REG NACA-RM-L8H12-1948 Yaw characteristics of a 52 degrees sweptback wing of NACA 64(sub 1)-112 section with a fuselage and with leading-edge and split flaps at Reynolds numbers fr.pdf

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1、RESEARCH MEMORANDUM YAW CHARACTERBTICS OF A 52O S-WFSTRACK WING OF NACA 641 -112 SECTION WITH A FUSEUGE AND WTTH LEADING-EDGE AND SPLIT FIAFS AT REYNOW M. . By Rein0 J. salmi Langley Aeronautical Laboratory Langley Field, Va. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON November 8, 1948 Pr

2、ovided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Lorspeed tests were made in the Langley 19-foot pressure tuzlnel to determine the aerodynamic characteristics in yaw of a 52O sweptback wing of aspect ratio 2.88 and taper ratio 0.625 with NACA im 43 perc

3、ent of the semispan, measured from the plane of symmetry. The fences were of cormtant height, being 60 percent of the maximum thickness of the local airfoil section, and extended. over the rear 95 per- cent of the airfoil chord. Tests The data were obtabed at Repolda nlmibera of 1.93 X 10 , 6 4 -35

4、X lo6, and. 6.00 X 10 6 with corresponding Mach nuaibers of 0.08, 0 -09, and 0.12, respectively. The stability derivatives were obtained from straight-line fairings of data obtained from tests at 00 and 250 angle of yaw. Extended angle-of-yaw tests were made at several angles of attack to cover the

5、yaw range from -50 to 25O Etngle of yaw. For the dng-alone testa the following flap configurations were used: (a) flaps neut-, (II) split flaps deflected, ana (c) split flaps and leading-edge flaps deflected with fences installed. For the wing- fuselage tests onlg the first and third flap config”ati

6、on8 were tested. Air-stream surveys were =de to dete-e the sidewash angles and dynamic pressures in a region approaiting the location of a vertical tail. The surveys were made with the Langle;g 19-foot tmnel 6-tube . rake (fig. 5) in a pke normal to the tunnel center ltne and 1.71 behind the center

7、of graviQ. (See fig. 6.) In some cmes, the sidewash sligles exceeded the values for which the rake hd been calibrated and extrapolations of the calibrations were necessary. The extrapolated values are shown by the dot-dash lines In the figure8 * All tail -surveys were made at a Reynolda number of 6.

8、00 X 106. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 CORRECTIOTS TO RATA NACA RM No. L8Hl2 The lift, drag, and pitching-moment data presented herein, have been corrected for support tare and interference effects and for air- stream misalinemen

9、t. The jet-boundary correctians to the angle of attack and drag coefficient were calculated frm reference 4, which accounts - for wing sweep, and are as follows: c The correction to the pitching-moment coefficient due to tunnel-induced distortions of the wing loading is: All of these corrections wer

10、e added to the data. No correctfons were applied to the rolling-moment, yawing-moment, and lateral-force coeff iclents RESULTS AND DISCUSSION The lift, drag, and pitchin;-mament characteristics for the wing with all flap canfigurations used Of the stall Which be On the baaing wing p-1 at a lift . co

11、efficient of about 0.90. .r baaing edge ocmd coincidentEtlly with IeaaTng-edge sepazation and an * . Directional stability and lateral force.- The plain wing had neutral directional stability at zero lift but gradua-lly increased in stability with increasing lift coefficient up to a CL of 0 .TO. Bey

12、ond this point, the directional stability decreased rapidly and the wing became direc- tionally unstable at a CL of 0.76. Tple instability aeem to colncide with the decrease in the slope of the curve. Although the wing became stable again at a CL of 1-03, it was unstable at the manlmum lift coeffici

13、ent- czv The lateral-force paramster Cy was negligible at lift coefficients .Ilr below 0 .TO but varied from a negative value of about -0 -0075 at a CL of 0.87 to a positive value of about 0.007 at the maximum lift coefficient. EPfect of Flaps on the Lateral-Stability Parameters Dihedral effect.- Th

14、e initial rate of Increase in effective dihedral with lift coefficient was slightly reduced by flap deflection. The maximum values of C2 were Increased, however, to 0 -0055 at a lift coefficient of 1.11 when the split flaps were deflected and to a value of 0.0065 at a CL of 1.29 when both the split

15、flaps and leading-edge flaps were deflected. The effective dihedral. remained at a large positive value at the maxirmrm lift coefficient when the leading-edge flap were deflectedj whereas with the plab w3ng and with only the split flape deflected, C became negative at the mum lift coefffcient. (Tuft

16、 surveys showed that the lea-edge flaps delayed the tip stall.) 9 z$ Directional stabiiity and lateral force.- Flap deflection extended the range of lift coefficient in which the directional stability increased with increasing lift. With the split flaps deflected, the wing became Provided by IHSNot

17、for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA F3f No. mHl2 directionally unstable at a lift coefficient of about 0 *95j and with both the leading-edge flaps and split flaps deflectad, the wing became unetable at a CL of 1.31. The large variations in directiona

18、l stability and lateral force which occurred at high lift coefficients for the plain wing were apparent also when the split flaps were deflected but were minimized when the leading-edge flaps were deflected. Fuselage Effects on the Lateral-Stability Parameters Dihedral effect.- The lateral-stability

19、 paramstem for the high-wing, low-wing, and midwing. cmbfnations are glven in figure 9 for flaps neutral and in figure 10 for flaps deflected. The low-wing combination had less dihedral effect than the wing alonej but, as the wing position was pro- gressively changed from low-w- to high-wing positio

20、n, the dihedral effect increased. The incremnt of increase in Cz between the low-wing and midwing combinations was about equal to that between the midwing and hi wing. The value of the increment in C at zero lift was about 0 -000 P with flaps neutral, and 0.0007 with the leading edge and split flaps

21、 deflected. The slopes of the C curves for the wing-fuselage combi- nations were slightly lower than for the wing alone for the flaps-neutral condition. The dihedral effect due to the midwing posi-t;ion was very small a8 had been expected. In general, the effects due to the fuselage were of the same

22、 itude as had been experienced on other sweptback $ and straight 1 and 3). Directional stability and lateral force.- The fuselage decreased the directional stability of the plain wing by an increment in which c”* varied from about 0 -0012 for the midwing canibination to approxktely 0 .OOl5 for the l

23、ow-wing combination. The Fncrement in Cn was almoat constant throughout the lift-coefficient range except when the leading edge and split flaps were deflected on the low-wing colriblnation; then a large poeitfve value of C occurred at zero lift, reducing the deetabilizing * y* coefficient, and at a

24、CL of about 0.75 this reiieving effect became negligible The midwing combination had the leaat side force of the three combi- nations, whereas the low-wing cambination had the greatest. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. L8Hl

25、2 9 Effect of Scale on Lateral-Stability Parameters A large scale effect was noted in the lateral-stability parameters for the plah wing when the Reynolds ntmiber was increased from 1.93 X 106 to 4.35 X 106, as shown in figure ll. (The scale effect in the range of Remolds number from 4 35 X 106 to 6

26、.00 X 106 was moderate for the stability parameters, except for C% at high lift coefficients. 1 The manlmlrm value of C at a Reynolds nWer of 1.93 X 106 was about one- half its value at R = 6 -00 X 106 and occurred at a much lower liFt coefficient. The directional stability and aide force were affec

27、ted in a sfmilar mer. 2* A very similar effect of Reynolds nzmiber waa observed for the 42* swept- back wing of reference 1, and consequent* it appears advisable to exercise caution when usfng lateral-etabiliQ parameters obtained at low Reynolds numbers, especially in the moderate to high lift range

28、 on swept wings with conventianal airfoil ahapes When the leading-edge and split flaps were deflected, the scale effect waa negligible throughout the range tested. Characteristics in Ertended Yaw Range The largest deviations of the stability parameters at high yaw angles frm those measured at tmaJl

29、yaw angles were obtaFned in the ana % variations at high angles of Bttack. At an angle of attack of 16.80 the various configurations with flaps neutral (fig. 13) shared a reversal tn slope for the variation of the yawing mament with angle of yaw at an angle of yaw of about 100, tending to make the m

30、odel less unstable. At an angle of attack of 23.30 (fig. 15), the directional instability due to the -elage increaaed rapidly between the yaw angles of loo and 13 when the lea-edge and split flaps were deflected. AIR-FLOW CHARACTERISTICS IN !EKE REGION OF A VERTICAL “L It is pointed out in reference

31、 6 and shown in reference 1, that the sidewash angles in the region of the vwtical tail may be affected by the wing-tip vortices when an airplane of low aspect ratio is yawed, especially at high lift coefficients when the vortices are strong. Provided by IHSNot for ResaleNo reproduction or networkin

32、g permitted without license from IHS-,-,-10 NACA RM No. mEl2 The results of the air-stream surveys (figs 16 to 19) show that the sidewash angles due to yaw are appreciable at the higher lift coefficients even for the plain wing, but they are only very slightly more negative than those for the 4.20 s

33、weptback wing of reference 1. The aspect ratio of the 42 sweptback wing was 3.94 as cmpared with 2.88 for the Po SWO- back wing discussed herein. The variation of the sidewash angles and dynamic pressure ratios at the tail with height above the fuselage center line were very similar to those for the

34、 42O uing, indicating that the effect of the wFng vortices is essentially t?m ame for the two wings. Unfavorable sidewash and wake characteristics occurred at the high angles of attack for the high-wing canbination, but these effects diminished 88 the wing position became lower. The greater height a

35、bove the wing wake and the end-plate effect of the wing on the fuselage vortices (as explained in reference 5) cawed the low-wing cambination to have favorable sidewash near the fuselage and also at the higher points. The effect due to deflecting the leading-edge flaps, and split flaps in cdination

36、with the fences wa8 to reduce slightly the negative side- wash angles and to came the decrease in qt/q due to the wiw wake to be more severe. This same effect due to flap deflection occurred on the 42O sweptback wing. 1 The results of an investigation of the aerodynamic characteristics in Yaw of a 5

37、2O sweptback wing Fn caznbination with a fuselage may be summrized as follows: 2. The combination of leading-edge flaps, split flaps, and fences extended the range of increase of effective-dihedral parameter with lift coefficient up to the maximum lift coefficient and increased the directional stabi

38、lity to higher lift coefficients. In all cases, hugever, the wing was directionally unstable near the maximum lift coefficient. 3. The low-wing and high-wing comBinatione had lower and higher dihedral effect, respectively, than the wing alone, and the magnitudes of the differences were comparable to

39、 those experianced on uncwcpt wings. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No mEl2 . 4. With flaps neutral, a large scale effect occurred for the lateral-stabiliQ pabmeters in the range of Reynolds rimer from 1.93 x 106 to 4.35 X 10

40、 and a very moderate effect in the range .- from 4 35 X 106 to 6 .OO X 106. With the conibinatim of l“-edge flaps, fences, and split flaps the scale effect was negligible. 6 5. The results of air-stream surveys shmd that the most favorable sidewash characteristics for directional stability occurred

41、for the low- wing combination and were about the 88me as those obtained rn a h2O sweptback wing. Langley Aermutical Laboratory Rational Advisory Committee for Aeronautics Langley Field, Va., 1. M, Reino J., and Fitzpatrick, J-8 E.: Yaw Characteristics and Sidewash Angles of a 42 Sweptback Circular-A

42、rc Wing with a Fuselage and with Isading-Edge and Split Flaps at a Rep01b IYumbe? of 5,300,000, KACA RM no. L7I30, 1947. 2. Salmi, Reino J., Cm, D. William, and taper ratio = 0.625; area = 4429 sq in.; E = 39.97 in. No dihedral or twist. (All dimensions in inches.) 0 Provided by IHSNot for ResaleNo

43、reproduction or networking permitted without license from IHS-,-,-NACA RM No. BEL2 Figure 3.- Geometry of flaps and fences for the 52 sweptback wing. (All dimensions in inches.) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure 4.- 52 sweptback

44、wing and fuselage mounted in the Langley 19-foot pressure tunnel. Low -wing configuration; fhps deflected. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,

45、-,-RAGA RMNo. L8H12 4 (a) Photograph of rake head. (b) Sketch of tube head. Figure 5.- mngley 19-foot pressure tunnel air-stream survey rake, Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking p

46、ermitted without license from IHS-,-,-NACA RE4 No. L9m2 19 I I I I- I I I I 1 /0 5“ oa * SeCfiOD A-A Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.04 0 :04 -.08 -.I2 cm -557 ru 0 Z E z 0 Z r! . . . . . . . . . Provided by IHSNot for ResaleNo repro

47、duction or networking permitted without license from IHS-,-,-. . . i.4 1.2 I. 0 .8 .6 CL .4 .2 0 -2 0 .04 .08 .lP ./6 .20 .24 .P8 .3P -36 .40 .44 .48 .52 .56 CD . v (b) CL plotted against CD. Figure 7.-. Concluded. E . I Provided by IHSNot for ResaleNo reproduction or networking permitted without li

48、cense from IHS-,-,-22 NACA REI No. LE3EL2 r- . I I, - I ,/ I “_ - “ “ - “_“ y I / Flaps neutral Split fhps deflected 0.575b/2 1.e. flaps and split flaps deflected with fences installed - -“ -“ Figure 8. - Variation of Czq, and C with CL for a 52 sweptback cn IJJ y$ wing with three flap configurations. R = 6.00 x 10 6 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,