NASA NACA-RM-L53B26-1953 Low-speed aileron effectiveness as determined by force tests and visual-flow observations on a 52 degrees sweptback wing with and without chord-extensions《.pdf

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1、Y -z RESEARCH MEMORANDUM LOW-SPEED AILERON EFFECTIVENESS As DETERNZINED BY FORCE TESTS AND VISCTAL-FLOW OBSERVATIONS ON A 52O SWEPTBACK WING WITH AND WITHOUT CHORD-EXTENSIONS Efy Patrick A. Cascro : 2 “ NATIONAL ADVISORY COMMITTEE P FOR AERONAUTICS WASHINGTON April 29, 1953 Provided by IHSNot for Re

2、saleNo reproduction or networking permitted without license from IHS-,-,-1x NACA RM 5326 NATIONAL ADVISORY COWTTEE FOR AERONAUTICS LOW-SPEED AILERON EFPECTIVENESS AS IEERMDED BY FORCE TESTS AND VISUAL-WW OBEERVATIONS OM A 52O SWEI*TBACK WIXG WITH AND WITHOWI CHORD-EXTENSIONS By Patrick A. Cancro A l

3、ow-speed investigation has been conducted in the 19-foot pres- sure tunnel at Reynolds numbers of 5.5 X Lo and 1.3 X 10 to determine the effect of leading-edge chord-extensions on the aileron characteris- tics of a 52O sweptback ufng. The King had an aspect ratio of 2.83, a taper ratio of 0.617, and

4、 symmetrical circular-arc airfoil sections, and was equipped with a 0.495-semispan aileron which extended from 0.415 to 0.910 semispan. In an attempt to simulate a more centrally located aileron, the outboard portion of the aileron was fixed to the 6 6 h rn wing and the resulting 0.370-semispan aile

5、ron wm tested. The results of the investlgation indicate that the values of the aileron effectiveness parameter Cz8 on the plain wing at zero lift for ailerons of 0.495 and 0.370 semispan were 0.00085 and 0.00063, respectively. However, at maximLrm lift the values of Cz8 were approx- imately 65 perc

6、ent of the values obtaine was from -25O to 25. All data have been reduced to standard nondimensiod coefficients. Stream inclination and jet-boundary corrections have been applied to the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5326 5 4

7、. angle of attack. Jet-boundary corrections hgve beeri applied to the pitching-moment and rolling-moment data (ref. 3). The rolling-mment Kfth 6a = Oo (see fig. 3) . The results for the various configuratiom are similar and indicate that the asymmetry due either to the wing or to the air stream wa n

8、ot constant throughout the angle-of-attack we. In an attempt to determine the cause of this variation, the wing was inverted and w-as tested through the angle-of-attack range. The data obtained were found to be essentially the same as those obtained Mth the wing erect. As a result, the variation of

9、Cz with a is probably due to an asymmetry of the air stream which is not constant with.angle of attack. In order to present data without the effects of tunnel air- stream asymmetry, values obtained from the faired curves shown Fn fig- ure 3 were applied as tares. - coefficient Cz varied wLth angle o

10、f attack a for each configuration Visual-flow studies on the plain wing and the wing equlpped with leading-edge chord-exteneions are shown photographically in figures 4 and 5. The lift ad *Le rolling-, pitching-, and yawing-moment chaxac- teristics obtahed for the plain wing configuration with a 0.3

11、70- and a 0.495-semispan aileron are presented in figures 6 and 7. similar extensions and in figures 10 and 11 for the wing with ex-sible leading-edge flaps. Representative cross plots of Cz against aileron from the curves of Cz plotted against 6a were used as the basis for the fairings of the curve

12、s of Cz plotted against a. Some scatter was encountered but it did not appear to affect materially the trends. In order to show the aileron effectiveness for a small range of aileron deflections through 6, = Oo, variatfons of Cz8 with angle of attack are presented in figure 14. . data are presented

13、in figures 8 and 9 for the wing equipped with chord.- - deflection 6a are presented in figures 12 and 13. The valuee obtained. DISCUSSION Visual-Flow Studies An opaque =quid was used in the rLsual-flow studies in the boundary layer over the swept King as presented in this paper. The investigation wa

14、s made during the early stages of that testing technique in the Langley 19-foot pressure tunnel, and as such the results obtained are not as complete as presently possible. These studies were made at a Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-

15、6 NACA RM 5326 Reynolds number of - 1.3 X 10 6 after tests at a Reynolds number of 3 .O X 10 6 had indicated no appreciable difference in flow patterns. I - As pointed out in reference 2, the changes in section lift charac- teristics brought about by the leading-edge-vortex flow on swept wings with

16、sharp leading edges produce changes in the pitching-moment charac- teristics throughout the angle-of-attack range. It was of interest in the present investigation to determine from visual-flow observations the location of the vortex and to define its path aa it moved over the wing. The procedw emplo

17、yed wa8 to allow a solution of lanrpblack and kerosene to flow into the boundary layer through a tube at the end of a strut-mounted probe. It was possible to move this probe spanwise and chordwise at will, as can be seen in figure 5. In this investigation, except for the condition sham in figure 5,

18、the solution was released at a chordwise position of approximately 0.05 and a spanwise position of 0.50b/2 for the configurations with and without chord-extensions. In addition, the solution was released at the inboard leading edge of the chord-extension. The results obtained are sham photographical

19、ly as figure 4. The interpretation of the flow studies is as follows: - When the leading-edge separation vortex (such as described in ref. 2) enveloped the chordwise position at which the solution was released, the solution flowed outboard and forward until it reached the position -re the vortex w o

20、f the leading-edge vortex undoubtedly results in a decrease in lift Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 - NACA RM 5326 at the tip sectioqs, and although the effect is destabilizing, the C, characteristics become more linear. Compare fig

21、s. 6 and 8 and figs. 7 and 9.) AB the angle of attack is increased, the leading-edge vortex emanating from the apex of the wtng apparently becomes stronger than the vortex emmating from the chord-extensions. The resultant flow tends to move outboard. It is interesting to note that at higher angles o

22、f attack (from a = 16O) the location of the resultant flow approaches that obtained for the plain wing configuration, (fig. 4( j ) ) . In Uke manner the pitching-moment characteristics approach those obtained with the plain wing configuration. A secondary vortex is evident along the leading edge of

23、the chord- extension (figs. the value of Czg WLn wing configuration is somewhat less than that obtained with the 0.495-semispan aileron. It is believed that the inboard portion of the aileron (0.370-semispan) is in a more stable region of flow. - Provided by IHSNot for ResaleNo reproduction or netwo

24、rking permitted without license from IHS-,-,-10 - NACA RM 1,5326 Since the vortex appears tg have no appreciable effect on the - aileron effectiveness, whether a 0.495- or a 0.370-semispan aileron is used, it is believed that the trailing-edge outflow common to swept wings is the primary influence o

25、n the aileron characteristics. I The results of an investigation of the lateral control character- istics of a 52O sweptback wing having an aspect ratio of, 2.83 and incorporating circular-arc airfoil sections, with and without chord- extensions and extensible leading-edge flaps and equipped with ei

26、ther a 0.495-semispan or 0.370-semispan aileron, are summarized as follow: The changes in pitching moment throughout the angle-of-attack range could be assocfated with changes in leading-edge flow. Although the leading-edge flgf yas markedly changed by the addition of chord- extensions, the aileron

27、effectiveness w these values were in accordance with those estimated by simple sweep theory. However, at maximum lift the dues of CI were approximately 65 percent of the dues obtained at zero lift for each configuration. With the 0.370-semispan ailerorr, the value of Czg was approximately 30 percent

28、 .lese at zero. lift than the value obtained with the 0.495-semiepan aileron. This reduction w-as attributed to geometric differences rather than aerdynamic differences. “ I Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va. .“. Provided by IHSNot for Re

29、saleNo reproduction or networking permitted without license from IHS-,-,-1. Furlong, G. Chester: Exploratory Investigation of Leading-Edge Chord-Extensions To i I ! .Of cz O :Of . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(a) a = oO. Phin wing

30、Wing witA chord-extension Wing wlth chord-extermion (b) a = 2O. Figure 4.- Visual-flow studies on a 52 amback biconvex wing with circulm-arc airfoil sections. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(a) a = 6“. -w%7 . . . . Provided by IHSNot

31、 for ResaleNo reproduction or networking permitted without license from IHS-,-,-. (f) a= loo. KZy= Figure 4.- Contiaued. L-77939 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(g) a = 120. (h) a = 14. Flgure 4.- ContFnued. L-77940 R Provided by IH

32、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. , Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(2) a =. 280. =” Figure 4.- Concluded. 577942 I . . 1 . . . . . H Provided by IHSNot for ResaleNo reproduction or

33、networking permitted without license from IHS-,-,-Figure 5.- Visual-flow st* on a 520 meptback biconvex rlng with circular-arc airfoil sections, ahowing the approldmate location of the rearward polnt of attachment of the wrtex. a = w. . Provided by IHSNot for ResaleNo reproduction or networking perm

34、itted without license from IHS-,-,-22 . . “ NACA RM L53B26 1 . “ r “ R a 2 n -10 v -20 D -15 -25 I u f b (a) Variation of and C with a. Figure 6 .- Aileron characteristics of plain wing with 0.370-semispan aileron. “ 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without lice

35、nse from IHS-,-,-NACA RM L53B26 . q-b (b) Variation of & with CL a,nd % with u. Figure 6. - Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 clllsll) NACA RM L53B26 (a) Variation of CL and Cz with a. “ , Figure 7.- Aileron characteristic

36、s of plain wing with 0.495-semispan aileron. “ . “ , . .- r- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.X . NACA RM L53B26 u -25 D -1s (b) Variation of Cm with and C, with a. Figure 7. - Concluded. - Provided by IHSNot for ResaleNo reproduction

37、 or networking permitted without license from IHS-,-,-26 /. 2 x0 .6 0 .02 .o/ - NACA RM L53B26 (a) Variation of CI, and Cz with a. Figure 8.- Aileron characteristics of wing with chord-extensions and 0.37O-semispan aileron. # c Provided by IHSNot for ResaleNo reproduction or networking permitted wit

38、hout license from IHS-,-,-MACA RM L53B26 (b) Variation of C, with % and Cn with a. Figure 8.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACA RE4 L53B26 (a) Variation of and Cz with a. Figure 9.- Aileron characteristics of wing wit

39、h chord-extensions and 0.495-semispan aileron. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MACA RM L5-26 - (b) Variation of with C, and Cn with a. Figure 9.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-