NASA NACA-TR-1146-1953 Aerodynamic forces and loadings on symmetrical circular-arc airfoils with plain leading-edge and plain trailing-edge flaps《带有普通前缘和普通后缘襟翼的对称圆弧机翼上空气动力和荷载》.pdf

《NASA NACA-TR-1146-1953 Aerodynamic forces and loadings on symmetrical circular-arc airfoils with plain leading-edge and plain trailing-edge flaps《带有普通前缘和普通后缘襟翼的对称圆弧机翼上空气动力和荷载》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TR-1146-1953 Aerodynamic forces and loadings on symmetrical circular-arc airfoils with plain leading-edge and plain trailing-edge flaps《带有普通前缘和普通后缘襟翼的对称圆弧机翼上空气动力和荷载》.pdf（41页珍藏版）》请在麦多课文档分享上搜索。

1、- 4 ,. i j , T ,. I. i ; ,b i 4 .- , 1 -, -_ -,I, . il , _- i L -I .“ ,j . -. - ./ ., 1 : , 8 -, : L ; . . 1. ; ! .A, ., .- , , .,. . . . ! *- . - I :, 8, ,/ ./ _I 3 _ :* 3 - 5 A7 S. 1 p A 1 SF t l.OOc -4 (a) 6-percent-thick airfoil. (b) lo-percent-thick airfoil. FIGURE I.-Symmetrical circular-arc a

2、irfoils with plain leading-edge flaps and plain trailing-edge flaps. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. ._. -.-.- _-. . - ._ .-_-. I . . +- ; .-1 -: 9 .- z-. 1, !, -( .i.: .I ,i : _I -. il . CIRCULAR-ARC AIRFOILS WITH LEADING-EDGE AND

3、TRAILING-EDGE FLAPS 3 .172c -FlOp skirt Chard line Airfoil center section,” Plain leading-edge flap - t-r- 24 Airfoil center section Plain trailing-edge flap Airfoil center section FIGURE 2.-Locat.ion of pressure orifices on B-percent-thick airfoil with a 0.15-chord plain leading-edge flap and a 0.2

4、0-chord plain trailing-edge flap. Plain Leading-Edge Flap Yi:; XlC - s 0 ; z 3 5 5 75 76 10 : :z 10 16. 1 13 :; 2:6 :z ;5 11 4 :56 15 - dc 0 . fA . :t 84 1: 08 1. 28 1. 47 1. 17 -. 15 -. 30 -. 57 -. 83 -1.21 -. 11 Plain Trailing-Edge Flap 0. 25 1. 54 1. 87 1. 53 1. 08 -1.53 -1. 08 -. 57 -. 29 -. 15

5、Airfoil Center Section Orifice the figures in which the data are presented. The airfoil 1 RESULTS AND DISCUSSION lift, drag, and pitching moment were measured and corrected 1 AIRFOILS WITH FLAPS NEUTRAL to free-air conditions by the methods described in reference 1. The flap section normal-force, ch

6、ord-force, and hinge- moment coeficients were obtained from mechanical integra- tion of the pressure distributions. Lift measurements of the models with the flaps neutral, with and without model end plates, indicated that the model end plates had no significant effect on the measured characteristics

7、. The section aerodynamic characteristics of the 6- and lo- percent-thick symmetrical circular-arc airfoils with the flaps neutral are presented in figure 4. X/C - The maximum section lift coefficients are 0.73 and 0.67 for the 6- and lo-percent-thick airfoils, respectively. This decrease in maximum

8、 section lift coefficient with increasing airfoil thickness is opposite to the trends that are shown by Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 REPORT 1146-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS the data for NACA 6-series airfoils (ref

9、. 1) through the same thickness range, but it is believed to be explainable in the following manner: As the thickness of the NACA 6-series airfoils is increased from 6 to 10 percent, the corresponding increase in the airfoil leading-edge radius results in improved air-flow conditions around the lead

10、ing edge at the high angles of attack. The increase in trailing-edge angle that results from increasing thickness tends to decrease the maximum section lift coefficient due to an increase in boundary-layer thickness on the upper surface. The favorable effect of a large leading-edge radius appears to

11、 predominate in this thickness range for the KACA 6-series airfoils and higher values of maximum lift are produced. For the circular-arc airfoils, however, the leading edges of both the 6- and lo- percent-thick airfoils are sharp and the air-flow conditions around the leading edges at high angles of

12、 attack are about the same. The effect of an increase in trailing-edge angle with increasing thickness therefore is a decrease of maximum lift. The lift-curve slopes are 0.097 and 0.090 for the 6- and lo-percent-thick airfoils, respectively. Because the air-flow (a) With model end plates. FIGURE 3.-

13、Front of a symmetrical circular-arc airfoil with and without model end plates in the Langley two-dimensional low-turbulence pressure tunnel. (b) Without model end plates. FIGI-RE 3.-Concluded. conditions around the leading edge of both circular-arc air- foils are probably very nearly alike through t

14、he complete range of angle of attack, the thicker boundary layer of the lo-percent-thick airfoil is probably the cause of the decrease in the lift-curve slope. The slope of the lift curve for the lo-percent-thick airfoil was measured at small positive or negative values of the lift coefficient to av

15、oid including the slight jog in the lift curve that occurs near zero lift. This jog in the lift curve has been noticed before in connection with sharp leading-edge airfoils (ref. 2) and appeared when the trailing-edge angle became large. Although a similar phenomenon may have existed on the 6-percen

16、t-thick air- foil, it was not of sufficient magnitude to be noticeable in the lift curve. The data (fig. 4) show no appreciable scale effect on the lift characteristics of either circular-arc airfoil with the flaps neutral through the range of Reynolds number investigated. The variation of the quart

17、er-chord pitching-moment coefficient of both the 6- and lo-percent-thick circular-arc airfoils indicates a forward position of the aerodynamic center with respect to the quarter-chord point of the airfoil. This variation of the pitching moment probably results from the relative thickening of the bou

18、ndary layer near the trailing edge on the upper surface with increasing angle of attack. The aerodynamic center of the lo-percent-thick airfoil is more forward than that of the 6-percent-thick air- foil. This shift in aerodynamic-center position is in fair quantitative agreement with data presented

19、in reference 3 which show that increases in trailing-edge angle or in the thickness of the rear portion of an airfoil cause the aerodynamic-center position to move forward. As is usually true when an airfoil stalls, the center of pressure of the circular-arc airfoils moves toward the rear and the qu

20、arter- chord moment coefficient increases negatively in the normal manner. The small negative pitching moment of both models at zero lift is attributed to asymmetrical loading resulting from very small model irregularities. For airfoils having sharp leading eclges, the drag coeffi- cient increases f

21、airly rapidly as the angle of attack departs from zero. In general, the drag coefficients clecrease with increasing Reynolds number in approximately the manner expected for ful1.y developed turbulent flow on both surfaces. In the case of the B-percent-thick airfoil, however, laminar flow apparently

22、was obtainecl over a fairly extensive portion of the upper surface at zero ancl negative angles of attack at Reynolds numbers of 3X lo6 and 6X106, as indicated by the lower drag for these conditions as compared with the drag obtained at a Reynolds number of 9 X 10”. AIRFOILS WITH FLAPS DEFLECTED IND

23、IVIDUALLY The lift and pitching-moment characteristics of the two symmetrical circular-arc airfoils with the plain trailing-edge flaps and plain leading-edge flaps deflected individually are presented in figures 5 and 6, respectively. The maximum section lift coefficients of the 6- and lo- percent-t

24、hick airfoils increased and the angles of attack for maximum lift decreased as the 0.20-chord trailing-edge flaps were deflected. The values of the maximum lift coefficients (fig. 5) for both airfoils were substantially equivalent at corresponding flap deflections. Provided by IHSNot for ResaleNo re

25、production or networking permitted without license from IHS-,-,-CIRCULAR-ARC AIRFOILS WITH LEADING-EDGE AND TRAILINGEDGE FLAPS 5 2.8 I .032 .028 3 0 ii i /II 2.4 ttH i i i .008 .004 2 0 ,F 5 I R 0. position s I X/C Y/C -.2 A* E 0 6 X106 0.222 0 E I I % -.3- ,032 I I .028 r,=- i 2.8 24 2.0 1.6 , I.2

26、.8 I I I I 2 ,024 ; s .G E ,020 8 5 ,016 s Z 2 ,012 ,008 -.8 -1.6 -24 -16 -8 0 8 I6 24 Section angle of ottock, mot deg I I 1 I I I I I I 2 c -.I .E 2 5 8 -.2 z E 2 -.3 - c .o, g -.I F I R 0.c. position s I I I I I I I I X/Y Y/C -t+l E -.2 q 6X IO6 0.214 0 E I I I 1 I I % -56 -1.2 -.8 -.4 0 .4 .8 1.

27、2 1.6 2.0 -.4 Section lift coefficient, Cl (a) 6-percent-thick airfoil. (b) lo-percent-thick airfoil. FIGURE 4.-Aerodynamic characteristics of two symmetrical circular-arc airfoils with flaps neutral Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6

28、REPORT 1146-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .8 -yI 6 -8 0 8 16 0 3x106 0 6 0 9 A I 4- v I 8 0 0 0 0 Section angle of attack, aO, deg (c) lo-percent-thick airfoil. FIGURE 4.-Concluded. Deflecting the 0.15-chord leading-edge flaps increased the maximum section lift coefficients and increas

29、ed the angles of attack for maximum lift (fig. 6) primarily by alleviating the negative pressure peaks that cause leading-edge sepa- ration near maximum lift. These pressure peaks are alleviated because the flow approaching the leading edge is more nearly alined at high angles of attack when the lea

30、ding- edge flap is deflected. The maximum section lift coefficients for the 6- and lo-percent-thick airfoils at the optimum deflection of the leading-edge flap, 30, are 1.17 and 1.15, respectively. The optimum flap deflection is defined as the flap deflection for highest maximum lift. At correspondi

31、ng deflections of the 0.15-chord leading-edge flap, the maximum section lift coefficients of both airfoils are essentially the same. At angles of attack well below those for maximum . lift, the leading-edge flaps act as spoilers on the lower surface of the airfoils and cause some reduction in lift.

32、These losses in lift increase as the flap deflection is increased. The variation of the increment in maximum section lift coefficient Ac, and increment in angle of attack at maxi- mum lift Acx, for both models with deflection of the 77UZZ leading-edge or trailing-edge flaps individually is summarize

33、d in figure 7. From figure 7, it can be seen that the leading- edge-flap deflection for maximum lift, the optimum deflection, occurs at approximately 30 for both the 6- and IO-perccnt- thick airfoils. No optimum deflection was obtained for the trailing-edge flap because the highest test deflection w

34、as still the most effective. The maximum section lift coefficients of both airfoils are approximately the same at corresponding flap deflections, but the increments of maxi- mum section lift coefficient obtained with flap deflection differ because of the lower maximum section lift coefficient of the

35、 lo-percent-thick airfoil with the flaps neutral. (See fig. 4.) Positive increments of the angle of attack for maxi- mum lift resulted when the leading-edge flap was deflected, JL but negative increments resulted when the trailing-edge flap was deflected (fig. 7). The pitching-moment characteristics

36、 of the two models (figs. 5 and 6) show that the aerodynamic center at low CQ (near the ideal lift coefficient) continues to move toward the leading edge as either the leading-edge or trailing-edge flaps are deflected. At higher angles of attack, the center of pressure always moves to the rear and c

37、auses the variation of pitching moment with angle of attack to become stable. The increments in angle of attack and lift coefficient at which this change in stability occurs show approximately the same variation with flap deflection as is shown in figure 7 for maximum lift. AIRFOlLS WITH FLAPS DEFLE

38、CTED IN COMBINATION The section lift characteristics of the two symmetrical circular-arc airfoils with the plain leading-edge flaps and plain trailing-edge flaps deflected in various combinations are presented in figure 8. The flap deflections that resulted in the highest maximum section lift coeffi

39、cient were R= 6X 108. ,I 278643-54-2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 REPORT 1146-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .8 i; E - .4 .s 0 4 16 24 -16 -8 SectIon angle of attack, czo, deg 8 16 24 (a) B-percent-thick airfoil. (b)

40、 lo-percent-thick airfoil. FIGL-RE 6.-Section lift and pitching-moment characteristics of two symmetrical circular-arc airfoils for various deflections of the 0.15-chord plain Icading-edge flap; R= 6 X 104. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS

41、-,-,-.I CIRCULAR-ARC AIRFOILS WITH LEADING-EDGE AND TRAILING-EDGE FLAPS 9 The section lift characteristics of the two models with the plain leading-edge flaps and plain trailing-edge flaps deflected 30 and 60, respectively, obtained at Reynolds numbers of 3 X 106, 6 X 106, and 9 X lo6 are presented

42、in figure 9. At Reynolds numbers between 3X108 and 9X10G, the data (fig. 9 (a) show no appreciable scale effect on the maximum lift coefficient of the 6-percent-thick airfoil. The section iift characteristics of the 6-percent-thick airfoil with the leading- and trailing-edge flaps deflected 27 and 6

43、0, re- spectively, are presented in figure 10 for Reynolds numbers from 0.70X lo6 to 2.29X106. In this range of Reynolds numbers, the maximum section lift characteristics of the 6-percent-thick airfoil are independent of scale. In the case of the lo-percent-thick airfoil (fig. 9 (b), however, some a

44、dverse scale effect (nearly 0.1) is indicated in the maximum section lift coefficient at Reynolds numbers between 3X106 and 6X10. Similarly, some adverse scale effect (fig. 8 (c) is indicated in the maximum section lift coefficient at Reynolds numbers between 3X lo6 and 9X lo6 with the leading- and

45、trailing-edge flaps deflected 36 and 60, re- spectively. At Reynolds numbers above 9X108, however. the maximum section lift coefficient of this combination remained npprosim.atel,y constant. -,!I - IO-percent thick alrfoil The section pitching-moment characteristics of the two airfoils with the lead

46、ing- and trailing-eclge flaps deflected 30” and 60”, respectively, (fig. 9) show that the aerodynamic center remains aheacl of the quarter-chord point for angles of attack greater than zero. In adclit,ion, the combined action of the leacling- and trailing-edge flaps caused the moment coefficients to

47、 increase negatively with increasing lift coeffi- cient until the angle of attack was high enough to alleviate the spoiler action of the leading-edge flap. As the lift coefficient was increased beyoncl this point, the moment became less negative until approximately 2.5 beyond the angle of attack for

48、 maximum lift, whereupon the moment curve breaks. LOW-DRAG-CONTROL FLAPS The lift and drag characteristics of the 6-percent-thick symmetrical circular-arc airfoil with the leading- and trailing- eclge flaps deflected arc presented in figure 11. Deflecting the leading-edge flap to loo decreased the s

49、ection drag coefficient of the 6-percent-thick airfoil at a lift coefficient of 0.3 about 40 percent by delaying the formation of a nega- tive pressure peak at the leading edge which causes separa- tion. In general, the leading-edge flap was more effective in extending the low drag range to higher section lift