NASA NACA-CB-L4G10-1944 Wind tunnel investigation of NACA 66(215)-216 66 1-212 and 65(sub 1)-212 airfoils with 0 20-airfoil-chord split flaps《对带有0 20翼弦分裂式襟翼NACA 66(215)-216 66 1-21.pdf

《NASA NACA-CB-L4G10-1944 Wind tunnel investigation of NACA 66(215)-216 66 1-212 and 65(sub 1)-212 airfoils with 0 20-airfoil-chord split flaps《对带有0 20翼弦分裂式襟翼NACA 66(215)-216 66 1-21.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-CB-L4G10-1944 Wind tunnel investigation of NACA 66(215)-216 66 1-212 and 65(sub 1)-212 airfoils with 0 20-airfoil-chord split flaps《对带有0 20翼弦分裂式襟翼NACA 66(215)-216 66 1-21.pdf（16页珍藏版）》请在麦多课文档分享上搜索。

1、l!l!lljippj-;,(,C ,.,“2T I.+.+”J.,* , .+ ,6-ww 1:.? !L- 140 1.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA C13iO . L4.G1ONAT1ONAL ADVIsoilycoMMITTEfiFOR AERoNAuTIcsC-OIWIDENTIALBULLETINWIIJ-TtThTNELINVESTIGATION OF NACA 66( 215 )-216,66,1-212

2、, A3D 651-212 AIRFOILS wImO.20-AIRFOIL-CHON3 SPLIT FLAPSBy l?ljcienF. Fullmer , Jr.SUMMARYAn.invest iatim was carried out in the NACA two-dii?lenSional low-turbulence pressure tunne1 of theNACA f16(215 )-216, 66,1-212, and 651-212 airfoil.sectionseqlippedwith split flaps having chord.s :?0percent o.

3、fthe airfoil.chord. The purpose was to determine themaxitum-l,j.ftcharacteristics of these low-drag airfoilSections withsplit flaps. All the present tests weremade at anywith chordwise laminations, and the surfaceswere paLnted and sanded “LliIti I aerOdynairlica.1y smooth.The slitflaps wee simulated

4、 by triangular blocks oflaiiinatedmahogany atta.chcd to the lower surface of themode1 One face of tb.eblock was cut to the centour ofthe flap portion of the airfoii lower surface. AI:jrpicalarrancment i.sshown in fi.ule1.Provided by IHSNot for ResaleNo reproduction or networking permitted without li

5、cense from IHS-,-,-,.I3RESULTS AND DISCUSSIONWiesection lift and pi.tching-momen.t characteristicsfor the NACA 66(215)-216, 66,1-212, and 651-21.2 airfoilsecti.on.sare presented in figures 2, 3, azmi.,respec-tivel. The lift and pi.tchi.ng-mcxnentcharacteristicsof the plain airfoil are included for c

6、omparison withtlheairfoils with flaps deflected. A comparison of theinaximumlift coefficients of the three sections testedin the present investigation is ,gtve.nin figure , withsimilar data fop the NACA 23012 airfoil from reference 2.Figu-,”e6 shows the variation of the increment of maximumsection l

7、ift coefficientAc with flap deflectionmaxfor the various airfoils.M examination of figure 5 3hows that higher maximumlfts vere obtained wih the plain NACA 651-212 airfotl.than with the plain NACA 66,1-212 airfoil. Khen thefap wpe deflated, however, the maximum lift coeffi-j.entsfor both airfoils wer

8、e approximately equal. Asimilar comparison between the two HACA 66-seriesairfoils shows that considerably hj.gherm.axim.umliftcoeif+.cientsfor all flap defleCtiOIW3were obtained ,riththe 1-percent-thick airfoil. The increments of maxirnrunlift c.oef.fici.entfor this airfoil sectjon were, on the:Pf:,

9、Z5Lper?ent hfgher than the increments obtainedwith the l!TACA6b,1-212 airfoil section. (See fig. 6.)We increased maximum lift coefficients for theriT;CA66(215)-216 air.fotlare attributed to the greaterthickness and consequent increase in leadin-edge radius.?igm”ep also shows that the maximum 1.ftcoe

10、fficientsobtained with the plain NACA 66(215)-216 air-oilat aReynolds r.umberof 6 x 106 erea;?proxirnatelythe sameas those obtained from tests of the NACA 25012 airfoilofreference 2 at an effective Reynolds numberof 3.5 x lo. For most flap deflections tested, theI values of CL and Act (figs. 5 and 6

11、) obtainedmax maxwith tb.e16-percent-thick low-drag airfoil were higherthan those obtained with the 12-percent-thick conventionalairfoil,IProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- .-1Provided by IHSNot for ResaleNo reproduction or networking p

12、ermitted without license from IHS-,-,-3RESULTS AND DISCUSSIONm. e section liftand.pitching-moment cha.ractieristicsfoithe HACA 66(215)-216, 66,1-212, and 651-212 airfoilsectlocs are presented in figures 29 3, and , resEec-ti.vel?. T lift and pitchin-filomentcharacteristicsofthe lain airfoilare inclu

13、ded for comparison withthe a.ifilswith flaps deflected A comparison of theii.;iurlift coefficients of the three sections tested.in the present investigation is given in figure 5, withshui.lardata forthe NACA 23012 airfoil from reference 2.FiU”Lae6 snows the va-riationof the incrffientof maxisectilon

14、lift coefficient ACJ with flap deflectionmu.for the various airfoils.An examination of figure 5 shows that higher maximumlj.ftsvere obtained with the plain NACA 651-212 airfoilthan lviththe plain NMM 66,1-212 airfoil. When thefls.swere d.eflected,however, the maximum lift coeffi-cinfisfor both airfo

15、ils were approximately equal. Aj.fl.lalcoflparj.sonbetween the two NACA 66-seriesairfoils shovschatconsiderably higher m-aximmm litcoeficients for all flapdeflections were obtained withthe 16-percent-thick airfoil. The increments ofmaximumlifb coefficient for this airfoil section were, on the:vepzge

16、 34 percent higher than the increments obtainediththe JjTACA661.-212 airfoil section. (See fig. 6.)The increasd maximum lift coefficients for the?ACA 66(215)-216 airfoil are attributed to the greaterthickness and.consequent increase in leading-edge radius.pf,Tljj. ZISO shows that the maximum lift co

17、efficientsobzained.with Khe plain NAC.A66(215)-216 airfoil at aReynolds number of 6 :106 *;ereapproxirlatelythe sameasYhose obtained-from tests of the NACA 25012 airfoilof relerence2 at an effective Renolds numberof . l+. For most flap deflections teted$ theValues of Cz and. Act (figs. 5 and 6) obta

18、inedmax maxwith the 16-percent-thick low-drag airfoil were higherWan those obtained with the 12-percent-tinickconventionalairfoil.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.SUMMARY OF RESULTSb-emaximm lift coefficients of three low-dragairfoils

19、 without s.ndwith 0,20-airfoil-chord split flapsobtained from tests at a Reynolds number of approxi.-m.2tsQ 6 x 106 are as fOliOws:LanS;j Memorial Aeronautical LaboratoryNational Advisory Co.mdttee for AeronauticsLangley Fieidj Va.1. Tacobs, Eastman N., Abbott, Ira H., and Davidson,THlton: Prelimina

20、ry Low-Drag-Airfoil and FlapData from Tests at Large Rejmolds N-i-rnbersandLow lurbulenceYand Supplement. NACA AC,yc 19*.2.“enzinger, carl ., and Harris, Thomas A.:Vind-Tunnel Investigation of N.A.C.A. 250123502Jjand 23030 Airfoils with Various Sizes ofSplit Flap. NACA Rep. No. 668, 1939.-.- - . ._

21、,=- .=, .,-.,-. :, . . . . . :.;. .: . . . . . .,. $ ,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-i. . .- .IA I. NACA 66(215)-216 AIRFOIL Stations and ordinates are givenin percent of airfoil chord -. .-.- -. -iI Upper surface I Lower surface- .

22、. . .=e ,-Station7;.50:wj .fj)i,).yj*3.00.$31:40128IiIiiIfo1.230blJ !1.852:j:oz4tia71 2.lL.O6.2767iw1.3668.7368.$1809.0929.06080875:;Zx .9258.yb9.972Ordinate.o-2.9 2$z- :o:4.930+.564-6.05J-6.422-6.676-6.838-6.902-6.654-6.68t-6.35 .- .802?:t: Z;-3.021-2.0 9-1.069-.2810I L,E. radius: 1.575Slope of rad

23、ius through L.E.: 0.084.NAT10NAL ADVISORYCOMMITTEZ FOR A330NAUT1CS.,. ,-”- - - .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. .- ., . _TABLE IT.- NACA 66,1-212 AIRFOILS”tationsand ordinates are givenairfoil chordb in percent of_, ._-. . .Upper su

24、rfaceStation(1-. .Ordinate.0a71 94-76.5226.8165.7592.7701.760a71 7920_. - - -Lower surfaceStationo.576a7134z1.3 32.6055.1167.62110.12215.11626.1.052L5.Q9130.075.Ordinateo-.847-1.010-1:25?-l,il-2.165-2.593-2.y5-5.529t- .9d2-322:. 5;:-4:863-4.903-4.869-)+.749:;,3a71 563-2.895-20167-1.424-.726-.160L.E.

25、 radius : 0.893Slope of radius through L.E.: 0.031.1NAT1ONALADVISORYCONUITTW3FOR A3RONATJTICSProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA C3 !?0 . 4G1O sl 7,.- . -. . . -I!TABLE 111.- NACA 651-212 AIRFOIL i, Stations and ordinates are givenii

26、n percent of airfoil .chord I.-e -i Upper surface. -. II Skation ! OrdfnateI:-)IIIII0.9701.176l.l+pl2.0582.919;:3(6oLower surface 1-p. .I+Station / Ordinate /-”-t-”-”“ !10IIJ !. _L.E. radius: 0;932Slope of radius throu L.E.:i 0.0811.-. INATIO?M.LADVISORYC(XWITT:.TFOR AZROVATJTI(X3Provided by IHSNot

27、for ResaleNo reproduction or networking permitted without license from IHS-,-,-zo.IcWOODEN BLOCKCOMMITIEEFORAERONAUTICSFigure l.- View showing the NACA 66(215)-216airfoilwith 0020c split flap deflected6000%Reynolds number, R, 6 X d. Teats, TDT 247, 568, 571.Provided by IHSNot for ResaleNo reproducti

28、on or networking permitted without license from IHS-,-,-NACA CB No. L4G1O Fig. 3“-24 -16 -8 0 8 16 24Section angle of attack,aOFigure .7with a.20-.2-.4-.6COATMITEFOR AERONAUTICSo._. .-. 4 0 .4 .8 1.2 1.6 2.0 z.4 2.8SectIonllftcoefflclemt,C,.- Section lift and pitchln%-moment characteristic for an NA

29、CA 66,1.212 .drfc.110.20c split flap; R, 6 x 10 . Tests, TDT b21L,570, 576, 602.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA CB No. L4G1O Fig. 4o-*t-l.-l,+2.8 I I2.4(deg)2.0 0 0(plainairfoil) -+Lol, lI 0701 I I I I R, 6 x 106. Tests, TDT 356,

30、 569, 599.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA CB No. L4G1O Fig. 5c!o2i)alal.-,NACA 66(215)-2162.86 x 106 (approxO)NACA 230122*k : EffectiveR = 3.5 x 106P (fromreference2)/ J NACA 66,1-212/ / R= 6 x 106 (approx.)/2.0 , 9/ f/ / r Xrlln

31、!+CL 919/ / , =/ L Lwatifi V -c=LC/ ,/1R=6 x 106 (approx.)“z 1.2.8I INATIONALAOVISORYCOMMITTEEFORAERONAUTICSI I I00 1 I I20 40 60 80Flap deflection, bf # degFigure50- Effect of flap deflectionon maximum sectionliftcoefficientfor the variousairfoilsections.Provided by IHSNot for ResaleNo reproduction

32、 or networking permitted without license from IHS-,-,-NACA CB No. L4G1O Fig. 6$ 1a)co2c)a)ccl%!oI I.2NACA66(215)-216= 6 x 106 (approx.).0 -NACA 23012fective R = 3.5x 106(fromreference2).8 N-CA66 1.212R = 6 x 1Q6 (apPrC)X.).6 R = 6 x 106 (approx.)a714NATIONALADVISORY.2 COMMITTEEFORAERONAUTICSo0 20 40 60 80Flapdeflection,bf , degFigure6.-Eectof flapdeflectiolon theilcrementof maxiinumsectionliftcoefficientfor thevariousairfoil-flaparrangements.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-