NASA NACA-RM-L9F14-1949 Effect of sweepback on the low-speed static and rolling stability derivatives of thin tapered wings of aspect ratio 4《后掠角对展弦比为4的薄锥形机翼低速静态和旋转稳定性导数的影响》.pdf

《NASA NACA-RM-L9F14-1949 Effect of sweepback on the low-speed static and rolling stability derivatives of thin tapered wings of aspect ratio 4《后掠角对展弦比为4的薄锥形机翼低速静态和旋转稳定性导数的影响》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L9F14-1949 Effect of sweepback on the low-speed static and rolling stability derivatives of thin tapered wings of aspect ratio 4《后掠角对展弦比为4的薄锥形机翼低速静态和旋转稳定性导数的影响》.pdf（37页珍藏版）》请在麦多课文档分享上搜索。

1、RESEARCH MEMORANDUM- EFFECT OF SWEEPBACK ON TKEZ LOW-SPEED STATIC AND ROLLING STABILITY DERIVATIVES OF THIN TAPERED WINGS OF ASPECT RATIO 4 William Letko and Walter D. Wolha.rt Langley Aeronautical Labratory Langley Air Force Base, Va. 5. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON August

2、 9, 1949 C I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECT OF WEEPRACK ON THE LOW-SPEED STATIC AND ROLLDG STABILITY DWIVMTVES OF TECN TAPERED WINGS OF ASPECT RATIO 4 By William Lstko and Walter D, W03hm-t SUMMARY A 1ow”speed investigation wa

3、s mde in the IangLey stabilltg tunnsl to determine the effect of sweepback on the static and rolling stability derivatives of a series of Uings, each of w-hich had a taper ratio of 0.6 and an aspect ratio of 4. The were of RACA 65A- section in planes parallel to the axie of symmetry asd had mepback

4、anglee of their quartemhord lFne of 3.6O, 32.6O, and 46.70. Most of the tests were =de with the wbgs in L.ameters cm the rolling etability derfvatives of wings have been fnvestigated inthe Langley stability tunnel by rueam of the roll-flox technique. (See reference 1.) The irrvestitiona have include

5、dthe effects of mpect ratio and mmep (refemme 2), taper mtfo (reference 3), dihedral (reference 4), and airfoil section (reference 5). AU of the vestl- gat ions were performsd at low Mach number8 and vith moderately thick wings. In order to obtain an indication of the rolling chamcterietice of swept

6、back vlngs at higher SUbSoniC speeds, a seriee of thin wing6 (XACA 65006 airfoil section) were tested in the Langley high-epeed 7- bg 10-foot wind tunnel at Mht flow and at Oo angle of yaw in rolling flow. For the straight-flow teeta at Oo angle of yaw, lift, straiboundary corrections (6LmUar to tho

7、se of refe- ence 7) based on unsmptrwing theory have been applied to the angle of attack, the drag coefficient, and the roll-nt coefficient Correctiona for blocking or support-gtrut tare8 ham not been applied to the results, Straight-Flow Characteristics The lift, drag, and pitch-nt chmacteristics o

8、f the three wings, each tested in caniblnation with the fuselage, are presented in figure 6. The pitch-nt results at low lift coefffcfents indicate that the aerodpauic center moved reazwaml, From 17.6 percent to 27.0 percent of the mean aeroagnamic chord, as the angle of sweep back was increased frc

9、m 3.60 to 46.70. The theoretical results given in reference 8 predict ahost no change in the aerodynamic-center location of plafn wing6 over this of sweep angles for the micular aspect ratio and taper ratio of the wings investi,eti. TEJ differences betmen theory and expriment -po%ably resulted from

10、the fact that a fuselage was used in the tests. Because each of the wfngs was comtructed in two semfspm segments with mounting blmka at the inboard ends 9 or attachment to a helage, true wing-alane chmacteristics could not be obtained. An attempt to simulate, as nearly as possible, the winmom caditi

11、on was made, however, for the 46.70 sweptback wing. The dng segments were supported by cover plates and the entire root regian was faired with balsa wood and clay. (See fig. 5.) Lift and pitching-moslent results obtained Kith this model (wing alone) and with the SEI wing in Provided by IHSNot for Re

12、saleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM LgF14 canibination with the fueeLage are ccmpwed in figure 7. The fwelage appeared to have very little effect on the general. shape8 of the lift and pitching+nament curves or on the mrodpamlc-center 108tiOn determinsd

13、 from the slope of the pitchiwnt cme at zero lift. For either the wing alone or the w%qpfueelas ccmibination, the aem dynamic center was anl;y about 1 percent of the mean aerodynamic chord behFnd the location (27 percent of the mew aerodynamic chord) given by the theoq of reference 8. Apparently, fo

14、r the 46.7 mptback wing the forvard location of the wing-fuselage JWture msulted in elimination 5f the usual unstable pitc-nt contributian of the Azsebge. , ?or the wings with emaller sweep angles, the location of the wing- ?uselage junsture was farther rearward and, in them cases, the :ontributicm

15、of the fmew to the pitching-mment characteristics . mems to have been a destabilizing effect, as ie normally expected. iuch an effect (an Increase of the -table pitchbg+ucment contribution If the Rzselage with a rearward shift of the -1- juncture) as found in testa of mid- configuratians with Eltrai

16、ght vings eported in reference 9. The reeulte of reference 9 for a midving onfiguration show that as the location of the quartelrchord line of he wing with respect to the fuaelage varied fram 9 to 44 percent of h8 fuselage length, the taerodynamic-center location of the configw sation varied fram 0

17、to about 6 percent forward of the location for ing alone. For the 3.6O mptback King with fuselage, the srodynamic-center location (17.6 percent of the mean eLer-c :hord) was 7.4 percent forward of the location mdicted by the theory )f reference 8 for the wing done. The results preHnted 5n figure 7 a

18、how that removal of the fusel= awed a reducticm in lift-curve elope (from 0.062 to 0.054) near zero. ift ; but even with the fusela removed, the lZft4urve elope m8 lightly higher than the theoretical value (0.052) even in refereme 8. he small displacements af the lift and pitchlrq+mmnt cmes for the

19、)lain wlng, relative to the cmee for the *fusela canbination, robably resulted frm s(llly9 camber introduced by the fairing of the :enter sectiob of the Xing. he lift data presented in figure 6 indicate an incmase u maxirmrm lift coefficient from 0.80 to 1.02 as the mepback is increased fram 3.6O to

20、 46.7. This result is in agreement w3th the ffndinga of another law-scah investigation (reference 10) and has been confirmed for Repolds numbers a8 high as 12 x 10 in a recent 6 investigation (unpublished) of winge having geametric properties ahoat identical to thoae used for the preeent Fqvestigati

21、on. At lift coefficients below 0.8, the lift cmes for the three engs are very nearly the same. Although the theoriea of reference8 8 and ll do predict a reduction in lift-curve elolpe of plain vi= with Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-

22、9 . increased eweep angle, such a, redudion, if it occurs, would be expected to be confined to a very am.ll range of lift coefficients (from about 4.2 to 0.2) for the present models, because above a lift coefficient of 0.2 (somewhere between 0.2 and 0.3) psrtial separation appears to tak3 place. TEI

23、 separation is Indicated from the comparison of the axprimontal drag CUTVBB with the curve obtained by adding the drag at zero lift to. the theoretical induced drag for elliptic wings of aspect raiAo.4. (See fig. 6.) For each of the wings the experi- Illental drag curve began to depart from the theo

24、retical relation at a lift coefficient same however, thie initial trend would be eqected to be maintained only mer the rw of lift coeffiolents for which the total drag is app?oxlmately equal to the drag at zero lift plm the induced drag. (See reference 2.) As is Indicated by the drag data of figure

25、6, this oonditian is satisfted only up to Ilft coafficiente of about 0.2 or 0.3. At ekh low lift .coefficients the ragnitudes of the theoretical values of the pwhg moment d.ue to rolling probably are withln the experimental accuracg of the rpeaems and, therefore, no initial nega.tive slope could be

26、detected. P The experimental results for the derivative C are compared ?e in figure 14 with remits calculated by a method (presented in refelr- erne 2) which includes cmsfderation of the drag easured under straight-flow conditions. In general, fafr agreement is obtained, altho- tlie predicted valuee

27、 of Cnp at high Uft coefficienta are too highly positive for the 3.60 and 32.6O aweptkk vi-. In refex- erne 2, through analysis of experinusntal da, the increment of C due to profile drag was found to be woportional to the slope of the curve of profile drag pldted agaimt angle of attack, and, the co

28、nstant of promioImlity was found to vary with mpect mtfo but to be essentially indspendent of the sweep angle. The ccunpa.rison presented llp Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-k canparison of value6 of C obtained by tbe roll-flow techni

29、que of the Langley stability tunnel with those obtained fram the free“rotatiop tests of the models in the hagley 7- by 109oot tunnal (referance 6) is psented in figure 13. Ijl general, the variation of Cz ulth lift coefficient is similar, and tho retlues of Ct are in good agreemnz. The Langley 7- by

30、 lO-foot tunnel reeulte are slightly higher, but tu8 difference can be attriblrted eilmrst entirely to the difference in Maoh number of the teats, as i indicated in figure 16, which ccmrpa;res experlnmnteil result6 obtained by the two technique8 with theoretical results (from reference 16) correerpc

31、rnding to the two t-t Mach runbere. The difference between the two theoretical curve8 is slmoet exactly equal to the difference between the two experhental curves. Both exprimental techniques yield values that we consistently larger than the theoretical VBlues, ectiveness parameter (pb/2V)B, Provide

32、d by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-therefore, is determined primarily by CJ All of the wings show reductions in rolling effectiveness as the lf9t coefficient is increased up to about 0.5. In the cam of the 46.7O Bweptback wing, this reduction a

33、mourrts to about 4-Q percent of the value at zero lift. At higher lift coefficients (b/2V) increase8 for all of the wings became C 2 decrease8 more rapidly than Cz8. The value8 of ( pb/2V)E presented b figure 18, however, (as previouely mentioned) neglect any possible effect of aileron deflection an

34、 the nRmping in roll or of rolling on aileron effectivenese. P P An investigation made in the Langley stability tunnel of a seriee of thin sweptback wings of aepect ratio 4, each tested in combination with a fuselage, indicate8 the following conclwiom: 1. The mxhm lif% coefficient of the -fwelage cc

35、adbinationa increased as the angle of weepback increased. At lift coefficients below 0.8, the lift cmes were- very nearly the 8- for all three models. The Uual effect of sweepback in reductng the lift-curve slope- appeared to be confined to the lift-ooefficient range between about 4.2 and 0.2 but wa

36、a leee than W but for the 3.6O aweptback wing, the fueelage is believed to have a deetabilizing effect, as le usually expected . 3. At low lift coefficients the derivative of rolling moment due to yaw varied linearly with lift coefficient , and the rate of variation increased wlth an increase in swe

37、ep angle in very mch the manner that is predicted by theory. The linear VariatioG were maintained mer only very -11 ranges of lift coefficient for the more highly ewpt win; however o Mean aerodynamic chord, ft 0.763 Taper ratio. . 0.60 Root chord, ft . 0.938 Tip chord, ft 0. s3 . Aileron : Typ5 . .T

38、rue contour, sealed g&B Chord, percent c 20 span, percent b/2 . 40 Inboard stabion, percent b/2 . 55 Outboard stat ion, percerrt b/2 . 95 Figure 2.- Sketch and dimensions of wings teated. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 WCA RM I?lh

39、 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I Fiaure 4.- The 46.7 aweaback wing with fuaelage mounted in the rolling-flow section of the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided b

40、y IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LgF14 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 F1 .gum 6.- Variationof the drag and pitchlng-momnt coefficiente and angle of attack fdth lift co

41、efficient for wing te&ed Kith a fuselage. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c Q c Figure 7.- Variation of the pitchingaronusnt coefficient and the angle of attack with lift coefficient for the 46.7 sweptback wing alone and for the 46.70

42、 sweptback wing in coaibination with the fuselage. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-26 NACA RM L9F14 $2 0 Figure 8.- Variation of the Lift cmfficient wfth angle of attaok of the 46.7O sweptback wing alone for vczrioue valuss of Reynold

43、s rider with and without transition strips cm ving leading edge. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WCA RM LgFl4 0 0 Figure 9.- Variation of the pitchin-mnt coefficient with lift coefficient of the 46.70 sueptback wing alone for.various

44、values of Reynolds number with and without transition strips on wlng leading edge. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 RACA RM L9Fl. , Figure 10.- Effect of trmmition strips and Repol& number on the lift- curve elope and location of th

45、e aerodynamic center for the 46.7O eweptback wing alone. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LgF14 # Figure U.- Variation of C C and C with lift Goefficient for YJIJ nq 2J( the wings tested with a fueelage. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-