NASA NACA-RM-L51F22-1951 Investigation of low-speed lateral control and hinge-moment characteristics of a 20-percent-chord plain aileron on a 47 7 degrees sweptback wing of aspect p 6).pdf

1、RESEARCH MEMORANDUM INVESTIGATION OF LUW-SPEED LATERAL CONTROL AND HINGE-MOMENT CHARACTERXSTICS OF A 20-PERCENT-CHORD PLAIN AILERON OM A 47.7O SWEPTBACK WING OF ASPECT RATIO 5.1 AT A REYNOLDS NUMBER OF 6.0 x 106 By William M. Hadaway and Rein0 J, Salmi Langley Aeronautical Laboratory NATIONAL ADVISO

2、RY COMMITTEE FOR AERONAUTICS WASHINGTON October 22, 1951 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 r INVESTIGATION OF LOW-SPm LATa CONTROL AND HINGE-“ CHARACTERISTICS OF A 2O-PERCENT-CHOFCO PLCUN By William M. Ifadaway and Rein0 J. Sam SUMMAR

3、Y The low-speed lateral control and hin -moment characteristics of a 20-percent-chord plain aileron on a 47. 79“ sweptback wing of aspect ratio 5.1 have been determined in the Langley lg-foot PI-SSU tunnel. The tests were made with and without high-lift and stall-control devlces . at a Reynolds numb

4、er of 6.0 X LO . 6 The results indicated that an airplane with a wing similar to the one tested may exhibft undesirable rolling oscillations and vibrations at moderate and high anglee of attack due to intermittent separation of flow over the wing. The static rolling moments obtained with large ailer

5、on deflections were greater in magnitude, however, than the rolling moments induced by the separated flow, thereby indicating that some degree of lateral control could be maintained. At zero angle of attack, a rate of change of rolling-moment coefficient with aileron deflec- tion C of O.OO080 was ob

6、tained which was in fair agreement with the calculated value. The addition of leading- andtrailing-edge flaps did not appreciably affect C at law lift coefficients. Because of the nonlinear characteristics the rolling-moment data, the value of the aileron effectiveness parameter Cz8 was not well-def

7、ined in the angle- of-attack range through which flow separation occurred. Emever, the data indicated that the rolling moments near maximum lift were about 70 percent of the values obtained at zero angle of attack for. large total aileron deflections. Measurements of the aileron hinge moments and ba

8、lance-chanhr pressures indicated that a ratio of the aileron nose balance to the aileron chord of 0.60 or more will be required to balance ccmpletely the internally sealed type of aileron. 28 Z8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA

9、FtM L5lF22 Previous investigations of plain flap-type ailerons on sweptback wings (references 1 to 3) have shown that the aileron effec1;iveness in the low lift range can be predicted by semiempirical means based on simple sweep theory. The effectiveness of the aileron at the higher lift coefficient

10、s, however, cannot be calculated because of. the early separation of flow over the wing (reference k). An investigation waa made on a 47.70 meptback wing of aspect ratio 5.1 employing a 20-percent- chord outboard aileron to provide infonaation on the aileron effectfveness on a wing of relatively hig

11、her aspect ratio and sweep than has previously been investigated. The tests were husde at a Reynolda number of 6.0 X 10 6 and a Mach number of 0.14. These testa are part of a general investigation of the subject wing and the longitudinal stability characteristics have been reported in references 4 t

12、o 6. SYMBOLS All data are referred to a system of wind axes originating at the quarter-.chord point of the mean aerodynamic chord projected to the plane of symmetry. Symbols used herein are defined a6 f olltrws : CL c, ct lift coefficient (F) pitching-moment coefficient (Pitchi;Emment -) yawing-mome

13、nt coefficient rolling-moment coefficient (Rollintrment aileron hinge-moment coefficient aileron-load coefficient ( At moderate and high qles of attack, how- ever, the effects of the leeage acrom tKe seal on “the hinge monents and rolling moments for the wing of the present investigation are believe

14、d to be smsll. The aileron-load coefficient was measured normal to the wing chod for all deflection angles. An andysis of the aileron pressure distri- bution data of reference 8, however, indicated that the chorltwise forces on the aileron are small; therefore the force normal to the aileron chord l

15、ine can be 8pp320Ximated by dividing Cz, by the cosine of 8,. RESUiTS AMD DISCUSSION The basic aileron data are presented in figures 4 to 8. The results have been summarized in figures 9 to 12. Investigations of lateral control characteristics at low speeds and high Reynolds nmbers are of primary co

16、ncern at high lift coefficients. The rolling-moment data in the high-lift-coefficient range obtained in the present investigation, however, exhibited scatter snd nonlinearity due to unsteady and unsymmetrical forces on the wing which resulted from intermittent separated flow. Considerable vibration

17、of the model also occurred in the separated flow range, especially at angles of attack near maximum lift. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA m L51F22 7 1 In an effort to establish the reliability of the data, some retests were made

18、whereas in the present investigation, separation occurred at an angle of attack well below Ch, as indicated by figures 4, 5, and 6. The results obtained frm tests on the wing of the present investigation do not necessarily mean that the same type of intermittent separated flow will be obtained on a

19、sfmilar wing having a different airfoil section. Also, the intermittent type of flow separation might be improved by the use of stall-control devices other than those tested. Rollinn-moment characteristics. - Rolling-mmnt data are presented for the entire angle-of-attack range tested; but the ailero

20、n effective- ESS parameter CZs, as detedned from a small range of aileron deflec- tions through 6, = Oo, WS not evaluated for angles of attack greater than 160 because of the scatter of data due to intemittent separated 0.00080 was obtained at Oo angle of attack for the plain wing (fig. 9). - flow i

21、n the higher angle-Of;8tt8Ck range. A Czg value of about - A corresponding Czg value of O.m72 was calculated by the method of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 BACA. RM L5lF22 reference 9. The aileron effectiveness decreased gradually

22、 8,s the angle of attack was increased from 0 to 16O. With leading-edge flaps deflected, a value of C of 0.00075 was obtained at zero mgle of attack and remained fairly constant to about U0, beyond which it decreased. With both leading- and trailing-edge flaps deflected, CIS was about 0.00080 at 0 a

23、ngle of attack and remained approximately constant to an angle of attack of go and then decreased as tke angle of attack was increased to 16O. 28 At the angles of attack at which separation first occurred, as indicated by the unstable changes in the pitching mament (figs. 4(c), 5(c), and 6(c), a lar

24、ge decreaae in the rolling moment occurred for large negative (up) aileron deflection angles (figs. 4(a), ?(a), and 6(a). The rolling moment for large positive (dam) aileron deflection6 decreased gradually throughout the angle-of-attack range for the plain wing, decreased only in the high-lift-coeff

25、icient range with leading- edge flaps deflected, and remained approximately constant to maximum lift for the combination of leading- and trailing-edge flaps. Comparisons of the rolling moments for various total aileron deflections are presented in figure 10. For total aileron deflections of L2O or m

26、ore, the largest rolling moments were obtained bel-ow 5O angle of attack for the plain wing configuration. Between 5“ and b6O angle of attack, the largest rolling mosnents were generally obtained with leading-edge flaps deflected, and beyond 16O the largest rolling mcments were generally obtained wi

27、th both leading- and trailing-edge flaps deflected. The rolling moments near the maximum lift coefficient for large total aileron deflection angles were about 70 percent of the values obtained at zero lift coefficient. The yawing moments at small total aileron deflections were negli- gible through t

28、he angle-of-attack range for all conf%gumti.ms. As 8atotal was increased, however, adverse yawing moments were obtained for all except very low angles of attack for each configuration. The variations of the yawing-moment coefficient with angle of attack for the highest total aileron deflection teste

29、d (40O) are shown in fig- ure 10. The adverse yaw was greatest with the plain-wing configuration except at maximum lift coefficients. Pitching-moment characteristics.- The effect of maxinnmup and dm aileron deflections on the pitching-moment characteristics are presented in figures 4 to 6. The incre

30、ments in pitching moment were larger for the up-aileron that for the dm-aileron for all wing con- Pi rations except at angles below bo for the plain wing and between 11 g;u and 18 for the leading- and trailing-edge flip configuration, where the increments appeared to be equal. The dashed lines on th

31、e pitching-moment curves indicate the change in pitching moment due to I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.2 + NACA RM L5lJ?22 9 the combined maximum up and down aileron deflections for each wing con- figuration. Deflection of the lead

32、ing-edge flaps reduced the total pitching-moment increment between the maximum up and dam aileron deflec- tions to about 70 percent of the corresponding plain-wing increment at zero angle of attack. As the angle of attack was increased, the differ- ence in total increments became Bmaller until they

33、were about equal at 20. With the leading-edge flaps deflected, deflecting the trailing- edge flaps increased the pitching-moment increment between the maximum up and dm aileron deflections throughout the angle-of-attack range. Hinge-moment characteristics. - The parameters ch8, c consequently, the v

34、alues of C were more nearly constant with increase in &v Figure 5.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 mAca RM L5D22 1.4 . .4 .P Q 72 -.4 0 Figure 5.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking pe

35、rmitted without license from IHS-,-,-KACA RM L51F22 .6 .4 -2 -. 4 0 -10 a -15 v -20 -02 Figure 6.- Aileron characteristics of the wing with leading- and trailing- edge flape deflected. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 12 .8 -. 4 . I

36、P .u8 .06 0 - /a ma. RM L51F22 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-L6 I. 2 x 0 .4 -2 0 .08 .04 . % a 20 (b6 00 v -20 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-26 mAeA RM L5IF22 702 .

37、o/ -.a? Figure 7.- The variation of rolling-moment coefficient with aileron deflection for various model configurations. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure 8.- The variation of aileron hinge-moment coefficient with aileron deflect

38、ion for various model conflgumtions. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 -. 06 .04 .02 0 Figure 9.- The effects of high-lift and stall-control devices on the aileron hinge-moment and effectiveneas parametere Ch , (:28, 6 pRa c&, 83-d p

39、qy c I I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-XACA RM L5B22 .o/ 0 -. 01 702 tiiiiiiiiiiiir-1 .04 .03 .02 hOO .03 c* .02 30 . u/ G2 .QZ .u/ 0 * Figure 10.- Variation of rolling-moment and yawlng-moment coefficfents with angle of attack for several model configurations and total aileron deflectim. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-