REG NACA-TR-705-1941 Wind-tunnel investigation of effect of interference on lateral-stability characteristics of four NACA 23012 wings an elliptical and a circular fuselage and ver.pdf

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1、REPORT No. 705WIND-TUNNEL INVESTIGATION OF EFFECT OF INTERFERENCE ONLATERAL-STABILITY CEARACTEIUSTICS OF FOUR NACA 23012WINGS, AN ELLIPTICAL AND A CIRCULAR FUSELAGEAND VERTICAL FINSBy RUFUS O. HOUSE and ARTHUE R. WALLACESUMMARY.4 wind-tunnel irmesiigation of the effect of wing-juselage interference

2、on lateral-stability characteristicsuws made in the NACA 7- by IiI-foot wind tunnel. FouriVACA 2301$ wings were teetedin combinati wiih twofuselages and iwo $n8, representing high-wing, low-wing,and midwing monopkne. !ihe fwelage are of eirculizrand elliptical cross 8ecti0n. The wing8 hare roundedti

3、p8 and, in plan form, one is rectangular and the otherthree are tapered 3:1 with rarkrw amount8of wxep.The rate of chungein the coejicients of rolling moment,yam”ng moment, and luteralforce with angle of yaw L+gicen in a form to 8how the increment tamed by wing-fuselage interference for the model wi

4、th no fin and theeffect of mung-jwelage interference on effecticene88.Re a carreqmndingdecreasein effectire dihedralum8obtainedfor the Ia-”ngcombination. With jlap8 neutral the maximum inter-ference efleci UXL8 forthe other three wings, which are tapered 3:1, the ordi-nates of the ellipses have been

5、 expanded in proportionto the taper of the individual Ieading or trailing edges.The NACA 23012 prdle is maintained to the ends ofthe wings and, in elevation, the maximum upperarfaceordinatea are in one plane. The srea of the rectangularwing is 3.917 square feet and the aspect ratio is 6.383;for the

6、tapered wings, the area is 4.101 square feet andthe aspect ratio is 6.097. The sweep angk of thetapered wings are 4.75, 4.75, and 14.00. Thewings were set at 0 incidence to the fusehige centerline in all positions.31Provided by IHSNot for ResaleNo reproduction or networking permitted without license

7、 from IHS-,-,-32 REPORT NO. 705NATIONAL ADVISORY COMMITTEE FOR AERONAUTICSA and B ore quadrontsof similare/lsas,.=!. -.:. .q.15”1 ,L - _3“ 1- I.- Sp/iffbp/5” I3“ .-. . ._ ,vT1FIOVBEl.Plan views and denatfons of the NAC4 2$312wfngs.The two fuselages used are shown in figures 2 and 3and the dimensions

8、 are given ti- tabIe I. The maxi-mum cross-sectiomd area of the two fuseIages is thesame. The circular fusekge, which was used for thetestsreported in reference 5, was made from dimensionsobtained from reference 8.The h were made to the NACA 0009“section and,in phin form, are representative of h now

9、 .in use.The area of the f3n of the circular fuselage is 53,7 andthat of the eIIipticaI fuselage is 56.2 square inches.These areas are given to the center of the fuaehge.The aspect ratios of the fins of the circuk.r and of theeIIiptioaIfuselagm are 2,20 and 2.26, rwpecthdy. The/-I+-.-IiFIGURE2.Drawt

10、m?of NACA !23312wlrw In mmbfnatfon wItlI drcuk+r fuseh.gc amlfin of NAC 0)29section.Wingpositionfinches),“w-/P-b/2=30”.-WFIGUB8.Drawing of NACA ZXt12wfng fn comblrrationwithelllptlonf fudgeand M of NACA OOWwdlon.distance from the assumed position of tlm center ofgravity -of the model to the trailing

11、 edge of the fin is0.455 times the wing span.The split flaps were made of -inch steel plates tmc,for the flap-deflected condition, were attached to thowing at an angle of 60., The flaps have a uhord 20 _percent of the wing chord, are tapered with the wingchord, and extend over the inboard 60 percent

12、 of thespan. For the micking and the high-wing positions,the center section of the flap was cut away to aNowforthe fuselage. The gap between the flap and the fuselagewas seaIed for alI kte,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-T7rlrKll fQQ,

13、2./o/d(b)-8 -4 0 4 8 /r /6 20Angle of offock,d ,deg(b) TaIMM wiUfi A, 4,76”;OkQUk fusdME.Flounm4.Llrt, draK,nad pltddng-monmntooa7Montdor eomplatamodel w a high.wlugmonoplane,. IIProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1-1 q,$g ?,* A -bP I I

14、/1 I#.4 I I A I I I 1/ I IiAngk of cn%ck d ,deg”-(d) ReOhIWUIacwti A, m elliptkd ftuelagaFnmmm4.Conthmed.,.:,;,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-/.61 I I I I I I I I I I 1 , Ix- -uo%( !hlM* Wh; A-4.75*; elliPtfd fU (a) TcmIrIMwlnu: A+.7

15、6”: ohnlu fnselauo.FIOLW 4,OonaludecL FmuBrn 6,LKt,dcag, and pltihlng-moment oMEclents of oomplete modo as a low-wing monoplnne.C d!PLiLwIlmMftIKI, (e) TarmrfK4WIUI!;A, 4.73;dlip:icnl rm!ag%!, : IProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1- 8=,

16、deg ga .-.2$,deg I Y I $I :1 “1 ! ,=151111111-.3$TI I I I I I I I /1 I I I I.-a1111111 l -_ t*Recfangulor wing-.00.0010(-.ik4c.; - - i +I I -a- -.= -d, -%)3:/ ?ner;A, 14.00-.001mlo ,.-+ -: +- = :, - -3:1 fuper;A, 4.75-.00/ “.001uL (i, - L -I -3:Ifaper;A, -4.75 k70011 1-/0-5Angk fatfack,;,deg 10 IsGU

17、RE 12.-VerIation of C,It with angleof Waek. NACA 23012w!. (Detsfrom reference 8.)Values of C,t and CYfor the complete model may boobtained in a similar manner.DISCUSSIONThe application of theory to the problem of the in-fluence of wing-fudage interference on latei%l-stability characteristics is diff

18、icult because of the com-plex flow involved. Several components of the flowand their probable effects will, however, be consideredin a qwditative manner,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INVESTIGATION OF EI?FECI OF INTERFERENCE ON LATER

19、AL-STABIIJTT CIL4WLCTERISTICS 43 ,-“t-t-i-t:=0I I .1-.004 -71Rectangular whg.Go4 1 I fo - . + AF -7004 1 I3;/ faper; A 14.00= 004J ,%;o / - T- T t04 L. . I-. - *+?iwf - ,.004 T0-.004 I3:1 iaper;A-4.75iI-lo -5t%gle%faffack,Z ,deg10 r5?lawm M.-lartatton d Cpt wfth angle of attack. AC.4 w. Sfrom refere

20、nw 3.)Change in span load distribution is beIieved to be animportant factor in wing-fuseIage interference. Boththe unqymmetricaI flow resulting when the fuseIage isyawed rind the flow over and under the fuselage con-tribute to a change in the Ioad distribution aong thespan of the wing.A region of in

21、creased pressure exists on the side ofthe fuselage toward the wind and a region of decreasedpressure exists on the down-wind side. The flow aboutthe wing wdl be modified depending on the position ofthe wing on the fuseIage. With the wing in the highposition, there wiII be an addition of lift on the

22、sidetoward the wind and a corresponding reduction in lifton the down-wind side. Thus a rolling moment shouldw.aultthat tenda to raise the leading wing tip. It isX seen that, with the wing in the Iow position, thehangein Ioading would be such as to produce a roIIingmoment in the direction opposite to

23、 that obtained withthe wing in the high position. With the wing in themidposition, this effect should be a minimum.An additional change in span Ioading is brought aboutby Iocal changes in wing rmgIeof attack caused by theRow over and under the yawed fuselage. With thewing in the high position, the a

24、ngle of attack of aportion of the wing near the fuseIage is increased on theside toard the wind and is decreased on the down-wind side. h opposite change in angle of attackprevaik with the wing in the Iovrposition; with the wingm the midposition, the change should be smaII.Thus, when the model is ya

25、wed, the two interferencefactors considered should give an increment of rollingmoment tending to raise the leading wing tip of a high-wing monopIane and to lower the Ieading wing tip of alow-wing monoplane. kmgitudinal position of thewing on the fuseIage should be an important factor inthe change in

26、 span loading just discussed because of thefuselage load distribution.The presence of the -wing exerts an appreciableinfluence on the flow about the fuselage. With eitherthe high-wing or the low-wing monoplane in yaw thewing acts as a motied end pIate, which should causean increase in lateral force.

27、 The presence of the wingshould also change the fuselage load distribution- Thusthe longitudinal position of the AU on the fuselageshould affect both the magnitude and the center ofpressure of the lateraI force and, consequently, themagnitude of the unstable yawing moment of the fuse-lage. The vorte

28、x field is apparently affected by inter-ference, which results in an induced lateraI flow at thetail. When the fuselage alone is yawed, vortices areshed at the top and the bottom of the fuselage aonwwhat like the tip vortices of a wing, the strength ofthese voitic= increasing with increase in Iatera

29、l force.U a wing is placed on the fuseIage in the high or the -Llow position, the lateral force should be incrertsedbecause of the end-plate effect and the vortices shouldincrease in strength. With the wing in the low position,however, the vortices shed from the bottom of thefuseIage are so affected

30、 by the wing that the iriducedIateral flow near the bottom of the fuseIage is greatIy,-reduced. SimilarIy, with the wing in the high position,the induced latemd flow is decreased at the top of thefuselage. These characteristics have been noted intiual observations of the flow by means of tufts-.L -.

31、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.44 REPORT NO. 70onHi,5I Iml I I I I /1I1 I L Io /,:/ ,_. /.Y L. t b-y- ? p.004 .w- J .4 1+_l“,0“ -.004 kAICY;I I-.-/0 105 AeOof otiock, ,deg /5+-l(b)Ealptlcalfuselage.frmcrm lr,0?m Yk e .I.m4 - 1“ -?-”

32、 “-c-+-_+%+ - L-, .=. - -rw L7;-41 1 1 k 1 1If?ecfunaularwire:r.51 ,T.mI I I I I I I I I I I I !.(W4 1 I t t t I II / .+ a71Jmf T T II I I 13:l”fai?r,A. /4.oo; r.u” I I I I. . .W%; ,.004I I I I1 I I 1. I Iti3:/taper;A, 4.75; 17,0- :004 t_tiii”iiiiiii HI , , , , , , 1 t k 1 1 1 1.iwf I I I I I I I I

33、I I Ir.+ * 1.I I - .-r- .70041 1 I-.oogo- .: I I I I I I I I t II Io 5 10 “15Angleofaffadq d ,deg(n)Ckmlmfoeelege.CDatafmreetengukr wing wesobtalnedkomreference 6.)FIGUEE 16.Increment of C/t tie h whnseIw interference. NACM 2M12Q7fnga.was computed for a wing of the same aspect ratio asthe fin from t

34、he data given in figure 4 of reference 9.The increase in Cr$ produced by the fin is about 10percent greater than this computed value. The changein C/$ with angle of attack is of the order expectedfrom the change in C=* produced by the fin and theverticaI-taiI position; the change in C.$ produced by.

35、002, I I I I I I I I I I I I 1t I.i/Ag/oLaAn 1High MI Law c$deg.ml Ovx+uD501 # ) I I tm- Ti.,Lw w;-. r n.-.00 I t IRecfang-,-, V,“mill I t ItI 1 fM=0-; .-1 I I t I .-.00 I II t t tt Rectangular wing;r.5: 17”-+ ItI I I I 1 t 1 I I I I I700/ I I I 1 I I3:1 faper; A , 14.WO;I,0”.0(7/o4I* L 1:Ooi t3-Ifo

36、per;A, 4.75; J7, O“.0011md - / a71700 t I I(a 3:ffaper;A,-4.75; l?,”1-10 -5 0 5 0 15r,o” t-.002.m/AIC+o700/ I I 1 1 I3:/+Oper;A ,4. 75”; r, 0 -(b) i,%0I I 1-5 0 5 /0 15Ang/e ofoffd, d ,deg(b) EllfPtloal fnselasa.FIGUItEld.-ConoIuded.0080 v.004i-.004.Lw4A. .- ./01Angleofutiuck, ,deg .-(b) Elllptbml f

37、nseke.FIWJEE17.-Comlrrdsd.Iage. At 0 angle of attack, the values of (?=$ andC.$ for the elliptical fuselage are nearIy twice as largeas those for the circdar fuselage. At this angle ofattack, CY$ is a minimum and Cm; is a maximum. Achange in angIe of attack in either direction from 0 isaccompanied b

38、y a marked increase in CYJ and by anrappreciable reduction in unstabIe yawing moment,indicating a movement of the center of pressure towardfhe rear.Angleofa#uck,d ,deg(a Cfr A, 4.75-;,01-!0 5 .AngleOofalfaOk,Z ,o%g lo !5a) CfrcularfuseIage. fDafa forrectangdar wing obtafnedfrom reference.)FIUGREl A

39、, 4.75; r, 0 y I-lo -5 10Ang/eOofoffock, ,deg/5(b) EIIfPtfcaI fnseFIG- IS.-Concluded.I t I.001r.00r ;14CI;-.002 +.00/3.7 faper:A, 4.75”!r. 0) HIfptfc81 fuselage.FIGUIU?,19.Ef7ectof wfng-fusekge fnterferenea on CYt due ba fm. h-ACA 28012-.causes a deorease in CrJ. Increasing the sweepbackincreases th

40、e effect of CL on cl. It should be notedthat, for the tapered wing with the leading edge str decreas-ing Cmtiincreas the stabIe yawing moment of thewings.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-48 REPORT NO. 705NATIONAL ADVISORY COMMITTEE FOR

41、 AERONAUTICSI I I.002 Wing Ioca+icmHighMid Low T,0“-,ooi.001- v- - -1 1-.09/.00!0,AZCI;7WI.00/,7001.00/01 I 1 i 13:f faper;A ;4.75;I,0” -.og+o 1-5 0.5 10 /5Angleofoffack,d ,deg(W CfrcnIarhelage- (Dnts torrcctfmmh wing m.sobtohed from refomnm 5.)FIGURE 19.-EfM of wing-fuselage interference on Cf$ due

42、 to fin. NACA 22J312w-WING AND FUSELAGEInorement of Cfi caused by wing-fuselage inter-ference,Values of AICequivalent to an increase indihedd from 2 to 6 were obtained with the wing inthe high position; with the wing in the low position, theeffective dihedral was decreased from 2 to 6. Thechmge in e

43、ffective dihedraI with the wing ig the mid-position was 1 or less. (See . 14.) These resultsare in agreement with the theory previously given.The reversal of the curves of AIC/$ for the low-wingmonopkme is of interest. The v no filIets wereused on”the present models,Of the other variables, fuselage

44、shape for a givencross-sectional area appeared to have an appreciableeffect on the value of Alct, much larger vrdues beingobtained with the elliptical than with the circular fuse-lage (fig. 14). Inmuch as the yawed elliptical fuse-lage develops more lateral force than the yawed circu-hr fuselage, th

45、e larger values of AIC obtained withLheelIitical fusdage are consistent with the theory ofthe chae in span loading.,.The eflects of wing taper and of longitudinal posi-Lion of the wing on the fuselage are interdependentbecauseeach of the wings was tested at a differentongitudimd position in order to

46、 ocatc the meanerodynamic center of the wing at the assumed centerof gravity of the model for all combinations. In general,Lhe absolute values of AIC/t increased as the sweepWM increased. This change is believed to be largelyFmuEE 2).-Vortex field caused by fnterferanw for the lo for the mid-wing co

47、mbination, the values are small. (See . 15.)When the high-wing or the low-wing monoplane isyawed, the wing acts as a partial end plate, whichtends to increase the Iateral force on the fuselage.Larger values of AICY$ were obtained with theelliptical than with the circuhw fuselage.The effect of dihedd

48、 on A,C=$ was hugely de-pendent on wing position (fig. 15 (a). h increasein dihedral was accompanied by a corresponding in-crease in AICyt for the high-wing monopkme; forthe low-wing mcdeI, the opposite effect was noticed;and, for the midwing model, the effect of dihedral wasinconsistent.Flap deflection acted to increase AICT* with thewing in both the mid position and