NASA NACA-TN-3162-1954 Effects of subsonic Mach number on the forces and pressure distributions on four NACA 64A-series airfoil sections at angles of attack as high as 28 degrees《在.pdf

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1、,. . -.,A? da71 ?=dNATIONALADVISORYCOMMIITEE .s0=FOR AERONAUTICS r=(n mco=w=Jl = “mTECHNICAL NOTE 3162 =Z zLOAN COPY: RETURIm-w.- (s IL)KIF?TLAND M=, N.EFFECTS OF SUBSONIC MACH NUMBER ON THE FORCES ANDPRESSURE DISTRIBUTIONS ON FOUR NACA 64A-SERIES.AIRFOIL SECTIONS AT ANGLES OF ATTACKAS HIGH AS 28By

2、Louis S. Stivers, Jr,Ames AeronauticalMoffett Field,LaboratoryCalif.WashingtonMarch 1954d.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.LNATIONAL ADVISORY COMMITTEE FORTECCa Nom 3162EFFECTS Ol?SUBSONIC MACH NUMBER ON THE FORCES AND.PmmmRE DISmIBUT

3、IONS ON FOUR NACA 64A-SERIESAIRFOIL SECTIONS AT A?GLESOF ATTACKAS HIGH AS 28By Louis S. Stivers, Jr.EWMMARYLift, drag, moment, and pressure-distributionmeasurements havebeen made for the NACA 64AO1O, 64A41O, 64AO06, and 64A406 airfoil sec-tions at high subsonic Mach numbers. The tests were made for

4、angles ofattack as high as 280 and for Mach numbers ranging from 0.30 to about0.93 with correspondingReynolds numbers varying from approximatelyO*9X 106 to 1.9X 106.* A comparison of the msximum lift coefficients from NACA TN 2096for 10-percent-chord-thickNACA 64A-series airfoil sections camberedwit

5、h a = 1.0 and a = 0.4 mean lines with those of the present report.for the NACA 64A41O airfoil section cambered with the a = 0.8 (modified)mean line indicated that the a = 0.8 (modified)mean line was superiorfor providing high maximum lift coefficients throughout the Mach numberrange, especially for

6、Mach numbers above about 0.6.As the angle of attack was increased above that for the maximumlift coefficient obtained at about 8 to 10 angle of attack, the sym-metrical.airfoil sections experienced no serious losses in lift coeffi-cient. In fact, the lift coefficients for the symmetrical airfoilsect

7、ions and for the NACA 64A406 airfoil section at angles of attackabove 24 reachedvalues greater than the respective initial maximumlift coefficients obtained at the lower angles of attack.A region of slight compression,heretofore undescribed, was estab-lished within the local supersonic region on eac

8、h of the airfoil sectionsnear the leading edge in place of an expected expansion. This leading-edge compression reg-ionwas formed just downstream of the abruptProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2eansion at the leading edge forattack that

9、 v=ied in some degreeNACA TN 3162ranges of Mach number and angle ofwith airfoil-sectionthickness ratioand camber. As indicatedby the measured pressures on the surface ofthe airfoil sections, the flow over the leading edge expanded to maximumlocal Mach numbers froml.6 to 2.0 before the start of the l

10、eading-edgecompression region. When the leading-edge compressionregion was establ-ished on the airfoil sections, the lambda shock wave, which usuallydeveloped in the flow at high Mach nurbers, was not formed on the samesurfacey leaving only the normal shock wave.For angles of attack above that for c

11、omplete separationof theflow over the upper surface of each airfoil section, the pressure coef-ficients on this surface for a constant Mach number were essentiallyunsd?fectedby cadber of the airfoil section or by a reduction in airfoil-section thickness ratio from 0.10 to 0.06. The correspondingpres

12、surecoefficients on the lower surface,however, were increased noticeablyby the increase in ember orme relative simplicityby the decrease in thickness ratio.I17TRODUC1710Nwith which the subsonic aerodynamic charac-teristics of unswept wings fiybe calculated from section data employingliftiti-line the

13、ory (see ref.-l) has been appreciated for many years and,more recently, has been an incentive for establishing a similarproceduresuitable for swept wings. One recent effort to determine local sectioncharacteristicsof sweptbackwings from two-dimension cd error Percent error0.3 0 -0.0007 to 0.0011 -5.

14、5 to 8.610 -.0003 to .0015 -1.0 to 4.928 .0117 to .0183 1.5 to2.4a717 o .0002 to .0004 1.5 to 3.110 .00 to .0080 2.0 tO 2.9a719 o .0001 tO .0016 .4 to 1.72 .0007 tO .0023 1.4 tol*7The errors in the test Mach ?mmibersand Reynolds numbers are less thankO.005 and 0.1 X 106, respectively.Provided by IHS

15、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3162 7RESULTS AND DISCUSSIONFORCE AND MOMENT DATALift CharacteristicsThe effects of Mach number on the section lift coefficients of theNACA 64AO1O, 64A41O, 64AO06, and fig. 3(c), a = 8;and fig. 3(d), w = 100).

16、 For 8 or 10 angles of attack, theseincreaseswere apparently caused by the rearward extension of localsupersonic flow over the forward portion of the upper surface, as isconfirmedby the pressure distribution data presented later in thisreport. In figure 4 the section lift coefficients for each airfo

17、ilsection are presented as a function of section angle of attack withMach number as a parameter. Maximum section lift coefficients are evi-dent for the lower Mach numbers of this figure at angles of attack ofabout 8 to 10. No serious losses in lift coefficient are noted forthe symmetrical airfoil se

18、ctions at higher angles of attack. At thehighest angles of attack shown the lift.:oefficientsfor these airfoilsections and also for the NACA 64A406” (b) the locaticm ofthe origin is not appreciably affectedly Mach number; and (c) this com-pression is not related to the normal shock wave but appears

19、rather tobe associatedwith the abrupt expansion region at the leading edge of theairfoil section. Furthermore, it should be realized that the two typesof mild compression do not appear simultaneouslyon the same surface ofthe airfoil section. In other words, when the compression that formsnear the le

20、ading edge is fully developed, no lambda shock waves formdownstream in the flow on that surface, but only normal shock waves.This will be evident in some of the schlierenphotographs which are pre-sented later in this report. In figures 13(f) to 13(i) it is observedthat the pressure increases associa

21、ted with the shock waves are morewidespread and less abrupt than those noted for the lower angles ofattack. Such a change in the character of the increases in pressureapparently results from the more pronounced boundary-layer separationwhich exists at the higher angles of attack and Mach numbers. Th

22、e extentof separation and the nature o-fthe shock waves at the higher Mach numberson the NAC!A64AO1O airfoil section at angles of attack of 6, 8, and 10are shown in the scblierenphotographs of figures 18(d) to 18(f). It isnoted in the photographs for the higher Mach numbers and angles of attackthat

23、the shock waves, although similar in shape to the lambda shock wavesu.w.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3162noted at the low angles of attack,cussed in that the oblique legs of.11differ from these previously dis-the waves appe

24、ar markedly stronger.In the pressure data for angles of attack from 4.2 to 8.2, amild pressure rise is observed on the upper surface that originates.near the leading edge and extends downstream to the abrupt pressureincrease associated with the normal shock wave. This slight compressionnear the lead

25、ing edge (hereinafterdesignated as the leading-edge com-pression) exists in the upstream portion of the local supersonicchnumber region where an expansion would be expected, indicating a changein the nature of the local flow over the upper surface in this region.To show the features of this leading-

26、edge compression region in moredetail, pressure coefficients on the upper surface for an angle ofattack of 6.2, given in table V, have been plotted for several Machnumbers above 0.61 in figure 19. It is noted in this figure that aslight pressure increase near the leading edge is just beginning for a

27、Mach number of 0.64, and as the Mach number is increased the compressionregion spreads downstream. Throughout the range of Mach numbers, how-ever, the origin of the compression remains essentially fixed betweenthe 2.5- and the 5-percent chordwise stations. Although the region isextensive and well de

28、veloped for Mach numbers of 0.71 and 0.77, the com-pression appears greatly diminished for Mach numbers of 0.82 and 0.85.At these higher Mach numbers, pressure increases of the type associatedwith lambda shock waves are apparent. To indicate the effect of free-% stream Mach number on the magnitude o

29、f the leading-edge compression,differences in local Mach numbers associated with the peak pressure atthe start of this region and the pressm at approximately the 0.10.chordwise station have been determined for several free-stream Mach num-bers, using figure 17. These differences in local Mach nunibe

30、rs AMzand the maximm local Mach numbers corresponding to the peek pressuresnear the leading edge hex are given in the following tableangles of attack of 4.2, 6.2, and 8.2:T0.63 to 0.64 - - -.66to .67 -0.34.71 -.10a71 74 -.05.76to .77 -.06.79 to .801 -.05ktlax1.241.351.441.431.401.37%= 6-POAMz-0.16-.

31、14-.10-.09-.06-.05MZm=1.601.611.581.581.561.51%= 8.20AMz-0.19-.17-.12-.10-.08-.06?msx1.401.441.411.381.521.60for.As the free-stresm Mach number is increased, a reduction of the differ-ences in local Mach numbers is observed for each angle of attack in the.Provided by IHSNot for ResaleNo reproduction

32、 or networking permitted without license from IHS-,-,-12 NACA TN 3162table, indicating a correspondingreduction in the strength of theleading-edge compression, It is also observed that the compression isassociated with high values of maximum local Mach numbers, especiallyfor the 6.2 and 8.2 angles o

33、f attack. These high values of local Machnumber (up to 1.61) are indicative of a strong expansion region justupstream of the leading-edge compression region.Substantiatingevidence that a compressionregion existed in theflow over the NACA 64AO1O airfoil section near the leading edge for con-ditions c

34、orrespondingto the data of figure 19 is given in the schlierenphotographs of figure 18(d). (The fixed bulbous shape which appears onthe forward portion of the upper surface in some of the schlierenphoto-graphs of this report is due to a chipped window in the wind-tunnel sidewall.) In the photographs

35、 of the present report, a light area is indic-ative of a compression region, and a dark area is indicative of anexpansion region.There is little evidence of shock-inducedcompression on the uppersurface of the airfoil section at angles of attack from 12.2 to 18.2,figures 13(j) to 13(m), and none at t

36、he higher angles of attack. Theextensive separationof the flow over the upper surface at these anglesof attack apparently obscured any effects of existing shock waves.NACA 64A41O airfoil section.-An examination of the pressure datafor the NACA 6h.A410airfoil section which are presented in figure 14a

37、nd table VI reveals that the characteristicsof the ressure distribu-tions within the local supersonicregions on this airfoil section are2enerally the same as those for the NACA 64AO1O airfoil section. TheLeading-edge compression region was formed on the upper surface atapproximately the same angles

38、of attack as for the NACA 64AO1O airfoilsection, but because of the camber, a compression region was also formedon the lower surface for angles of attack from -5 to OO. An inspectionof the differences in 10CSL Mach numbers in this region has indicatedthat the leading-edge compression on the NACA 64A

39、41O airfoil sectionwas stronger on the lower surface and weaker on the upper surface thanthe correspondingcompression on the symmetrical airfoil section.Furthermore, the maximum local Mach numbers associatedwith the peakpressures near the leading edge are greater on the lower surface andless on the

40、upper surface for the cambered airfoil section. Local Machnumbers as high as 1.7 to 1.8 were attained on the lower surface of thecambered airfoil section at angles of attack of -5 and -4.Schlieren photographs for the NACA 64A310, a . 1.0, airfoil section(differingfrom the NACA 64A41O airfoil section

41、 in amount and type ofczuiber)are presented in figure 20 to corroborate the foregoing remarksconcerning the characteristicsof the pressure variations within the.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA m 3162 13local supersonicregions on

42、the NACA 64A41O airfoil section. esephotographs were made during the investigationreported in reference 5.It is observed in figure 20 for an angle of attack of -4 that a rela-tively strong leading-edge compression region was formed on the lowersurface at Mach numbers above 0.71, and that lambda wave

43、s were estab-lished on the upper surface at Mach numbers greater than about 0.81.The characteristicsof the leading-edge compression region on a cambered10-percent-thick airfoil section, as revealed in the photographs of fig-ure 20, are much the same as those for the NACA 64AO1O airfoil section.It is

44、 noteworthy that evidences of a leading-edge compression regionare apparent in the pressure-distributiondata of reference 11 for sec-tions of a 45 sweptbackwing of aspect ratio 3 employing the NACA 64A41Oairfoil section. The compression region appeared at the outboard stationsat subsonic free-stream

45、 Mach numbers of 0.86 and above for angles ofattack from about 7 to 10, and was established immediately dowmtresmof a strong expansion region along the leading edge wherein the localMach numbers attained values as high as 1.9.NACA 64AO06 airfoil section.- The pressure coefficients for theNACA 64AO06

46、 airfoil section are given in figure 15 and table VII. Acomparison of the coefficients for angles of attack of t2 and alsofor *l” at given chordwise stations, particularly near the leading edgeindicates that the model of the NACA 64AO06 airfoil section was not= perfectly symmetrical. Measurements ha

47、ve indicated that the asymmetryis due to small construction inaccuracieswhich, for this model, werelarger than usual. The ordinates around the leading edge and on the. lower surface near the leading edge were very close to those specified.On the upper surface, however, the ordinates between the 0.5-

48、 and aboutthe 10-percent-chordpositions were greater than the specified ordinates,the maximum difference being approximately O.1-percent chord. It shouldbe recalled, however, that the asymmetry of the lift coefficient data,shown in figure 3(c), is very small and is less than that observed forthe NAC

49、A 64AO1O airfoil section in figure 3(a). Irregular values ofcertain pressure coefficients near the leading edge, which probablyresulted from orifice errors, are also observed at angles of attackof -10, 0, and 1 and at the trailing edge at angles of attack from-1 to 10 for some of the Mach numbers. The curves have been fairedthrough these values.A comparison of the nat