NASA NACA-TN-3529-1955 The transonic characteristics of 36 symmetrical wings of varying taper aspect ratio and thickness as determined by the transonic-bump technique《根据跨音速撞击技术决定的变.pdf

《NASA NACA-TN-3529-1955 The transonic characteristics of 36 symmetrical wings of varying taper aspect ratio and thickness as determined by the transonic-bump technique《根据跨音速撞击技术决定的变.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TN-3529-1955 The transonic characteristics of 36 symmetrical wings of varying taper aspect ratio and thickness as determined by the transonic-bump technique《根据跨音速撞击技术决定的变.pdf（132页珍藏版）》请在麦多课文档分享上搜索。

1、Provided by IHSNot for Resale-,-,-TECH LIBRARYKAFB,NM,CNATIONALAndynamicADVISORY C FOR AERONAUTICS lllillllllllllllMllMMllclhb495!l!ECIINICALNOTE3529THE TRANSONICOF VARYINGCHARAC=STICS OF 36 smmmucfu wmwsTAPER, ASPECT RATIO, AND THICKNESS ASD BY THE TRANSONIC-BUMPTECHNIQUE=By Warren H. Nelson, Edwin

2、 C. Allen, andWalter J. I33, and 0.6), and 2 (taper ratios of 0.33, 0.5, and 0.72), andNACA63AOOX sectionswith thickness-to-chordratios of 8, 6, 4,and 2 percent.The results indicate that the greatest effect of taper on the lift-curve slope occurred for the wing having the highest aspect ratio andthe

3、 thinnest section. This effect, which was to increase the lift-curve slope with increasing taper ratio, diminished as the.aspectratiowas decreased and/or the thickness increased. Increasing taper ratiogenerally increasedthe over-all center-of-pressuretravel for all thewings in going from subsonic to

4、 supersonic speeds.INTRODUCTIONA comprehensiveinvestigationhas been initiated in the Ames 16-foothigh-speed wind tunnel to determine the transonic aerodynamic charac-teristics of wings having various aspect ratios, thichessesj cambers,and plan-form taper ratios. The experimentaldata on the,effects o

5、faspect ratio, thicbess, and camber have been reported in references 1and 2, smd analyses of the data using the transonic similarityruleshave been reported in reference 3. -%upersedes recently declassifiedNACA RMA3129by Warren H. Nelson,Edwin C. Allen, and Walter J. Kkumm, 1953. . .-. .-. . _._. _ _

6、 . . _ .Provided by IHSNot for Resale-,-,-2 NACA TN 3329.The purpose of this report is to present that part of the general. ,investigationinvolving the effect of plan-form taper.Three basic plan-form taper ratios, 0, 0.2, and 0.9, were investi-gated for wings having an aspect ratio of k and NACA 63A

7、-series sym-metrical sectionswith thickness-to-chordratios of 2, 4, 6, and 8 per-cent. These wings, in turn, were reduced in span by cutting off thetips to give aspect ratios of 3 and 2. The resulting plan-form taperratios for the wings having an aspect ratio of 2 were 0.33, O.z,and 0.72.NOTATION% d

8、rag coefficient,twice semispan dragc minimum drag coefficient* friction-dragcoefficient assumed equal to the minimum dragcoefficientat 0.6 Mach number;“-(L ,% minimum pressure-drag coefficient,assumed equal to ,( )Ckin %CL lift coefficient,twice semispan liftqs% pitching-momentcoefficient,referred t

9、o 0.21j 6,twice semispan pitching momentqszAMMLsvb2aspect ratio, mean Mach numberlocal Mach numbertotal wing area, Mce wing area of semispanmodel, sq ftvelocity, ft/sectwice span of semispanmodel, ftlocal wing chord, ft* b/2czdymean aerodynamic chord,fob/2C dyl,. Provided by IHSNot for Resale-,-,-NA

10、CA TN 3529 3!J. dynamic pressure, p, lb/sq ft2tF thickuess-to-chordratioY spanwise distance from plane of symmetry,fta angle of attack, degtip chordtaper ratio, root chordP air density, slugs/cu ftdCLx slope of lift curve measured at zero lift, per degdCm slope of pitching-moment curve measuredAPPAR

11、ATUS AND MODELSThe tests were conducted in the Ames 16-footat zero lifthigh-speed wind tun-nel, utilizing a transonicbump. A description of the bump may be._.found in reference 4. The aerodynamic forces and moments were measuredby means of a strain-gagebalance mounted inside the bump.A photograph of

12、 one of the wings of aspect ratio 4 mounted on thebump is shown in figure 1. The principal plan-form dimensions of thewings are shown in figure 2. The profiles used were NACA 63AOOX sectionswith thickness-to-chordratios of 2, 4, 6, and 8 percent (ref. 5). Threebasic wings of aspect ratio 4 having ta

13、per ratios of O, 0.2, and 0.5 andequal areas were constructedof steel for each thickness-to-chordratio.Aspect ratios of 3 and 2 were obtained by successivelycutting off thetips of the wings. The following aspect-ratioand taper-ratio combi-nations were included in the tests:Aspectratio Taper ratio4 0

14、 0.20 0.503 .14 a71 33 .602 a7133 .50 .72The mean aerodynamic chord of the wings changed with the taperratio; this change was such that.at any one aspect ratio the mean aero-dynamic chord of the wing with the highest taper ratio was about. . . - . - - . . Provided by IHSNot for ResaleNo reproduction

15、 or networking permitted without license from IHS4 NACA TN 352980 percent that of the wing with the lowest taper ratio. The wing tipswere faired by us one-half of the wing thickness at each chord station ,as a radius. JA fence 3/16 inch from the bump surfacewas used to prevent theflow through the ga

16、p between the wing and bump surface from affectingthe flow over the wing.overThe lift, drag, anda Mach number rangenumber range depended onTESTS AND PROCEDUREpitching-moment characteristicswere obtainedof 0.6 to 1.10. The correspondingReynoldsthe wing mean aerod “c chord. tbe extreme/range of Reynol

17、ds numbers being rom about 1:4 million to 2.0 million.The angle-of-attackrange, in general, was from -2 to the stall, or towhere the root bending stress became critical.Aldecreasingthe aspect ratio and/or increasingthe thiclmess ratiodecreased the magnitude and the variation with Mach number of the

18、lift-curve slope. These effects on rectangul= wings have been correlatedinreference 3 by use of the transonic similarityrules. The effects of- -Provided by IHSNot for Resale-,-,-NACA TN 3529 5taper on lift-curve slope were not nearly so pronounced as were thoseof aspect ratio and thickness. The larg

19、est effects of taper occurredfor the wings of aspect ratio 4 where the taper ratio varied from Oto 0.5. (It should be pointed out that, in these tests, as the aspectratio was reduced, the values of taper ratio increased and the range oftaper ratio decreased.) Moreover, the effects of taper were grea

20、testfor the thinnestwings and decreased as the thicknesswas increased.For the wings of aapect ratio 4, increasingthe taper ratio increasedthe lift-curve slope, this effect being greatest for taper ratiosbetween O and 0.2. The 8-percent-thickwings, having taper ratiosof 0.2 and 0.5 (aspect ratio 4),

21、had the type of variation of lift-curveslope with Mach number associatedwith shock stall. In the case of thewings of aspect ratios 2 and 3, there was little effect of changes intaper ratio on the lift-curve slope, except for thewing of aspectratio 3 having a thicbess ratio of 2 percent. For this win

22、g, therewas an increase in lift-curve slope with increasingtaper ratio through-out the Mach number range.( )The variation of minimum pressure drag coefficient CCDfwith Mach number for the wings is shown in figure 14. The effect ofchanges in thiclmess is ite apparent; decreasing the thickness ratioca

23、used a reduction in pressure drag, as would be expected. However, theeffects of taper and aspect ratio are not so apparent. In”general, therewere no large, significant,or consistent effects of taper ratio oraspect ratio on the minimum pressure drag, although some drag reductionwas obtained at the hi

24、gher Mach ntibers by decreasing the aspect ratiofrom 3 to 2. In the case of the 8-percent-thickwings, the effects oftaper appear large in light of a previous subsonic investigationreported in reference 6. In reference 6, the effect of taper on full-span wings having the same plan forms and sectionsw

25、as investigatedupto a Mach number of 0.94. Data from reference 6 have been included infigure 14(a) for comparison. In general, the data of reference 6 indi-cate there is no effect of taper ratio on the Mach number of dragdivergence except for the wings of aspect ratio 2 where the differenceamounts t

26、o 0.01 Mach number. This is in contrast to the data for thepresent investigationwhere for the tigs of aspect ratio 4 there is adifference of 0.025 in drag-divergenceMach number. The drag rise forthe wings of reference 6 is also more abrupt than that for the presentinvestigation. The reason for the l

27、ack of aeement between the datais not understood; however, to some extent it maybe the re6ult of thedifferences in the testing techniques.The variation of pitching-moment-curveslope with Mach number isshown in figure 17. In general, increasingthe taper ratio shiftedthe center of pressure toward the

28、leading edge and increased the over-allcenter-of-pressuretravel in going from subsonic to supersonic speeds._ _ - - - .-. . . -Provided by IHSNot for Resale-,-,-NACA TN 3529CONCLUDINGREMARKS.The eatest effect of taper ratio on the lift-curve slope occurredfor the wing having the highest aspect ratio

29、 snd the thinnest section;this effect, which was to increase the li”ft-curveslope with increasingtaper ratio, was considerablyreduced by decreasing aspect ratio and/orincreasingthickness. In general, increasing taper ratio had an adverseeffect on the pitching-moment coefficient, in that the over-all

30、 center-of-ressure travel was increased in going from subsonic to supersonicspeeds. In the case of the drag coefficient,the effect of taper ratiowas not consistent.Ames Aeronautical LaboratoryNational Advisory Committeefor AeronauticsMoffett Fieldj Calif., Sept. 29, 19531. Nelson, Warren H., and McD

31、evitt, John B.: The Transonic Character-istics of 22 Rec, Symmetrical Wing Models of Varng AspectRatio and!l!hichess. NACATN 3501, 1955.2. Nelson, Warren H., and Kiwmn, Walter J.: The Transonic Characteris-tics of 38 CaniberedRectangdar W3ngs of Vary Asct Ratio andThickness as Determined by the Tran

32、sonic-BumpTechnique.NACATN 3502, 1955.3. McBetitt, JohnB.: A Correlationby Means of the Transonic SimilarityRules of the ExperimentallyDetermined Characteristicsof a Seriesof Rectangular Wings. NACA Rep. 1.253,1955.4. Axelson, John A.,and Taylor, Robert A.: Preljhinary Investigationof the Transonic

33、Characteristicof an NACA Submerged Inlet.NACARMA5CZ13, 1950.5. Loftin, Laurence K., Jr.: Theoretical and Experimental ta for aNumber of NACA 6A-SeriesAirfoil Sections. NACA Rep. 903, 1948.6. MLen, 13dwinC.: Ekper3mental investigationof the IYfects of Pls,n-Form Taper on the Aerodynamic Characteristi

34、csof SymmetricalUnsweptWings of Vsrying Aspect Ratio.NACARMA53C19, 1953.Provided by IHSNot for Resale-,-,-NACA TN 3529 77. Allen, Edwin C.: Experimental Investigation of the Effects ofPlan-Form Taper on the Aerodynamic Characteristicsof Sym-metrical Unswept Wings of Varying Aspect Ratio.NACARM Aj3C1

35、9,1953.- . - _ _ . ._ _ . - .Provided by IHSNot for Resale-,-,-. Provided by IHSNot for Resale-,-,-NACA TN 3529 9A-17628.lFigure 1.- Photograph of the wing having an aspect ratio of 4and a taper ratio of 0.2 mounted on the bump._ . . . - . - Provided by IHSNot for Resale-,-,-10 NACA TN 3529Bump surf

36、oce1=0.694-1a0 .94 .98 L02 /.06 /.h8(a) t/c = 0.C8; A = OFigure 4*- The variation of Uft coefficient w3th engle of attack for the a6pect-ratio-4 wings. !2g-,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS/III(b) t/c = 0.08; h = 0.2Figure h. - Continued.P

37、rovided by IHSNot for Resale-,-,-I1.21.0 .8-b: .6,-Q“$ .4038 .2do-.2-.4-4 0 4 8 /2 /6 20 24 Angle of OttUCh, Q, deg vaof+ +:+!+ +for M of 0.60,70 .75 ,80 ,85 ,90 .94 .98 k02 /.06 1/0(c) t/c = 0.08; A = 0.5 sFigure 4.- Continued. EljProvided by IHSNot for ResaleNo reproduction or networking permitted

38、 without license from IHSI/.2-.4-4 0 4 8 /2 /6 20 24 Angle of attuck, a, deg vUof+ +:+i“-+qfor M of 0,60.70 ,75 ,80 ,85 .90 ,94 .98 L02 L06 LIO(d) t/c = 0.06; A = OFigure 4.- Continued.GProvided by IHSNot for Resale-,-,-/.2/.0-.4/.-4 /“M/ /l r Yr ibf M r$f o 0.60f “ dA. ,70. . -L1v/?,120AI I I I I I

39、 I I I I I I I I I I I I I I I I I./3.80nr7,94,98I no,.I/-Y I I I-1 1.06I I I I I I I I I I A I,foI I I I I I I I I I I I I I I I I I I I I I I I I 1 1 1 1 1 1 1 1 1 I I I-4 0 4 8 /2 /6 20 24 Angle of attack, a, deg 1Qoff +(+(+ +:for M of 0.60.70 .75 .80 .85 .80 .94 .98 L02 106 LIO(e) t/c = 0.(%; A

40、= 0.2Figure lt. - Continued. EProvided by IHSNot for Resale-,-,-1II1.21.0Q4 ,8-.2-.4-4 0 4 8 /2 /6 20 24 Ang/e of attack, a, deg vaof+ +(J + +(+for M of 0.60,70 .75 ,iO ,85 ,90 .94 .98 L02 L06 I!IO(f) t/c = 0.06; A = 0.5Figure 4.- Conttiued.sProvided by IHSNot for Resale-,-,-I I I I I I I.vI I 1 I 1

41、 1 1 1 1 1 1A -. .,/.2/. o .8 .6*u“$ ,4QQQ .2,-o-.2-.4-4 0 4 8 /.2 /6 20 24 Angle of attack, a, deg vaof+ + ;h+for M of 0.60.70 .75 ,80 .85 .80 .94 .88 102106 1/0(g) t/c = 0.04; A m c1Figure 4.- Continued.Provided by IHSNot for Resale-,-,-for M of 0.60.70 -.+5 .80 ,85 . A = 0.2Figure 4.- Continued.t

42、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSD3o1.4/.2/.0-,2-.44 O 4 8 /2 /6 20 24 Angle of attack. a. deaaof+i-+c)+:(! qfor M of 0.60,70 ,75 ,80 .85 .90 ,94 .98 I!02 1.06 LIO(i) t/c = 0.04; ?!= 0.5Figure 4.- f!ontinued.,Provided by IHSNot for Resale-,

43、-,-II(J) */c = O.; A= oFigure 4.- Contfnued.Provided by IHSNot for Resale-,-,-N/.21.0.8.6.4.20?2-,4-4 0 4 8 /2 /6 20 24 Angle of ottock, a, deg wOof+ :+: +h+for M of 0.60.70 .75 .80 .85 .90 .94 .j8 L02 /.06 /108(k) t/c E O.CQ; A = 0.2Figure 4.- Continued.Provided by IHSNot for Resale-,-,-/.2/. o-.2-

44、.4-4 0 4 8 /2 /6 20 24 Angle of oftock, a, tfeg T.-Uof+ : :+(!;:+for M of 0.60.70 .75 .80 .85 .80 .94 .98 i02 /.06 108(2) t/c = 0.(32; A= 0.5l?ie 4.- concluded.Provided by IHSNot for Resale-,-,-I/.2-.2-,4-4 0 4 8 12 16 20 24 Angle of attack, a, deg v.-aof +;:+ for M of 0.60.70 .75 .80 .85 .90 .94 .9

45、8 102 /.06 108(a) t/c = 0.08; A -0.14Figure . - The vamiatlon of lift coefficient with e of attack for the aspect-ratio-s wtngs.I,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSIII(/.2/.0-.4-4 0 4 8 /2 /6 20 24 Angle of attack, a, deg Uof+ +($! :+:+for M

46、 of 0,60,70 .75 ,80 .85 ,90 .94 .98 L02 L06 110(b t/C = o.oaJ A = 0.33Figure 5.- Continued.Provided by IHSNot for Resale-,-,-mm-4 0 4 8 /2 /6 20 24 Angle of ottack, a, deg for M of 0.60.70 ,75 .80 .85 .80 .94 ,88 102106 LIO.I(c) t/c = 0.08; A = 0.60Figure .- Continued.Provided by IHSNot for ResaleNo

47、 reproduction or networking permitted without license from IHS/.281.0 .8%e .6“2Q$ .4Q2 .2.$+0-.2-.4-4 0 4 8 /2 /6 20 24 Ang/e of attack, u, deg vIllUof+ f f y + ;+for M of 0.60.70 .75 .80 .85 .90 S4 .98 102 /.06 /.08(d) t/c = O.ti; A = o.lhFigure 5.- Continued.-5Provided by IHSNot for Resale-,-,-W03Ia I I I 1 I 1 1 C*)I I w I M PIr98I I I I I I I I II I Q 1.02A ll)R