NASA NACA-TR-903-1948 THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A-SERIES AIRFOIL SECTIONS《若干NACA 6A系列翼剖面的理论性和实验性数据》.pdf

《NASA NACA-TR-903-1948 THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A-SERIES AIRFOIL SECTIONS《若干NACA 6A系列翼剖面的理论性和实验性数据》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TR-903-1948 THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A-SERIES AIRFOIL SECTIONS《若干NACA 6A系列翼剖面的理论性和实验性数据》.pdf（25页珍藏版）》请在麦多课文档分享上搜索。

1、tJ_. ;- _O_ flagged symbols indicate NACA 64A-series sections with standard roughness.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-THEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A-SERIES AIRFOIL SECTIONS 19T,ift.-The section angle of zero

2、 lift as a function of thick-ncss ratio is shown in figure 12 for NACA 64- and 64A-seriesairfoil sections of various cambers. These results show thatthe angle of zero lift is nearly independent of thickness andis primarily dependent upon the amount of camber for aparticular type of mean line. Theore

3、tical calculations madeby use of the mean-line data of figme 3 and reference 1indicate that airfoils with the a=0.8 (modified) mean lineshould have angles of zero lift less negative than those withthe a= 1.0 mean line. Actually, the reverse appears to bethe case, and this effect is due mainly to the

4、 fact that air-foils having the a= 1.0 type of mean line have angles of zerolift which are only about 74 percent of their theoretical value(reference 1), and those having the a-0.8 (modified) meanlines have angles of zero lift larger than indicated by theoryThe measured lift-curve slopes correspondi

5、ng to the NACA64-series and NACA 64A-series airfoils of various cambersare presented in figure 13 as a function of airfoil thicknessratio. No consistent variation of lift-curve slope withcamber or Reynolds number is shown by either type of air-foil. The increase in traihng-edge angle which accompani

6、esremoval of the cusp would be expected to reduce the lift-curve slope by an amount which increases with airfoil thick-ness ratio (references 3 and 4). Because the present datafor the NACA 6A-series sections show essentially no varia-tion in lift-curve slope with thickness ratio, it appears thatthe

7、effect of increasing the trailing-edge angle is about32“ - (/VACA_ 0-_4-2_-400 0 2r_ 00:4“.2(o-_44 8 /2 /6 20 24A/r-foE fh/c/ness_ percent of chordFnURE 12.-Variation of section angle of zero lift with airfoil thickn_s ratio and camber forsome NACA 64-series (reference 1) and NACA 64A-series airfoil

8、 sections. R=6Xll_ 6./4,s,./,2LQ ./6.oo -L“4 .060el)_ o OB .2 _ NACA 6#h-ser/exo ., I ISmooth_Rough J NAC, 64-series-I I , I I I4 8 / 2 16 20 24A/rfoH th/cktless_ percent of chordFn_uBl_ 13. Variation of lift-curve slopc with airfoil thickness ratio for some NACA 64-scrio_(reference I) and N A (A 64

9、A-sl!rics airfoil sections of w_rious camhcrs hoth in the smoothcondition and with standard leading-edge roughness. R=6X10“; flagged symbols indicattlNACA 64A-series sections with standard roughness.balanced by the increase in lift-curve slope with thicknessratio shown by NACA 6-series sections. The

10、 value of thelift-curve slope for smooth NACA 64A-series airfoil sectionsis very close to that predicted from thin airfoil theory (27rper red|an or 0.110 per degree). Removing the trailing-edge cusp from an airfoil section with standard leading-edgeroughness causes the lift-curve slope to decrease q

11、uiterapidly with increasing airfoil thickness ratio.The variation of the maximum section lift coefficient withairfoil thickness ratio and camber at a Reynolds numberof 6X108 is.shown in figure 14 for NACA 64-series andNACA 64A-series airfoil sections with and without standardleading-edge roughness a

12、nd simulated split flaps deflected60 . A comparison of these data indicates that the char-acter of the variation of maximum lift coefficient with airfoilthickness ratio and camber is nearly the same for the NACA64-series and NACA 64A-series airfoil sections. The magni-tude of the maximum lift coeffi

13、cient appears to be slightlyless for the plain NACA 64A-series airfoils and slightlyhigher for the NACA 64A-series airfoils with split flaps thancorresponding values for the NACA 64-series airfoils. Thesedifferences are small, however, and for engineering applica-tions the maximum-lift characteristi

14、cs of NACA 64-seriesand 64A-series airfoil sections of comparable thickness anddesign lift coefficient may be considered practically the same.0I(b)0 4!I (?zi 1,NACA 64-series- o .4 _! - -/ . _-“1“-1 - - -.2NACA 64A-ser,es t I I ,2 I2 IG 20 Z4A/rfoH /h/c/_ness_ percent of chord(a) Airfoil with simula

15、ted split flap deflected 60“.(b) Plain airfoilI,o;_sRt,: 14. -Variation of maxinmm section lift cocllieicnt with airfoil thickness ratio andcamber for some NACA 64-series (reference 1) and NACA 64A-serit,s airfoil sections withand witlmut simulated split flaps and standard roughness. R_6X10_; flagge

16、d symbolsindicate NACA fi4A-serics airfoils with standard roughness.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2O REPORT NO. 903-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICSA comparison of the maximum-lift data for NACA 64A-series airfoil section

17、s, presented in figures 4 to 10, withsiinilar data for NACA 64-series airfoil sections indicatesthat the scale-effect characteristics of tile two types of sectionare essentially the same for the range of Reynolds numberfrom 3 X l06 to 9 X 106.Pitohing moment-Thin-airfoil theory provides a meansfor c

18、alculating the theoretical quarter-chord pitching-momentcoefficients of airfoil sections having various amounts and./J“1,-_ 0(D0 -./“_ -_3I I% I IfNACA 64A-ser/es_o 0o .2-0 .4“-2I I _li(NACA G4-ser/es)Iu I_ “.4-4o 4 8 /2 /6 20 24Aii,“*foi/ 7/“t/ck,“tes$ t percent of c_oi.“d(a) Plain airfoil. ,(b) Ai

19、rfoil with simulated split flap deflected 60,FIGURE 15,-Variation of section quarter-chord pitching-moment coefficient at zero angle ofattack with airfoil thickness ratio and camber for some NACA 64-series (reference 1) andNAOA 64A-series airfoil sections with and without split flaps. R=6X 106; flag

20、ged symbosIndicate NAC.& 64A.series airfoils with 60 simulated split flap, -./O% “-,08o_ -.06d -.04_ :02E0J_ACA 64A2/0 /,v,4cAe3Az, .,v,cA64,A2/z-I !i:oNACA 64aA2/5-/,-.0211 /NACA 64A410- /NA CA/. YVACA 64-2/0. _VACA 641-2)2.IVACA 642-215y-.04 -.06 _ 08T/Teore f/col moment coe?_clenf for o/rfoilmeo,

21、“t /me obouf querfer-chord pofHf641-4/2 /“-,/0FIGtT:aR.16.-Comparison of theoretical and measured pitching-moment coefficients for someNACA 64-series and 64A-series airfoil sections. R=6XI06.types of camber. Calculations were made according to thesmethods for airfoils having the a= 1.0 and a=0.8 (mo

22、difiedmean lines by using the theoretical mean-line data presente,in figure 3 and in reference 1. The results of these calculations indicate that the quarter-chord pitching-moment coefficients of the NACA 64A-series airfoil sections having tha=0.8 (modified) mean line should be only about 87 perceno

23、f those for the NACA 64-series airfoil sections with tha=l.0 mean line. The experimental relationship betwee:the quarter-chord pitching-moment coefficient and airfoithickness ratio and camber, shown in figure 15, discloses thathe plain NACA 64A-series airfoils have pitching-momencoefficients which a

24、re slightly more negative than those fothe plain NACA 64-series airfoils. The increase in thmagnitude of the pitching-moment coefficient of NACA 64Aseries airfoils as compared witlt NACA 64-series airfoilbecomes greater when the airfoils are eq uipped with simulate,split flaps deflected 60 . A compa

25、rison of the theoretic_and measured pitching-moment coefficients is shown in figur16 for NACA 64-series and 64A-series airfoil sections. Thescomparative data indicate that the NACA 64A-series sectionmuch more nearly realize their theoretical moment coefficientthan do the 64-series airfoil sections.

26、Similar trends haybeen shown to result when mean lines such as the a=0.type are employed with NACA 6-series airfoils (reference 1)_erodynamic eenter.-The position of the aerodynamicenter and the variation of the moment coefficient with lifcoefficient about this point were calculated from the quarter

27、chord pitching-momentdata for each of the seven airfoiltested. The variation of the chordwise position of the nerodynamic center with airfoil thickness ratio is shown in figur17 for the NACA 64-series and 64A-series airfoil section.,Since the data for the NACA 64-series airfoils showed nconsistent v

28、ariation with camber, the results are represente,by a single faired curve for all cambers. Following this samtrend, the position of the aerodynamic center for the NAC_64A-series airfoils shows no consistent variation with cambe_The data of figures 4 to 10 show that the variations in thReynolds numbe

29、r have no consistent effect upon the chordwise position of the aerodynamic center.Perfect fluid theory indicates that the position of thaerodynamic center should move rearward with increasinairfoil thickness and the experimental results for the NAC_64-series airfoil sections follow this trend. The d

30、ata, c_. .2827._ .26o 0I _i I- a .2 _ _CA 64A-se_/es-NACA _4-_er/e_ .i-I“IIL_fIf2 4 6 8 /0 12 14 /6 /8 20 2,A/rfo_7 f/_/ckmess, percent chordFIGURE 17.-Variation of chordwise position of aerodynamic center with airfoil thickncratio for some NACA 64-series (reference 1) and 64A-series airfoil section

31、s of differc_cambers. R=6XI0 _.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TI-IEORETICAL AND EXPERIMENTAL DATA FOR A NUMBER OF NACA 6A-SERIES AIRFOIL SECTIONSreference 5 show important forward movements of the aero-dynamic center with increasing

32、trailing-edge angle for agiven airfoil thickness ratio. The results obtained for tileNACA 24-, 44-, and 230-series airfoil sections (reference 1)reveal that the effect of increasing trailing-edge angle pre-dominates over the effect of increasing thickness because theposition of the aerodynamic cente

33、r moves forward withincreasing thickness ratio for these airfoil sections. For theNACA 64A-series airfoils (fig. 17) the aerodynamic center isslightly behind the quarter-chord point and does not appearto vary with increasing thickness. These results suggestthat the effect of increasing thickness is

34、counterbalanced byincreasing trailing-edge angle for these airfoil sections.CONCLUSIONSFrom a two dimensional wind-tunnel investigation of theaerodynamic characteristics of five NACA 64A-series andtwo NACA 63A-series airfoil sections the following conclu-sions based upon data obtained at Reynolds nu

35、mbers of3X106, 6X106, and 9X106 may be drawn:1. The section minimum drag and maximum lift coef-ficients of corresponding NACA 6-series and 6A-series airfoilsections are essentially the same.2. The lift-curve slopes of smooth NACA 6A-series airfoilsections appear to be essentially independent of airf

36、oilthickness ratio, in contrast to the trends shown by NACA6-series airfoil sections. The addition of standard leading-edge roughness causes the lift-curve slope to decrease withincreasing airfoil thickness ratio for NACA 6A-series airfoilsections.3. Tile section angles of zero lift of NACA 6A-seri(

37、airfoil sections are slightly more negative than those (comparable NACA 6-series airfoil sections.4. The section quarter-chord pitching-moment coeificientof NACA 6A-series airfoil sections are slightly more negativthan those of comparable NACA 6-series airfoil sectiomThe position of the aerodynamic

38、center is essentially indpendent of airfoil thickness ratio for NACA 6A-series airfosections.LANGLEY _EMOR1AL AERONAUTICAL LABORATORY,NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,LANGLEY FIELD, VA., _l/iay 6, 1957.REFERENCES1. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis SJr.: Summary

39、of Airfoil Data. NACA Rep. No. 824, 1945.2. Jacobs, Eastman N., Ward, Kenneth E., and Pinkerton, Robert M.The Characteristics of 78 Related Airfoil Sections from Tests i_the Variable-Density Wind Tunnel. NACA Rep. No. 460, 19333. Purser, Paul E., and McKee, John W.: Wind-Tunnel InvestigatioJof a Pla

40、in Aileron with Thickened and Beveled Trailing Edges ola Tapered Low-Drag Wing. NACA ACR. Jan. 1943.4. Jones, Robert T., and Ames, Milton B., Jr. : Wind-Tunnel Investigation of Control-Surface Characteristics. V-The Use ofBeveled Tlailing Edge to Reduce the Hinge Moment of a ControSurface. NACA ARR,

41、 March 1942.5. Purser, Paul E., and Johnson, Harold S.: Effects of Trailing-Edg(Modifications on Pitching-Moment Characteristics of AirfoilsNACA CB No. L4130, 1944.u. s. GOVERNMENT PRINTING O.:FICE: 1950Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-