NASA NACA-TN-3607-1956 Effect of thickness camber and thickness distribution on airfoil characteristics at Mach numbers up to 1 0《当马赫数为1 0时 厚度 弧形和厚度分布对机翼特性的影响》.pdf

《NASA NACA-TN-3607-1956 Effect of thickness camber and thickness distribution on airfoil characteristics at Mach numbers up to 1 0《当马赫数为1 0时 厚度 弧形和厚度分布对机翼特性的影响》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TN-3607-1956 Effect of thickness camber and thickness distribution on airfoil characteristics at Mach numbers up to 1 0《当马赫数为1 0时 厚度 弧形和厚度分布对机翼特性的影响》.pdf（77页珍藏版）》请在麦多课文档分享上搜索。

1、co I 00cc.)u-in “i“NATIONAL ADVISORYTECHNICAL NOTE 3607EFFECT OF THICSS, CAER, AND THICKNESS DISTRIBUTIONON AIRFOIL CHARACTERISTICS AT MACH NUMBERS UP TO 1.0By Bernard N. D-z:,idd:.-. ,i :=L J., . . . .)II-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-

2、,-,-. . IWH UBRAriY KAtti, NMN NATIONAL ADVISORY COMITTEE FOR AERONAUTICS lIllIUIM013LL4413TECHNICALNOlX3607EFl?ECTOF CKNESS, CAMBER, AND THICKNESS DISTRIEWONON AIRFOm CEMRACTERISTICSAT MACH NUMBERS UP To l.dBy Bernsxd N. IM.ey and Rich contrary to low-speed results the (n/d)= increased as eitherthe

3、 thickness ratio or the camber was decreased. At all Mach nmrsthe normal-force coefficientfor (n/d)H generally increased tithincreases in thickness ratio and camber with forwsrd movement of theposition of madlmml thjmkness. The trends of the data in the highestMach number range indicated that the no

4、rmal-force-curveslopes of alJ-airfoils tested sre approximatelyequal at Mach number 1.0, the valuebeing about the same as at low speeds.INTRODUCTIONEesigners of aircrsft and aircraft propellers have repeatedlyexpressed the need for airfoil-sectiondata in the trsnsonic-speedrange. Almost all section

5、data in the subsofic speed range have beenobtained from closed-throattunnels which inherently 13mit the speedrange of the tests to Mach mxnbers less than the choking value, gen-erally about 0.9. :F41.9192.ti32.5572.T772.8962.97-I2.9932.9b52.8252.6532.4*2Jea1.907l.lm1.Zg-.243-.m-.IX6-.o13Station ord3

6、nat4 station Ch-dlmteo.*.G?41.1072.3334.81.27.3049.80314.8x?w.82324.8%ay?:39:93244.96249.991Z.5g2:W3.4CCI4.227,4.8T75.3B2pJ6:x95:Fl5.95.*y!324:XI03.6172.7J1.870.942.01.3n o-.331-.373-.*-.474-.b94-.457-.418-.=-.=5-.I.24-.02Z.ca5:%.%1.663.8ti.W1.0s21.1021.057.844.592.W-.OI.3.61.2.8j6L3932.66-(s.mT.696

7、ti.ti15.1882Q.172%,069S4J%3159.94764.93669.9187#g84:88789.92194.9fiJm.oalL.E. radius: 0.246T.E. radius: 0.014Slopofrus though LX.: 0.238L.E. Mu: 0.246T.E. radius: 0.014Sloy of radiusthroughL.Il.: O.S .- . . _ . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license

8、from IHS-,-,-22 NACA TN 3607DiffuserCompressed -air the-1IIEnd-plate ossemblyAirfoil model?(Ar Nozzle blockEntronce coneA/“(a) Pictorial representation.Figure l.- Larigley4- by 19-inch semiopen tunnel. .-. .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS

9、-,-,-NACA TN 3607 23Station(inches)70$o 7).- a 64-OXX . - :7 a .9 1.0Machnumber, Mref.8)/It, = 0.10I4II /t& = o.ci3- .-0 /“/! II /I ! II/ It /1I / I,“7 .8 .9 1.0Machnumber, Ma717 .8 a719 1.0I&chnumber, MFigure 6.- Comparison of zero-liftdrag coefficients obtained from testsin the Langley 4- by 19-inch semiopentunnel with those obtained frcmtests by other methods.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-