptback wing is described here The horizontal tails had a modified biconvex airfoil section (parallel to free stream) with a thickness of 4 percent chord at the root and 3 percent chord at the tip. These tails, when ro

30、tated about a hinge line which was swept back 13.3, were capable of deflection angles from 0 to -2OO. tail section consisted of a 4-percent-thick modified biconvex airfoil (parallel to the free stream). The vertical All tests were made with the inlets open and the inlet spike positioned to provide t

31、he proper engine airflow at a Mach number of 1.20. A sketch of the steady-state forced-roll apparatus installed in the Langley high-speed 7- by 10-foot tunnel is shown in figure 5. The model was 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-mount

32、ed on a six-component wire strain-gage balance of the type normally used for static tests of sting-supported models. Electrical signals from the strain- gage balance were transmitted to the data recording equipment by wire leads, slip rings, and brushes. interchangeable couplings between the balance

33、 and the rotating sting support. A more complete description of the mechanical operation of this apparatus is presented in reference 1. Variation of angle of attack was obtained by means of TESTS AND CORRECTIONS Damping-in-roll tests were made in the Langley high-speed 7- by 10-foot tunnel over a Ma

34、ch number range from 0.40 to 1.20 for the 72.5O wing sweep position and from 0.40 to 0.80 for the 20 wing sweep position. attack was varied from -5O to 20. per foot (per 30.48 cm) with Mach number is presented in figure 6. The angle of The variation of the test Reynolds number The support system def

35、lected under load and these deflections, combined with the effects of model product of inertia and any initial displacement of the center of mass of the model from the roll axis, introduced centrifugal forces and moments when the model was rotated. The contribution of these centrifugal forces and mo

36、ments to C2, Cn, and Cy are, to the first order, symmetrical about zero rolling velocity. The rolling derivatives Czp, Cnp, and were therefore reduced from data obtained at several rolling veloci- ties having equal magnitude but opposite sign so that the centrifugal contri- bution would be canceled.

37、 The angles of attack have also been corrected for deflection of the balance and support system under load. Cyp In an attempt to fix transition, 0.10-inch-wide (0.254-cm) strips of No. 120 carborundum grains (mean particle diameter of 0.0049 inch (0.0124 cm) were placed on the model. These strips we

38、re applied around the fuselage 1.65 inches (4.19 cm) back from the nose, around the inlets 0.40 inch (1.01 cm) from the leading edge, and at 0.25 inch (0.64 cm) rearward perpendicular to the leading edges of the wing, horizontal tails, and vertical tails. For all tests at transonic speeds (M 0.80) w

39、ith the open-slot configu- ration of the tunnel, no jet-boundary or blockage corrections are necessary and therefore were not applied to the data. However, for the tests at subsonic speeds, with the closed-slot configuration of the tunnel, jet-boundary correc- tions estimated using reference 4 were

40、applied to the angle of attack, and blockage corrections estimated using reference 5 were applied to the dynamic pressure and Mach number. PRESENTATION OF DATA The derivatives presented herein are referred to the stability system of axes and are nondimensionalized with respect to the wing in a 16O s

41、weptback 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-position. outline of the contents of the data figures is presented: For convenience in locating a particular set of data, the following Figure Variations of rolling stability derivatives with

42、 Mach number for: Configuration A with FW2oVH and glove off; it = 0 . . . . . . . Configuration A with FwzoVH; it = 0 . . . . . . . . . . . . . . Configuration A with FW20VH; .it = -10 . . . . . . . . . . . . . Configuration A with FWOVH; it = -20 . . . . . . . . . . . . . Configuration A with FW2oV

43、 . . . . . . . . . . . . . . . . . . . Configuration A with FW20 . . . . . . . . . . . . . . . . . . . Configuration A with FW20 and glove off . . . . . . . . . . . . Configuration B with FW20VH; it = 0 . . . . . . . . . . . . . Configuration A with FW VH; it = 0 . . . . . . . . . . . . . . Configur

44、ation A with FW50 . . . . . . . . . . . . . . . . . . . Configuration B with FWs0VH; it = 0 . . . . . . . . . . . . . . Configuration A with FW72.5VH; it = Oo . . . . . . . . . . . . . Configuration A with FW72.5VH; it = -10 . . . . . . . . . . . Configuration A with FW72.5VH; it = -20 . . . . . . .

45、 . . . . . Configuration A with FW72.5V . . . . . . . . . . . . . . . . . . Configuration A with FW72.5 . . . . . . . . . . . . . . . . . . Configuration B with FW723VR; it = 0 . . . . . . . . . . . . . Configuration A with FVH; it = Oo . . . . . . . . . . . . . . . Configuration A with F alone . .

46、. . . . . . . . . . . . . . . . Contribution of vertical and horizontal tails to C - a = 1.0 Comparison of experimental and estimated Contribution of vertical and horizontal tails to 50 Contribution of horizontal tails to czp; a = l.oo . . . . . . . IP for wing-fuselage combinations . . . . . . . .

47、. . . . . . . . . . . . . . . . cnP and Cy cnP P 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 RESULTS AND DISCUSSION Experimental Results The basic data of this investigation were obtained as the variation of forces and moments with wing-tip helix angle and angle of attack. The

48、derivatives Czp, from these data at values of Results with Ciifferent data symbols were obtained at different values of angular velocity. In general, rolling velocity is seen -to have only a slight effect on the rotary stability derivatives in the low angle-of -attack range. example, see fig. 8.) pb

49、/2V at each test Mach number and Cy were then extracted c“p P pb/2V of equal magnitude but opposite sign. (For However, at the higher angles of attack, where wing stall 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-is encountered, rolling velocity is seen to h