NASA NACA-TN-1813-1949 A study of flow changes associated with airfoil section drag rise at supercritical speeds《在超临界速度下和翼剖面阻力增长相关的流量改变的研究》.pdf

《NASA NACA-TN-1813-1949 A study of flow changes associated with airfoil section drag rise at supercritical speeds《在超临界速度下和翼剖面阻力增长相关的流量改变的研究》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TN-1813-1949 A study of flow changes associated with airfoil section drag rise at supercritical speeds《在超临界速度下和翼剖面阻力增长相关的流量改变的研究》.pdf（32页珍藏版）》请在麦多课文档分享上搜索。

1、J-TECHNICALNOTENo. 1813 -#A STUDY OF FLOW CHANGES ASSOCIATED WITH AIRFOILSECTION DRAG RISE AT SUPERCRITICAL SPEEDSBy Gerald E. Nitzberg and Stewart CrandallAmes Aeronautical LaboratoryMoffett Field,Calif.-Washington. . . . . . . ., _ . .I-/-.Provided by IHSNot for ResaleNo reproduction or networking

2、 permitted without license from IHS-,-,-TECHLIBRARYKAFB,NMIlfllllllllllnlullnllnlNATIONAL ADVISORY COMMITTEEFOR AERONMJTWS 0144830TECHNICAL NOTE i-To. 1813ASTODYOF l?LOWCHANGES Assccm WITH AIKFonlSECTION DRAG RISE AT SOI?ERCRITICAZSPEER3By Gerald E. IVitzbergand Stewart CrandallSUMMARYA study of exp

3、ertientalpressure distributionsandteristios for severalmoderately thick airfoil sectionssection characwas made. Acorrelation appears to exist between the drag-divergenceMach numberand the free+ tieam Mach number for which sonic velocity occurs at theairfoil crest, the chordwise station at which the

4、airfoil surface istangent to the free+ tream direction. It was found that, since the Mchnumber for which sonic velocity occurs at the airfoil crest can beesttited satisfactorilyly means of the Erandtl+lauert rule, a methodis provided whereby the drag+iivergenceMach numler of an airfoilsection at a g

5、iven angle of attack oan be estimated frcm the low-speedpressure distributionand the airfoil profile. This method was usedto predict with a reasonable degree of accuracy the drag-diverwhereas for other types no appreciable drag rise occursuntil the Mach number of the free stream is osidera%ly above

6、thecritical.No adequateDthod has been presented for predicting the freestream Mach nurriberaat which the various suercriticalflow ohangesOclmra71 The pwpose of the zwsent reort is to investigate theseflow ohanges for twcwihens ional transonicflow past conventional andlow-drag airfoil sections. The i

7、nvestigationis based prtmarily ona systematicanalysis of eerimental data for airfoil sections oflercent+hord thickness. Although this thickness is greater thanis generally desirablefor usethat airfoil+hape effects arelayer effects.accdclM%Ma%at high speeds, it has the advantageless lilmly to he mask

8、ed by boundary-SYMBUISairfoil seotion angle of attiokairfoil chordairfoil section drag coefficientairfoil section lift ooeffiolentratio of local velocity to local velocity of soundfree-streamMach nuuiberat which sonic velocity is firs%reached at airfoil crestcriticalIkch numler of airfoil section (f

9、ree-abeamMach number at which local sonic velocity is firstreached on airfoil surface)drag-divergenceMach number (Mach number at which slopeof curve of drag coefficientversus Mach number attainsa value of 0.10)ratio of freewhereas for other cases it does not begin until the free-treamMach number is

10、considerably greater than the critical. Sme high-speed pressure distributionsand section characteristics (reference1)for the NACA 652-215 (a = 0.5), 66,2-15 (a = 0.6), 0015, 23015, and4415 airl?oilsections have been studied in an attempt to detezminethe flow changes which are associated with the app

11、earance of this moreor,less abrupt drag rise. In order to supplement these data,simultaneouspressure distributionsand schlierenpictures of the flowaround the NACA 23015 airfoil section were obtained in the Amesl-by 31/2+Poothigh-speedwind.tunnel. Before considering thespecificproblem of the flow cha

12、ns associatedwith the appearanue ofthe supercriticaldrag rise, the data for this airfoil section will bediscussed in detail.- z . . . . . . , - . -.,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 IJACA0. 1813Measurements at SupercriticalSyeeds for

13、IW2A 23015 Airfoil at 20 IhcidenceL and, at least up to a Mach nwhber of 0.8, this pressure-coefficientvariation with free-streamM but,for these cases, it ssts the possibiliof predicting the peak pressure ooefficient. M figure 10 the ,eQeri-mental local I although, of course, this theory is not vad

14、for a supesonic region of limited extent normal to the airfoil chord. Xt i8seen that a close approximation to the experimental local Mach numberdistribution can be obtainedby taking one-half of the Prandtl+eyerlhch number incrementper degree of surface turning. Figure 11 ShOWSthat, for the airfoil s

15、ections comidered in figure 9, a usefulapproximation to the local lkch nuniberat the airfoil crest for I 1945.2. Idndsey, W.F., IkQey, l?ernardN., and Humphreys, Milton D.:The llowand Force Characteristics of SupersonicAirfoilsat Hi Subsonic Speeds. MICA TN IVo.12U, 1947.3. Allen, H. Julian, Heaslet

16、, Max. A., and itzberg, Gerald E.:The lhteracticm of B-nmdary Iayer and Ccmgmession Shook andits Effect Upon Mrfl Pressure Distributions. l!lACARM No.A7A02, 1947. “ -#I.- - . - . . . . .- -. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, NACA TIvN

17、o. 1813r Xl.Free -streum MQch numbe MO -Figure /.- Section lift and df og coefficients 0s functions offree - s)reum Much number for NACA 230/5 uirfoilsection d 2 angle of oftuck. ,. .,. .+- .-, _ , - . - ._ . . . - -.-.!.= - . . . . -.,- .,.Provided by IHSNot for ResaleNo reproduction or networking

18、permitted without license from IHS-,-,-,.,a71.,- ,.,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. .ChofowwSwhir x/cFigure 2, Simw7aneous/y obtane.d pressure okitributions and sch/ieren photographs forA!ACA 23015 oirfoll section; Q, 2T 15Provided

19、by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-._. . .,. +.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-iIawvwe stafkw,x/cqj!./ /. - 45 - .5.55 .6 .65 .7 .75 .lFree- streum Mach wnbe Me wHgure 8.- 6bundar.

20、es of supersonic re ion on upper surfoce offNAGA 23015 airfoil section us a unction of free -streumMuch number fw vurious angles of attack.h*. -.- .- -. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN NO. 1813 29Mach numbe MeFiwe 9.- Location of forwurd sonic point us o function of free-shwm Mach number for several NAGA airfoil sections. _ _ . .- . . -. _ - . ._ _ _ _ _ _Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-