NASA NACA-TN-2587-1951 Influence of wing and fuselage on the vertical-tail contribution to the low-speed rolling derivatives of midwing airplane models with 45 degree sweptback sur.pdf

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1、 -wLo03CD.NATIONAL ADVISORY COMMITTEEFOR AERONAUTICSTECHNICAL NOTE 2587INFLUENCE OF WING AND FUSELAGE ON THEVERTICAL-TAIL CONTRIBUTION TO THE LOW-SPEEDROLLING DERIVATIVES OF MIDWING AIRPLANEMODELS WITH 45 SWEPTBACK SURFACESBy Walter D. WolhartLangley Aeronautical LaboratoryLangley Field, Va.Washingt

2、on. . . . . . . . .,-. . - .- - . -Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-IlB.t.TECHLIBRARYi.AFB,NM1lllMMlwlllilMllNATIONAL ADVISORY CCMMITWE FOR AeronauticOOLIkihL5TlIfJCf3L NOTE2587INFLUENCE OF lQITGAND FUSELAGE ON TEEVERTICAL-TAIL CONTRIW

3、TIOIVTO THE LOW-SPEEDROLILCNGDERIVATIVES OF MIDWING AIRPLANEMODELS WITH 45 SWEFIJ31CKSURFACESBy Walter xD.WolhartSUMMARYAn investigationwas made to detezmine the influence of the wtigand fuselage on the vertical-tail contribution to the low-speed rollingderivatives of midwing airplane models with 45

4、 sweptback surfaces.The results show that the vertical-tail contribution to the rollingderivatives of midwing or near-midwing configurations can be calculatedwith good accuracy throughout the angle-of-attackrange by using availableprocedureswhen corrections have been made for the effects of fuselage

5、 “and wing sidewash at the tail due to roll.The mutual wing-fuselage interference increments of the wing-fuselageconfigurations investigatedshowed no consistent effect of fuselagelength. The incrementswere usually rather small and did not vary appreci-ably with angle of attack except that the increm

6、ent in yawing moment dueto rolling became quite large at angles of attack above 160.The contribution of the fuselage alone to the rolling derivativeswas small in comparisonwith the effects of angle of attack for theother components of the models investigated.! INTRODUCTIONRecent advances in the unde

7、rstanding of the principles of high-speedflight have led to significant changes in the dsign of the prihcipalcomponents of airplanes such as the incorporationof large amounts ofsweep of the wing and tail surfaces. Although the effects of changes inwing design on the stability characteristicshave bee

8、n extensively investi-gated, there is little information available on the influence of changesin the other components of the ame. In order to provide such infer- ,mation, the Langley stability tunnel is conducting a series of investigationst - . . . . - . . - - .-. .- Provided by IHSNot for ResaleNo

9、 reproduction or networking permitted without license from IHS-,-,-2 NACATN 2587Qwith a model having various interchangeablecomponents. The effects onthe low-speed static lateral stability characteristicsof variations inhorizontal-tailsize and location, and of vertical-tail size and length.and of fu

10、selage shape and length are presented in references 1 and 2,respectively. The effects of variations in vertical-tail size and lengthand of fuselage length on the yawing stability characteristicsare pre-sented in reference 3.As part of this general investigation,the influence of the wingand fuselage

11、on the vertical-tail contribution to rolling derivativeshas been detemninedby the method of interference increments (refer-ence 4), and the results are presented herein. These results are usedto check the validity of present methods of estimating the vertical-tailcontribution to the rolling derivati

12、ves as well as to derive an empiricalrelation for estimating the fuselage sidewash due to roll.The data presented heretificients of forces and momentsof axes with the origin at theSYMBOE3are in the form of standard NACA coef-which are referred to the stability stemquarter-chordpoint of the wing mean

13、 aero-%Cndynamic chord. The positive directions of forces, moments, and angulardisplacementsare shown in figure 1. The coefficientsand symbols are (defined as follows:()a71CL lift coefficient qswcJJ()Ddrag coefficient q%Cy()Ylateral-force coefficient q%.c(%JLtrolling-moment coefficient qspitching-om

14、ent coefficient P used with subscripts 1 to 5 to denote thevarious vertical tails (see fig2)F fuselage; used with subscripts 1 to 3 *O denote the variousfuselages (see fig. 3)APPARATUS AND TESTSThe tests of the present invesfigation were made in the 6-foot-diameter rolling-flow test section of the L

15、angley stability tunnel. Thissection is equipped with a motor-driven rotor which imparts a twist tothe air stresm so that a model mounted rigidly in the tunnel is in afield of flow similar to that which exists about an airplane in rollingflight (reference5). Forces and moments on the model were obta

16、inedwiththe mode1 mounted on a single strut support which was in turn connectedto a conventional six-componentbalance systern. 8All components of the model used in this investigationwere con-structed of mahogany. ,Sketchesof the components of the models are pre-sented as figures 2, 3, and 4. The var

17、ious vertical tails and fuselagesare referred to hereinafter by the symbol and number assigned to them inr figures 2 and 3. All vertical tails bad 45 sweepback of the quarter-chord- . . . - . . - . - -. .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-

18、,-6 NACA TN 2587line, taper ratio of 0.6, and NACA 65AO08 profiles in,planes paralle1 tothe fuselage center line. The ratios of tail area to wing area were *chosen to cover a range representative of that used for current high-speed airplane configurations. The tails were mounted so that the rootchor

19、d coincidedwith the fuselage center line and the tail length wasalways a constant percent of the fuselage length (:=00+ ethree fuselages (finenessratios of 5.00, 6.67, and 10.00), havingcircular-arcprofiles and circular cross sections, are shown in figure 3.The coordinates of the fuselages are given

20、 in table I.,The wing had however, .%Pmeaaured and calculate (referene7) values of C% axe in poor agree-,ment, particularly at angles of attack above about “. The breaks inthe curves of the rolling derivatives are partly attributed to flowseparation from the wing. I? is expected that, for wings with

21、 highlypolished surfaces, an increase in Reynolds number would delay this flowseparation to scmewhat higher angles of attack. pointed out in reference 7, an indication of the Muiti.ng rangeover which flow does not separate from the wing can be obtained byCL2locating the initial break in the plot of

22、CD - against angle of. .-. - . -.- -. . - . - - -. .-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12attack. A plot of this- a break ti the curve atincrement forapproxhate lyNACA TN2587.the wing alone in figure 12 shows4 gle “ofattack. Inspection .

23、of figures 8 to 11 for wtig-on configurations shows break in the curvesof the rolling derivatives at approximately the same angle of attack.A compsrtion of wing-off and wing-on data of figures 8 to 11 showsa decrease in the vertical-tail contribution to the rolling derivativeswhen the wing is added

24、to the fuselage-tail configuration. This changecan be attributed to sidewash in the region of the vertical tail causedby symmetrical span loadtig on the wing due to roll (reference6).With the exception of vertical tail V5, the vertical-tail contri-bution to CZp was ather small compared to the wing c

25、ontributionbecause of the short distance from the centertail to the axis of roll.CyPareforThe fuselage alone generay contributesand C% throughout the angle-of-attacksmall, h.Before going into a discussion of the various interference increments,it should be pointed out that interfen?nce incrementsusu

26、ally are assumedto apply to airplane configurationswhich are scnmwhat similar in designto the model used in evaluating the increments. One of the factors whichaffects the magnitude of the interference “incrementsis the height of thewing relative to the fuselage center line (reference9). Since, for t

27、hepresent investigation,the wing was located on the fuselage center line,interference increments caused by the wing-fuselage combination are con-sidered applicable only to midwing or near-midwing configurations. Thetiterference increments caused by the fuselage are expected to be ltiitedto similar f

28、uselage-tail configurationswith respect to fuselage shapeand tail location.Wing-fuselage interfence.- The interference increments of the wing-fuselage combination, Al increments,are presented in figure 13 withan average curve faired through-the test points of the three wing-fuselage ,combinations ti

29、vestited. The Al increments of C and C% %=erather small and do-not vary appreciabwith angle of attack except that _ .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 2587 - / ulcnp becomesquite large at angles of attack above 160. terfenceof t

30、he wing-f%selage combinationsapars to have the greatest effecton CZ for wgles of attack above about 8 and tends to decrease thePdsmping-in-rollof the combination: There was no consistent effect of .fuselage length for the range of fuselage sizes tivestQated.Fuselage interferenceon tail, wing-off.- R

31、eference 6 has impliedthat fuselage sidewash at the vertical tail may influence considerablythe tail contributionto the rolling derivatives at high angles f attack.The increments due to fuselage interference on the tfil, A3 increments,are believed to be mainly sidewash effects resulting from vortice

32、s associ-ated with lateral forces which develop on the fuselage due to roll. TheA3C% increment occmring at a = 0 for some of the fuselage-tail/configurations (differencebetween me=ured and calculatedvalues of()c% v )shown in fig. 17 was subtracted from the data since the sym-metrical fuselages teste

33、d are assumed to contribute.nosidewash due toroll at 0 angleof attack. The A3C% incrementswere then convertedrinto average effective sidewash angles at lk in terns of the wing-tip helix angle as mentioned previously. These results are used toderive an empirical relation for estimating the fuselage s

34、idewash angles.In determiningthis empirical relation, considemtion is given tosome of the factors which might be expected to influence considerablythe fuselage sidewash angles such as fuselage size, /distance from Ev 4to the axis of roll, and the angle of attack of the fuselage. Thesefactors havez -

35、 (Zv Cosbeen combined in the parameteia- )Zvsina. The values of which equ=bdq obtained for all theh$fuselage-tailconfigurations investigatedare plotted against the param- 1eter 1941.TheoryRat io.,and Its Application toZ.f.a.M.M., Bd. 19,.Provided by IHSNot for ResaleNo reproduction or networking per

36、mitted without license from IHS-,-,-NACA TN 2587 21i-11. Michael, William H., Jr.: Analysisference on the Tail ContributionsTN 2332, 1951. ,of the Effects of Wing Inter-to the Rolltig Derivations. NACA12. Bird, John D., Lichtensteti, Jacob H.j and Jaquet, Byron M.: Inves-tigation of the Influence of

37、 Fuselage and Tail Surfaces on Low-Speed Static Stability and Rolling Charactertics of a Swept-WingModNACA* L7H15,.1947ii. . .“. ,.,i -. .- . . . . - ._ . .- - - - - - -.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 NACA TN 2587TABLE I.- FUSEL.2

38、.25.3.000.756Fuselage: F1 F2 F3Length, ft,. . . . . . . . . . . . . . . . 2.50 3.34 5.00Fineness ratio . . . . . . . . . . . . . . . . . 5.00 6.67 10.00Tail length, ZV, ft (all tails) . . . . . . . . l. 1.39 2.09. Tail-length ratio, ZV/, (all tails) . . . . 0.347 0.464 0.697r Vertical tail: V1Aspect

39、 ratio, Av . . . . . . 1.0Taper ratio . . . . . . . .o.6Quarter-chordsweep angle,Av,deg. . . . . . . . . . .45NACA airfoil section . . 65AO08Area, , sqft . . . . . . 0.169sjjft . . . . .o.408Mean chord - . . . . 0.417%fiPerpendicular distance fromfuselage center line toCv4 of vertical tail, zv,ft .

40、. . . . . . . . . . 0.ig2. -. . -v1.00.64565AO080.3380.5830.592V31000.64565AO080.5060.7100.725Vk2.00.64565AO080.3380.82+0.416V52.00.64565AO080.6751.159o.5j20.267 0.325 0.375 0.532=s=. . - - ._. .,.- -.-. -. . - . .Provided by IHSNot for ResaleNo reproduction or networking permitted without license f

41、rom IHS-,-,-24 NACA TN 2587.TABLE III.- CONFIGURMIIONSINVEXWIGATEDwingoff Wing onConfiguration(a)wW+F1W+ F1+V2W+ F1+V4W+F2W+ F2+V1W+ F2+V2W+ F2+”V3W+ F2+V4W+ F2+V5W+F3W+ F3+V2W+ F3+V4Configuration Figure Figure(a)6,7 I- -FlF1 + V2F1 + V48(a) 8(b). .F2F2 + V1F2 + V2F2 + V3g(b)9(a)F2 + V4F2 + V5 10(a)

42、 10(b)F3 .F3 + V2F3+V4U(b)n(a)%Wation (For details, see table II and figs. 2 to 4): wF fuselagev vertical tsil.,*. .- . .-_. _ ,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-II- . . - .-.LFigure 1.- System ofb “-+L “Y6-)” Y_ #YQ.ztaxes used. krrows

43、 indicate positiw dire;tion ofangles, forces, and momnts.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-IIk-8.7 (b)Vz.z -1.0; y2.48+6 Sqh.6z-1() V4. AV4=2.0;4=48.6w h%7/d 18.539 /L J1- 6.6+10.7-+(C)Va.AV-1.0;-72.9Qh.*52-II+8.7-!(e)v AV5 -2.0;6-9-7.2

44、E4Jn.Figwce 2.- -MiOIW Of W?.tiCd- tailS tested. A = 0.6; A = 45; pro-f m, 17ACA 65Ao08. All Umemlons are in inch9E., , ._ _ _ _Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-27,.I,E-g t=oI.Figure 3.-Dimensions of fuselages tested; profile ordinates

45、 in table I.dhensions are in inches. - - . .- - . .- - .- .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28. /7A =45bw =36.00 + Cf.w=Ii.25 CiwFigure 4.- Dtiensions and location6.75+jdimensions are.of wing andin inches.vertical“.-tails. All.,. - ._

46、-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1I. . ., /“._-”-.“#-r /-”Figure .- View of model, configuration W + F2 + V2, mounted in the6-foot-diameter rolling-flow test section of the ey stabilitytunnel.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-