1、1 , I NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS TECHNICAL NOTE No. 1671 EFFECT OF TAPER RATIO ON LOW-SPEED STATIC AND YAWING STABILITY DERIVATIVES OF $5 SWEPTBACK WINGS WITH ASPECT RATIO OF 2.61 ay William Letko and John W. Cowan Langley Aeronautic+ Laboratory Langley Field, Va. Washington July 19
2、48 , Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECHNICAL NOTE No. 1671 EiJjFECT OF TAPEZRATIO ONLCW-SI?E3DSTATICANDYG STABILITYDERATIV33S OF 45SiKlECBACKm WITH ASFEET RATIO OF 2.61 By William Letko and. John W. Sowan SUMMARY A low-speed wFnd-tu
3、nnel inveatition was made to determine the effect of taper ratio on the static and yawing stability derivatives of three tapered wings each with a constant sweepback of 450 at the quarter- chord line and with an aspect ratio of 2.61. The wings were of the NACA 0012 airfoil section in planes perpendi
4、cular to the quarter-chord line. The win had taper ratios of 1.00, 0.50, and 0.25. The results of the tests in straight flow indicated that the Lift- curve slope at low and moderate lift coefficients was almost unaffected by chasges Fn taper ratio. The aerodynamic center at low lift coefficients shi
5、fted rearward ll percent of the mean aeromc chord with a decrease in taper ratio from 1.00 to 0.25. For low and moderate lift coefficients, the rate of change of effective dihedral with lift coefficient decreased as the wing taper ratio decreased. The results of the tests in yawing flow indicated th
6、at the rate of change of rolling moment due to yawing with lift coefficient decreased 88 the wing taper ratio-decreased. The chsqe in taper ratio had little effect on the yawing moment due to yawing. In general, the trends Fn the characteristics resulting from effects of taper were indicated with fa
7、ir accuracy by the available theory. INTRODUCTION Estimation of the dynamic flight chmacteristics of airplanes requires a knowledge of the component forces and moments resulting from the orientation of the airplane tith respect to the air stream and from the rate of angular motion of the airplane ab
8、out each of its axes. The forces and moments resulting from the orientation of the airplane normally are expressed as the static stability derivatives, which are readily dete3xnined from conventional wind-tunnel tests. The forces and moments resulting from the angular motions (rotary derivatives) ha
9、ve generally been estimated from theory because of the lack of a convenient,experi- mental technique. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA TN No. 1671 The recent application of the rolling-flow and curved-flow principle in the Langl
10、ey stability tunnel has made possible the determination of both rotary and static stability derivatives with about the same ease. Preliminary tests (data obtained In the Langley stability tunnel) have indicated that althou measured parallel, to.axis of symmetry mean aerodynamic chord Lb, cq distance
11、 measured perpendicular to the Uds of sYnrmetry distance of quarter-chord point of any chordwise section from leading edge of root section distance from leading edge of root chord to quarter chord of mean aerodynamic chord aspect ratio (b2/S) Lb2 cx 3 taper ratio (Tip chord/Rootchord) an All the tes
12、ts were made at a dynamic pressure of 24.9 pounds per square foot, which corresponds to a Mach number of 0.13. The test Reynolds numbers based on the mean aerodynamic chord of the models varied from about l,lOO,COO for the wing with a taper ratio of 1.00 to l,23O,OOO for the wing with a taper ratio
13、of 0.25. The characteristics of the wings were determined in both straight and yawing flow. In the straight-flow tests, six-component measurements were made for each wing through an angle-of-attack range from about -4 up to and past the angle of msximum liftrat angles of yaw of O“ and %O. The tests
14、in yawing flow were made at en angle of yaw of 0 and for four different wall curvatures corresponding to the .values of rb/2V of 0, 4.0316, -0.0670, and -0.0883. In the yawing-flow tests each model was tested through an angle-f-attack range from about A0 up to and past the angle for maximum lift coe
15、fficient. . c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1671 5 CORRECTIONS Approximate corrections, based on unswept-wing theory, for the effects of the jet boundaries have been applied to the an however, the angle of-attack for max
16、imum lift coefficient is increased with a decrease in taper ratio. The effect of taper ratio on the pitching-mnt coefficients can be seen in figure 4. The position of the aeromc center for each wing was computed from these data for 1stcoefficients up to 0.6 and. the data are presented in figure 6 to
17、gether with the theoretical variation of the position of the aerodynamic center with taper ratio as obtained from reference 2. The experimental data indicate a rearward. shift (equal to about 11 percent of the M.A.C.) of the aerodynamic center as the tayer ratio is decreased from 1.00 to 0.25. This
18、movement is in fair agreement with theory; however, the experimental data in each case indicate a more rearward position of the aerodynamic center than that predicted by theory. For the wing with taper ratio of 1.00 (see fig. 4), a-re erward shift in the aerodynamic center of about ll percent mean a
19、erodynamic chord occurs at a lift coefficient of ( CL2 about 0.6 lift coefficient at which the drag increment CD - z starts to rls a . Only a small shift in aerodynamic center occurred at this lift coefficient for the ting with taper ratio of 0.50 and almost no shift for the wing with taper ratio of
20、 0.25. Some chmges in aerodpamic- center position, therefore, can probably be avoided through the proper choice of taper ratio. For these tests the stabilizing shift of the aerodynamic center encountered above a lift coefficient of 0.6 for the wing with taper ratio of 1.00 was minimized by decreasin
21、g the taper ratio. The effect of taper ratio on-the static lateral-stability derivatives can be seen in figure 7. The values of Ct $ increased almost linearly but at slitly different rates with lift coefficient for each wing tested and reached a llaximum at a lift coefficient ofabout 0.6. The maximu
22、m value of Cl * obtained decreased with a decrease in taper ratio, probably because the tip-stalling tendencies of sweptback wings are aggravated by a reduction in taper ratio. For low and moderate lift coefficients, the rate of change of effective dihedral C2 with lift coefficient CL decreased with
23、 a decrease in taper ratio! X and this variation is shown in figure 8, which is a plot of “lda a CL 0.6. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 NACA TN No. 1671 ./ c 4- o ./ ./ Cnr 0 , Figure Il.- I I I I 1 . -” - (S-4 - t-t7l-T Comparison of experimental and theoretical variation of C, with r lift coefficient for wings tested. V = 0. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
