NASA NACA-RM-L52A29-1952 Wind-tunnel investigation of the aerodynamic characteristics in pitch of wing-fuselage combinations at high subsonic speeds aspect-ratio series《在高亚音速展弦比系列下.pdf

《NASA NACA-RM-L52A29-1952 Wind-tunnel investigation of the aerodynamic characteristics in pitch of wing-fuselage combinations at high subsonic speeds aspect-ratio series《在高亚音速展弦比系列下.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L52A29-1952 Wind-tunnel investigation of the aerodynamic characteristics in pitch of wing-fuselage combinations at high subsonic speeds aspect-ratio series《在高亚音速展弦比系列下.pdf（43页珍藏版）》请在麦多课文档分享上搜索。

1、RM L52A29RESE DUMWIND-TUNNEL INVESTIGATION OF THE AERODYNAMICCHARACTERISTICS IN PITCH OF WING-FUSELAGECOMBINATIONS AT HIGH SUBSONTC SPEEDSASPECT -RATIO SERIESBy Richard E. Kuhn and James “W. WigginsLangley Aeronautical LaboratoryLangley Field, Va. .-. “,%+Y-Provided by IHSNot for ResaleNo reproducti

2、on or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM.C NACARM L52A29 Illllll!llllllllllllllilllllllllll!JIJ43834 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.-UWIND-TUNNELRESEARCH MEMORANXMINVESTIGATION OF TEE AERODYNAMICCHARACTERISTICS IN PITCH OF WING-FUSELAGECOMBINATIONSAT

3、 HIGH SU2SONIC SPEZDSASPECT-RATIO SERIESBy Richard E. Kuhn and James w. wigginsSIMMARYAn investigation was conducted in the Langley high-sp=ed 7- by 10-foot tunnel to determine the effect of aspect ratio on the aerodynamiccharacteristics in pitch of wing-fuselage combinations with 45 sweep-back at t

4、he quarter-chord line and 0.6 taper ratio at high subsonic4 speeds. Generally good agreement was obtatied between the theoreticalwing-fuselage and wing-alone lift-curve slopes and the expertientaldata, although the absolute magnitudes given by the wing-alone theory.were somewhat low. The experimenta

5、l wing-fuselage aerodynamic-centervariation with aspect ratio agreed fairly well with wing-fuselage theoryat a low.Mach number for which the compar-isonwas made. The resultsshowed little variation of the aerodynamic center with Mach number upto the force-breakMach number. Above this point all wings

6、exhibited arapid rearward movement of the aerodynamic center. The drag-rise Machnumber tended to increase slightly with increase in aspect ratio. Belowdrag rise, the zero-lift drag (wing plus wing-fuselage interference) ofall three wings was approximately the same. The drag due to lift gener-ally de

7、creased with an increase in aspect ratio but generally showedonly small variations with Mach number. Increases in aspect ratio pro-duced an increase in maximum lift-drag ratio. Above the drag rise Machnumber, all wings exhibited a marked decrease in maximum lift-drag ratio.INTRODUCTION4A systematic

8、research program is being carried out in the Langleyhigh-speed 7- by 10-foot wtid tunnel to determine the aerodynamic char-.acteristics of various arrangements of the component parts of research-type airplane models, including some complete model configurations. .Provided by IHSNot for ResaleNo repr

9、oduction or networking permitted without license from IHS-,-,-2 NACA RM L52A29Results are being obtained on characteristics in pitch, yaw, and duringsteady rolling up to a Mach number of aboutO.95. The_models are mountedon a sting-type support system. Reynolds numbers range between 1,500,000”and 6,0

10、00,000, depending on the wing plan foxmm and test Mach numbers.The wing plan forms are simil= in genralj to the plan forms inves.tigated at lower Reynolds numbers”duringa previous research programwhich utilized the transonic-bump technique for obtaining results attransonic speeds. Some of the result

11、s obtained from the transonic-bumpprogram have been summarized in reference 1.” Some higher-scale tests ofsimilar or related wing plan forms have been performed in other windtunnels (references 2 to 4). A comparison of aerodynainlccharacteristicsin pitch as obtained by different test techniques has

12、been rerted inreference 5.The present paper presents results which show the effect of aspectratio on the pitch characteristics of wings having a sweep angle of 45,a taper ratio of 0.6,and an NAM 65Aoo6 airfoil secton in combination “-with a fuselage. In order to exdite the issuance of the results, o

13、nlya limited analysis has been made, although comparisons of some of themore significant characteristicswith available theory are presented.-COEFFICIENTSAND SYMBOLS.The symbols used in the present .r ?redefined the followglist. All forces and moments are presented relative to the quarter chordof the

14、 mean aerodynamic chord.CL lift coefficient (Lift/qS)CD drag coefficient (Drag/qS) % pitching-moment coefficient (Pitchingmoment/qSE)q dynamic pressure, pounds per square foot (Pv2/2)s wingF meanarea, square feetaerodynamic chord (M.A.C.), feet ( b2c0$c local wing chord, feet -. .-().cave average wi

15、ng chord, feet Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARM L52A29 3b.Pb vMRabALKCLA9%. A(a/bcL)YYTfiAt/4cZc/cLcaveSubscripts:Fa71 W?sspan, feetair density, slugs per cubic footfree-stream velocity, feet per secondMach numberReynolds number

16、 of wing based on Fangle of attack, de”greeslocal angle-of-attack change due to distortion of wings,degreeslift increment due to distortion of wings, poundscorrection factor for c due to wing distortionlift-curve slope (acLf2A3-.The wing-fuselage combinations investigatedare shown in figure 1.All wi

17、ngs had sm NACA 65AO06 airfoil section parall,elto the fuselagecenter line. A common alumimun fuselagewas used, the ordinates of w”chare shown in table I. The aspect-ratio-2 and -6 wings were constructedof solid aluminum alloy. The aspect-ratio-4wing was of comsite con-struction, consisting of a ste

18、el core and a bismuth-tin covering to givethe section contour.The three wings used in this investigationrepresent only a part ofthe family of wings being studied in a more extensive program; therefore,a simplified system for designating the wings (similarto that used inreference 4) is being utilized

19、 for this program. For in this case, the design-lift-coefficientbzero and the thiclmess is 6 percent of the chord.The models were tested on the sttig-type sup?ort_systemshown infigure 2. With this support system the model can be r=motely operatedthrough a 28 angle range. The internally mounted elect

20、rical straip-gage balance used is shown installed in the fuselage in figure 3.The teststunnel throughTESTS AND CORRECTIONSwere conducted in the Langley high-speed 7- by 10-foota Mach number range from approximately 0.40 to 0.95.The size of the models used caused the tunnel to choke at correctedMachn

21、unbers of from 0.95 to 0.96, depending on the wing being tested. Theblocking correctionswhich were applied were determined by the velocity-ratio method of reference 6 which utilizes experimental pressures mess-_ured at the tunnel wall opwsite the model. The corrections determinedin this manner were

22、checked by the theoretical method of reference 7and, in general, good agreement was observedjalthougl.abovea Machnumber of O.$X?the values obtained in reference 7 were somewhat higher.The jet-boundary correctionswhich were applied to the lift and dragwere calculated by the method of reference 8. The

23、 correctionto pitchingmoment was considered negligible.,.-. No tare correctionswere obtained; however the results of reference 9indicate that for a tailless sting-mountedmodel, similar to the models b“reported herein, the tare corrections to lift and pitching moment were ,.negligible. The drag data

24、have been corrected to correspnd to a Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L52A29 5pressure at the base of the fuselage equal to free-stream static pres-.sure. For this correction, the base pressure was determined by measuring -the

25、 pressure at a point inside the fuselage about 9 inches forward ofthe base. This correction, which was added to the measured drag coeffi-8cient, amounted to a drag coefficient ticrement that increased frm avalue of 0.001 to 0.004 for the wing-fuselage configuration and from0.001 to 0.002 for the fus

26、elage alone, as the Mach number increased fromo.4tlo 0.95.The angle of attack has been corrected for the deflection of thesting-support system under load.The test wings were lmown to deflect under load; accordingly, in aneffort to correct the measured data to corresnd to the rigid case,correction fa

27、ctors for the effect of this aeroelastic distortion weredetermined. Two types of distortion-were considered: (1) The twist ofthe Wingabout its elastic sxis, and (2) the spanwise change in angleof attack due to bending of the wing under load. Both types of distor-tion increased markedly with increasi

28、ng aspect ratio but with 45 sweepthe change in angle of attack due to bending is the predominant factor.A preliminary deflection analysis showed practically no deflection ofthe aspect-ratio-2 wing.* The correction factors for the effects of aeroelastic distortionwere determined from static loadings

29、of the wings. ti an attempt to. approximate this distortion, an elliptical load distribution was shnu-lated by applying loads at four spanwise mints along the quarter-chordline of each wing. The change b angle of attack fi )=however, above aMach number of 0.91the aerodynamic center moves rapidly rea

30、rward forall three wings, as would be expected.The theoretical wing-fuselage aerodynamic-center locations, as pre-. dieted by reference 12, are in fairly good agreement with the experi-mental results (fig. 16). The small discrepancies shown may be due inpart to the fact that the effect of the presen

31、ce of the fuselage on theatheoretical span-load distributions which are obtained from reference 10,is not considered in the theory of reference 12. It will also be notedthat at the highest aspect ratios, the wing-alone theory of reference 10,Provided by IHSNot for ResaleNo reproduction or networking

32、 permitted without license from IHS-,-,-8 NACARML52A29which reference 12 uses as a basis, is in -or agreement with the wing.alone data of reference 4 (fig. 16, M = O.).At the higher lift coefficients the pitching-momefitcurves of theaspect.ratio and 6 wings (figs. 8 and 9) iridicatedestabilizingbrea

33、ks, with the break for the aspect-rabio-6 wing being more severe andoccurring at a lower lift coefficient than that of the aspect-ratio-4wing. The wing of aspect ratio 2 exhibits adefinitestabilizingtrend at the higher lift coefficients. These effects are in agreementwithsomethe correlation presente

34、d in reference 13.Drag CharacteristicsDrag at zero lift.-At Mach numbers below the force break there isvariation in minimum drag coefficient between the three wing-fuselage configurations (fig. 14). Inasmuch as a common fuselage”wasused and the wing area of the three wings varied with aspect ratio,

35、theincrement of drag coefficient attributable directly to the fuselage alsovaried as shown in figure 17. To give a better comparison or minimumdrag coefficients,the wing plus wing-fuselage interference drag isplotted in figure 18. The wing plus wing-fuselage interferencedragwas obtained by subtracti

36、ng the fuselage-alone drag (fig. 17) from thewing-fuselage drag of figure 14. The slight difference shown can beattributed partly to interferenceeffects and partly to the relativeaccuracy of the results. The drag-rise Mach number tended to increaseslightlywith increase in aspect ratio (fig. 18); thi

37、s same effect hasbeen noted in reference 14.Drag due to lift-At”thelower lift coefficients the drag due tolift decreased with an increase in aspect ratio (fig. 19) as would beexpected. The drag due to lift of the aspect-ratio-2 and 4 wings isnot affected by Mach number (fig. 20), but the drag coeffi

38、cient of theaspect-ratio-6wing decreased at the higher Mach numbers. The data ofreference 1 show the same trend for this wing. The reason for thisreduction is not understood but it can possibly be attributed to thewashout of the wing due to distortion. Figure 20 also presents a com-parison with the

39、theoretical values (givenapproximately by CL2/fl)for the condition of theresultant force normal to the local relativewind. It wilJ be noted that the experimental drag due to lift breaksaway from the theoretical curve at a low lift coefficient, indicatingthe possibility of an early loss of leading-ed

40、ge suction because ofleading-edge separation. -“s-.a71.- .aProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2(2 NACA RM L52A23 9Lift-Drag Ratios.The results shown in figure 14 indicate that Mach nmber has littleeffect on the maximum lift-drag ratio up

41、 to the drag-rise Mach number,bbut above this point a rapid decrease occurs. The maximum lift-dragratios increase with increasing asct ratio, as expected. However, dueto lower minimum drag, the aspect-ratio-2 wing-fuselage combination hasa msxinnnnlift-drag ratio about as high as that of the aspect-

42、ratio-kwing-fuselage combination. It will also be noted (fig. 21) that athigh lift coefficients a very substantial gain in lift-drag ratio isobtained with increasing aspect ratio at the higher Mach numbers. Reduc-tion in aspect ratio is seen to reduce the lift coefficients at whichthe maximum lift-d

43、rag ratio occurs (fig. 21). bCONCLUSIONSThe results of the tivestigation of the effect of aspect ratio andMach number on the aerodynamic characteristics in pitch of 45 swept-back wings with 0.6 taper ratio and an NACA 65AO06 airfoil section indi-cated the following conclusions:. 1. The variation wit

44、h Mach number of the lift-curve slope, as pre-dicted by wing-alone theory, was in good qualitative agreement with theexperimental results, although the absolute magnitudes were somewhatlow. The theoretical lift-curve-slope variation with aspect ratio, aspredicted by wing-fuselage theory, was in good

45、 agreement with experi-ment at a low Mach number for which the comparison was made.2. The theoretical wing-fuselage aerodynamic-center variation withaspect ratio, as predicted by wing-fuselage theory, shows fair agree-ment with the experimental results. The experimental aerodynamic centershowed litt

46、le variation with Mach number up to the force break; however,above the force-breakMach nunbers all wing-fuselage combinationsexhibited rapid rearward movements of the aerodynamic center.3. The zero-liftdrag (wing plus wing-fuselage interference) of allthree wings was approximately the same at Mach n

47、umbers below the dragrise. The drag-rise Machnumber tended to increase slightly with.increase in aspect ratio.4. The drag due to lift generally decreased with an increase in.aspect ratio and showed only small variations with Mach number withinthe range of these tests.a71Provided by IHSNot for Resale

48、No reproduction or networking permitted without license from IHS-,-,-10 NACA RM L52A295. Theratio. Allratio abovemaximum lift-dragwings exhibited aratio increasedwith increase in aspectmarked decrease in maximum lift-drag *the drag-rise Mach number. . A.“Iangley Aeronautical Laboratory National Advisory Committee for AeronauticsLangley Field, Va. ,-.-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L52A29REFERENCES111. Polhamus, Edward C.: Summary of Results Obtained by Transonic-BumpMethod on Effects of Plan Form