NASA NACA-TR-1071-1952 Theoretical symmetric span loading due to flap deflection for wings of arbitrary plan form at subsonic speeds《在亚音速下 任意平面机翼襟翼偏转引起的理论对称翼展载荷》.pdf

《NASA NACA-TR-1071-1952 Theoretical symmetric span loading due to flap deflection for wings of arbitrary plan form at subsonic speeds《在亚音速下 任意平面机翼襟翼偏转引起的理论对称翼展载荷》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TR-1071-1952 Theoretical symmetric span loading due to flap deflection for wings of arbitrary plan form at subsonic speeds《在亚音速下 任意平面机翼襟翼偏转引起的理论对称翼展载荷》.pdf（41页珍藏版）》请在麦多课文档分享上搜索。

1、REPORT 1071THEORETICAL SYNUW2TIUC SPAN LoG E To ELAP ELEcTo FOR s OFARIWIRARY PLAN I?ORNI AT SUBSONIC SPEEDS 1BY JOEX DEYOENGSUhlhIARY.A imPled ijiing-wrface them-y is appiied to the problem ofemrlwting span loading due to flap deject ionjor arbitrar.r.rwingplan forms. U“ith the resulting procedure,

2、 the e$eck of j?apde.fl.ection on the span loading and associated aerodynamicchuracten”stics can be ean”ly computed for any wing which isgYmmt”calabo Ut the root chord and which has a straight quar-ter-chord line owr tht wing semispan. The C$PCLSof cornpresR-biity and spanwiw criation of section lif

3、t-curw sopc aretak(m into account by the procedure.For the case f straight-tajwred u-, a sweep parameterdefinvi as .i8=tan-1 (tan .i),lp ancI a w-chord distributionptirarneter H, defined by(qJ .+s discussed fn reference I, a, is not limited h small values but cm be ss Iarge sn em.zleas dwired, provi

4、ded separation does not omur.c The efieets of mmpres$l%fliw errd seecion lift-eume elope ere eqDfmIent to a change inwing plan hrme and em be mken “intoaceouni by a proper adjustment of the a,. wdueswhereK, ratio of experimental section lift-curve slopeat span station v to the theoretical -ralue of2

5、T, both at the same JIach numberPc, *U chord at span station vd, scccIefactor which has the following values:0.061 for v=l.234 for v=?.381 for P=3.320 for v=4Equation (2) can be w-ritten in an alternateH, in terms of U-LP geometry parameterssignificant; thusUH,=d, _ 1, ( A and theequivalent, angle-o

6、f-attack distribution for unit flap deflectionaJtiL for each case is Jisted in tabIe .44. The -raIues of aJ81Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24S REPOR 1071-NATIONALADvIs0R3” COMMITTEE FOR AERONAUTICSgiven in table .44 are for (cl/c)=

7、1. The results can be easilymodified for the case of (cf/c) 2. This limit to whichda/d8 can be used will be considered further in the sectionDiscussion when comparisons with experimental results arcmade, ._3. Arbitary span.wise distribution of jlap cford cJc) =zmrialdcl: The flap can be divided into

8、 several parts eachhaving constant da/d and the load distribution due to eachpart determined. The total distribution is then t.lus sum_ ofthese individual load distributions.Lift coefficient,-The lift coefficients due to flap deflectionfor the same flap configurations previously discussed cau befoun

9、d as follows:1. Full. wing-chord jfaps (c,/c) = 1: The spanwise loadingdue to flap deflection is, in general, so complicated thatProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-THEORETICAL SYMMEIC SP.0” LOADING DUE TO FLAP DEFLECTION FOR ,WLWGS ATntl

10、nwrical integration based on four spanwise values cannot.br perfomed aceuratelj- with conventional integrationformulas. However, in appemlix A a spwial integrationformula (which needs but four spanwise dues) is developedwhich applies to the dfierent flap spn. Equation (.A6)can be written aswhere, fo

11、r each of the cases of equations (4) and (5), the hmvalues are gi-ren byhl 0.293 0.299 0.300 0.352 0.301h, a 301 0.301.544 .544 .555 .561 .3.56 .536h-35s.725 .725 .734 .Z?2 . .726 .ixh, .392 :% .392 .105 .395 .393 .393,.2. Constant fracfion of un.ng-chord yaps (cf/c) =consfant:The flap effectiveness

12、 is given by(7),. Arbitrary span wise Wribufion oj jZap chord (c.t/c)=wria ble: Flaps for which cf/c varies spamvise on the wingcan he considered as equivalent to a wing-twist distribution.The effective s-ymmetric twist of the wing is gien by(8)whtw daldfi is now a function of spa.nwke position. Ift

13、,quation (8) is a continuous distribution such that it can beplotted by specifykm its value a fo semispan points, thenthv wing can be considered to be twisted and solutions founddirmtIy as for basic loading.?Ilen equation (8) is discontinuous, the ae of attackcan be divided into spanwise steps of eo

14、mtant. angle ofattack and the total lift can be found by the summation ofthe lift due to each spanwise step. The lift of a spamvisePC.,step is obtained from a curve of Lift. coefficient as arequired to obtain spanwise center of pressure and induceddrag are not. dewloped here. Eowewr, the integration

15、formulas given in refererice 1 to obtain these characteristics _can be used -with acceptable accuracy for the case of loadingdue to flap deflection.No rigorous procedure for determining pitching momentscan he developed from this method. As noted in reference1, ho-wever, a good approximation for a wi

16、ng without flapscan be obtained if the wing sweep is Iarge since moments dueto shifts of spanwise load overshadoxr those due to shifts ofchordwise load. Vilere the effect of flaps is to be considered,as in the present case, it ppears unlikely that the shifts inchordwise loads can be sa.feIy neglecte

17、d. Therefore anestimation of pitching moments due to flaps should not onlyincIude the moment due to spanw-ise load redistribution,which this method -wilIgive accurately from the longiudinalmoment of the center of load, but also should include anestimation of the moment due to chordwise redistributio

18、n.Additional considerations,Several items pertaining to -the usage of the method should be considered.1. Addifice nafure oj loading and spanwise angle of attack:The Jinear relation between angle of attack and loadingdistribution of equation (1) states that all loadings areadditive if the respective

19、angle-of-attack distributions areadded.Thus, the aerod.ynamie coeflkients that restit from theintegration of the Ioading distribution are additie if theydepend linearly on the loading distribution. Hence, lift.coefficient, rollingmoment coefficient., and pitching momentare additive; whereas induced

20、drag, spant+e center ofpressure, and aerodynamic center are not.The additive concept is very useful for the determhationof loading due to flap deflection. The loading due to anarbitrarF span flap having arbitrarF position on the wingsemispan is simply found by adding or subtracting knownIoadings due

21、 to flaps at other Iocations on the wing. Forexamde, if the spa-nwise ends of the flaps ar at the same.,span stations, thing-clord inboard flaps for flap spans measnred from theplane of symmetry outboard. The lift clue to outbomdflaps for flap spans measured from tbe wing tip inboard canbe obtained

22、from figure 5 by use of the relations of equation( 10). For fuII wing-chord flaps located arbitrarily on tlewing semispan, the lift can be obtained from figure 5 as in-dicate in the following example sketch:Throughout the figures. . . . is the comtant spanwise-section lift-curve sloI)edr the aerageo

23、f a small variation. For Iarge spanwise variations ofKthat foowthe function given in cqua.tion (B5)developed m appendix B, tbe pzrmneters 5.4/x_ and x can be reacedby theparameters - and X,respectively.For large spanw-ise wriatfons of K thst do not follow the curveof equation (B5), the sinm.taneouse

24、quations fortbegenend solution cm besolmdfor arbitrary distributions of., vauesof H, can be obtained conveniently from figure zWith the full wing-chord vtdues given above, tllc lift dueto conshant- fraction of wing-chorcI and flaps of arbitraryspanwisc chord distribution can be lound through Ilsc of

25、equations (7) and (9) with the da/d it, is shovn I1o;various forms of symmetric. Ioading can be found quicklythrough use of simplified lifting-surface thcoI y and vtilues ofthe a,m coefficients which are presented in graphical form.It can aIso be sho;vn that if the symmetric. load distributionis kno

26、wn, from any computational method or from experi-ment, then the same a. coefficimts can be used to find cor-responding values of the downwash in t.ho center of the own-wash wake. It is the purpose of this section of the reportto make available the method for completing dcmmwashwithout att wupting to

27、 explore the possibilities or limitationsof the method. In vimv of the importance of being al)k t _predict, downwash, it is considered that inclusion of this matwrial herein is justifiable although not directly related to themain purpose of the report. It is important to rememberthat the method can

28、bc used for any case of symmetric lofid- _ing and is not restricted to that discussed in this report.As noted previously, the a,. coefficients provide a verysimple means of finding the clew-nwash in the center of thedownwash wake of the wing due to any symmetric spa.nviseIoading_ distribution for an

29、y sweep angle. The downv-ashcan be found at spcrific points in the region bctirecn thowing tips from near the quart er-c.herd line t o an in fhlit edistance downstream,Equation (1) was derived (referece 1) to fintl the down-wash at the effecti-c three-quarter-chord line for which thedistance downsre

30、am from the quarter-chord line is K(c/z).The gemeral expression for the distance downstream fromthe quat,er-chord line isxy tan A$where z is the longitudinal distance measured in the hori-zontal plane from the Jving-root quarter-chord point to anarbitrary downwash point. Substituting this general CX

31、-pression .in equation (2) givesd,11($/0) 7Ptan ith theH, values given by equation (13).It should be noticed that, because thc chordwise loadingis concentra-ted at the one-quarter-chord line, equation (13)is independent of aspect ratio and spanwise distribution ofwing chord, w-bile from equation (14

32、) it can be seen that,dowmvash depends on these wing geometric values onlythrough the loading distribution (7.,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-THEORETIC.4L SYMMETRIC SPAN!ilquatio; (14) fl give the downwasbLOADING DUE TOange le to ari

33、.t rary s-mmetric loading distribution for all values of $ q,tan . for .42, as dcussed earlier. However, from thelinearized compressible-flow theorj- the aspect ratio becomesa-n effectire aspect ratio giren by B.-1.which approaches zeroas M+l .0. Thus the use of two-dimensional da/d thus a deflected

34、 flapresuIts in a discontinuity in the spanwise distribution ofangle of athldi. Spanwise loading for a discontinuous dis-tribution of angle of attack can he found through the simul-taneous solution method of equation (1). The number ofequations, however, must be greater than four to obtainaccurate r

35、esults and then the computation efforts becomeexcessiw. (Reference 1 gives the method for an arbitrarynumber of equations.) llowever, an alternate procedure,originally developed in reference 2, provides sufficient accu-racy with a reasonable number of equations and this proce-dure is the basis of th

36、e following development.The zero-aspect-ratio theory of rcferce 3 shows thtit fora given sweep tmgIe, as aspect ratio approaches zero, span-wise loading becomes independent of plan form. Thus, thealtered boundary conditions (see Development of lklet,hod)required for the presen theory to give spanlvi

37、se loading forgiven flap spans can be determined from he loading givenby reference 3, which applies to all plan fortns. Thesealtered boundary conditions are obtained from equation (1)using the a,. values for zero aspect ratio and the loading dis-tribution for zero-aspect-ratio Wings given by referen

38、ce 3for gicn flap spans.The zero-aspect-ratio values of a., are given ly figure Ifor 11,=0 or are given by the 13,. coefficients of refwwnee 1.These a, coefficients are t.abulatd as follows:TABLE Al% (FOR nt=7 AhTD A.=0),+A 41-10.4624 2.0720 -0.22422 3.8284 5.656s3. 0 -: 62S42.3ss3 4.329EL.- -:.154s

39、-.2928 0 -1.7072 4.0000With these ar coefficients and equation (1), symmetricspanwise Ioading can bc found for zero-aspect-ratio wings.As a comment on the accuracy of the simplified lifting-surface theory for determining additional loading Widlm=7, the solution of equation (1) with .4=0, a values ga

40、ve,to the accuracy of three decimal places, the c,lliptic loadingdistribution characteristic of zero-aspect-ratio wings.The flap end is arbitrarily chosen for the presen t.hcoryas the mean value of the spanw-ise trigonometric coordinatesof the points at which the boundary condition is appliccl findw

41、hich bracket the discontinuity or fiap spanwise end. Thusfor m=7, three flap spans can be defined for both inboardand outboard flaps. Let q, be the flap span, and 6 the spnn-wise point at the end of t.h ihpj thenqr=cos o for inboard flapsv,= 1cos 0 for outhoad flapsThese flap spans are given in the

42、following table:TABLE A2DEFINED FLAP SPANS.Inboard OutbofirdI COW -11111 111, IV Y W“-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-THEORETICAL SYMMETRIC SPAN LOADING DUE TO FLAP DEFLECTION FOR WINGS AT SUBSONIC SPEEDS 253For the flap spans listed

43、in table .42, the exact span load-ing distribution for zero.pect,ratiocan be found fmreference 3. With the a,s values listed in table Al and theexact values of Q1, (72, G3, and G4determined from reference3, equation (1) gives the boundary condition or twist re-quired for the present. method to produ

44、ce the loading dis-tribution for each case listed in table .42, that is,;=10.4524 $3.8284 O.2928 =_.0720 +5.6568 ! ,.,1. 111 w v“ “ jw+- _h, 0.2991 0.2994 0-2929 0.30;0 o.3)14 0.3009-,;hj .5541 .5544 .5549 .5605 .5563 .5556h .7248 .7250 .7252 .7339 .7275 .7259-Qh .3922 .3921 .3922 .4050 .3950 .3930

45、INTERPOLATION FUNCTION FOR ,SPANWISE LOADING DISTRIBUTIONThe method of this report gives spanwise loading due toflap deflection at four span stations. Since the completeloading distribution is not sufficiently defined by these values,the values of loading at otJcr span stations cannot be inter-polat

46、ed accurately by direct use of the interpolation equationand table of reference 1,The direct use of the interpolation table of reference I canbe obtained by converting the present loading due to flapdeflection to a loading approxtimaing additional Ioading dis-tribution. The conversion factors are ob

47、tained from valuesof the zero-aspect-ratio theory and given by the ratio ofequation (Az) to sin pn for each of the four given spanstations. Define ,=(G/ 0:0229_ _._. -=- . .,=.-L. -. :?. - .-The interpolated loading at span station k is derminedfrom the summation_ . ()=n+-enk - (A9), ,. $:-where equ

48、ations (A2) and- (A8) provide values of Rk vhichare tabulated as follows:.-. . TABIJ3,+8 .,. -.- .-4_46 .8201- .,/7 .135i.3324 .7882 .9556 . M42 .2136 .? = _With the Rz values of table i18 and the loacIing ratios determined from equation (9), the loading G, can be foundat span stations k=;, $ and (q= O.981, 0,831, 0.556and 0,195, respectively).APPENDIX BAIDS TO THE METHODLINEAR ASYMPTOTE OF u,For large vahms of H, the a,n functions become linearlyproportional to /+, (see,e.g,fig.1). Since this linear charac-Provided by IHSNot for ResaleNo reproduction or