NASA-TM-X-3160-1975 Experimental and theoretical low speed aerodynamic characteristics of the NACA 65 sub 1-213 alpha equals 0 50 airfoil《NACA 65(下标1)-213透明度为0 50的机翼的实验性和理论性低速空气动力特.pdf

《NASA-TM-X-3160-1975 Experimental and theoretical low speed aerodynamic characteristics of the NACA 65 sub 1-213 alpha equals 0 50 airfoil《NACA 65(下标1)-213透明度为0 50的机翼的实验性和理论性低速空气动力特.pdf》由会员分享，可在线阅读，更多相关《NASA-TM-X-3160-1975 Experimental and theoretical low speed aerodynamic characteristics of the NACA 65 sub 1-213 alpha equals 0 50 airfoil《NACA 65(下标1)-213透明度为0 50的机翼的实验性和理论性低速空气动力特.pdf（76页珍藏版）》请在麦多课文档分享上搜索。

1、ANDNASA TECHNICAL NASA TM X-3160MEMORANDUMI-(N ASA-TM-X- 3160) EXPERIMENTAL AND N75- 15601THEORETICAL LOW SPEED AERODYNAMICCHARACTERISTICS OF THE NACA 65 SUB 1-213,ALPHA EQUALS 0.50, AIRFOIL (NASA) 74 p HC Unclas4.25 CSCL 01A H1/01 09062EXPERIMENTAL AND THEORETICAL LOW-SPEEDAERODYNAMIC CHARACTERISTI

2、CS OFTHE NACA 651-213, a = 0.50, AIRFOILWilliam D. Beasley and Robert J. McGheeLangley Research CenterHampton, Va. 236656 ATIONA ARONATIANDSPA ADMINISTRATION WASHINGTON D. FEBRUARY 91NATIONAL AERONAUTICS AND SPACE ADMINISTRATION “ WASHINGTON, D. C. . FEBRUARY 1975Provided by IHSNot for ResaleNo repr

3、oduction or networking permitted without license from IHS-,-,-1. Report No. 2. Government Accession No. 3. Recipients Catalog No.NASA TM X-31604. Title and Subtitle 5. Report DateEXPERIMENTAL AND THEORETICAL LOW-SPEED February 1975AERODYNAMIC CHARACTERISTICS OF THE 6. Performing Organization CodeNAC

4、A 651-213, a = 0.50, AIRFOIL7. Author(s) 8. Performing Organization Report No.William D. Beasley and Robert J. McGhee L-977310. Work Unit No.9. Performing Organization Name and Address 505-06-31-01NASA Langley Research Center 11. Contract or Grant No.Hampton, Va. 2366513. Type of Report and Period C

5、overed12. Sponsoring Agency Name and Address Technical MemorandumNational Aeronautics and Space Administration 14. Sponsoring Agency CodeWashington, D.C. 2054615. Supplementary Notes16. AbstractLow-speed wind-tunnel tests have been conducted to determine the two-dimensionalaerodynamic characteristic

6、s of the NACA 651-213, a = 0.50, airfoil. -The results werecompared with data from another low-speed wind tunnel and also with theoretical predic-tions obtained by using a viscous subsonic method. The tests were conducted over a Machnumber range from 0.10 to 0.36. Reynolds numbers based on the airfo

7、il chord varied fromabout 3.0 x 106 to 23.0 x 10617. Key Words (Suggested by Author(s) 18. Distribution StatementLow-speed airfoil data Unclassified - UnlimitedReynolds number effectsExperimental-theoretical comparisonSTAR Category 0119. Security Classif. (of this report) 20. Security Classif. (of t

8、his page) 21. No. of Pages 22. Price*Unclassified Unclassified 74 $4.25For sale by the National Technical Information Service, Springfield, Virginia 22151Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EXPERIMENTAL AND THEORETICALLOW-SPEED AERODYNAMI

9、C CHARACTERISTICS OF THENACA 651-213, a = 0.50, AIRFOILBy William D. Beasley and Robert J. McGheeLangley Research CenterSUMMARYAn investigation was conducted in the Langley low-turbulence pressure tunnelto determine the low-speed two-dimensional aerodynamic characteristics of theNACA 651-213, a = 0.

10、50, airfoil. The results are compared with data from anotherlow-speed wind tunnel and also with theoretical predictions obtained by using a subsonicviscous method. The tests were conducted over a Mach number range from 0.10 to 0.36and an angle-of-attack range from -100 to 200. Reynolds numbers, base

11、d on the airfoilchord, were varied from about 3.0 x 106 to 23.0 x 106.The results of the investigation showed that the maximum section lift coefficientat a constant Mach number of 0.22 increased rapidly as Reynolds number increasedfrom about 3.0 x 106 to 9.0 x 106 and attained a value of about 1.7 a

12、t 9.0 x 106; furtherincreases in Reynolds number had only small effects on the maximum section lift coeffi-cient. The stall was abrupt below Reynolds numbers of about 9.0 x 106 and gradual athigher Reynolds numbers. The application of a narrow roughness strip near the leadingedge resulted in only sm

13、all effects on the lift characteristics at a Reynolds number ofabout 6.0 x 106, whereas extensive roughness wrapped around the leading edge forwardof 5-percent chord resulted in a decrease in maximum section lift coefficient of about13 percent. Increasing the Mach number at a constant Reynolds numbe

14、r of about 6.0 x 106was found to have large effects on the maximum section lift coefficient as a result of theflow over the airfoil becoming supercritical and the maximum section lift coefficientdecreased about 30 percent when the Mach number was increased from 0.10 to 0.36. Sec-tion lift and pitchi

15、ng-moment coefficients obtained at low Reynolds numbers for the smoothairfoil were in good agreement with results from another low-speed wind tunnel; however,there were differences in drag coefficients in the lift coefficient range where the laminarbucket would be expected. Comparisons of experiment

16、al section lift coefficients, pitching-moment coefficients, and chordwise pressure distributions with those calculated from aviscous flow theoretical method were good as long as no boundary-layer flow separationwas present; however, the theoretically calculated drag coefficients were generally lesst

17、han the experimental drag coefficients.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTIONResearch on both advanced technology and conventional airfoils has received con-siderable attention over the last several years at the Langley Research

18、 Center. Partic-ular emphasis has been placed on obtaining data at high Reynolds numbers to study theshock-wave boundary-layer interaction phenomena and to compare results measured invarious ground test facilities. The present investigation was conducted to obtain thebasic low-speed two-dimensional

19、aerodynamic characteristics of the NACA 651-213 air-foil over a broad range of Reynolds numbers. In addition, the experimental results havebeen compared with theoretical data obtained by using a viscous subsonic predictionmethod. The NACA 651-213 airfoil was selected to be representative of conventi

20、onalairfoils because of the existence of flight data and wind-tunnel data obtained in variousresearch facilities.The investigation was performed in the Langley low-turbulence pressure tunnelover a Mach number range from 0.10 to 0.36. The Reynolds number, based on airfoilchord, varied from about 3.0

21、x 106 to 23.0 x 106. The geometrical angle of attack variedfrom about -100 to 200. The operational characteristics and a new calibration of thetunnel are presented in an appendix.SYMBOLSValues are given in both SI and the U.S. Customary Units. The measurements andcalculations were made in the U.S. C

22、ustomary Units.a mean-line designationPL-PCp pressure coefficient,qoCp,critical critical pressure coefficient equivalent to a local Mach number of unityc airfoil chord, cm (in.)cc section chord-force coefficient,CP d() - CP d)Forward(t/c)max Cp Aft(t/)maCd section profile-drag coefficient, Wake d(h)

23、cd point drag coefficient, 2 11 12 P“ /2 q2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-cI section lift coefficient, Cn cos a - cc sin acm section pitching-moment coefficient about quarter-chord pointCp (0.25 - ) d) - Cp 0.25 - dc, section normal-

24、force coefficient, s. C d() -pu.s. d()h vertical distance in wake profile, cm (in.)M free-stream Mach numberp static pressure, N/m2 (lb/ft2)q dynamic pressure, N/m2 (lb/ft2)R Reynolds number based on free-stream conditions and airfoil chordt airfoil thickness, cm (in.)x airfoil abscissa (see fig. 1)

25、, cm (in.)z airfoil ordinate (see fig. 1), cm (in.)a geometric angle of attack, degp density, kg/m3 (slugs/ft3)Subscripts:L local point on airfoilmax maximummin minimum1 tunnel station 1 chord length downstream of model3Provided by IHSNot for ResaleNo reproduction or networking permitted without lic

26、ense from IHS-,-,-2 tunnel station downstream of model where static pressure is equal to free-stream static pressure and total pressure is assumed equal to totalpressure at station 1undisturbed stream conditionsAbbreviations:1.s. lower surfaceu.s. upper surfaceMODEL, APPARATUS, AND PROCEDUREModelThe

27、 development of the NACA 6 series airfoils is discussed in detail in reference 1.The NACA 651-213, a = 0.50, section was obtained by linearly increasing the basicthickness distribution of the NACA 651-012 airfoil and combining this thickness distri-bution with the a = 0.50 mean camber line for a des

28、ign lift coefficient of 0.20. Thenose radius was increased as the square of the thickness ratio. The airfoil sectionshape is shown in figure 1, and table I presents the measured airfoil coordinates.The airfoil model was machined from an aluminum billet and had a chord of60.63 cm (23.87 in.) and a sp

29、an of 91.44 cm (36 in.). The model was equipped with bothupper and lower surface midspan orifices located at the chord stations indicated intable II. Grooves were machined in the surface of the aluminum model and pressuretubing was routed in the grooves to desired orifice locations. The tubes were p

30、otted inplace with a plastic resin and orifices were drilled through the plastic into the tubing.The plastic was then machined to reform to the original surface and the airfoil surfacewas hand polished in the chordwise direction with number 400 silicon carbide paper toprovide a smooth aerodynamic fi

31、nish.Wind TunnelThe Langley low-turbulence pressure tunnel (ref. 2) is a closed-throat, single-return tunnel which can be operated at stagnation pressures from 1 to 10 atmosphereswith tunnel-empty test-section Mach numbers up to 0.42 and 0.22, respectively. Themaximum unit Reynolds number is about 4

32、9 x 106 per meter (15 x 106 per foot) at aMach number of about 0.22. The tunnel test section is 91.44 cm (3 ft) wide by228.6 cm (7.5 ft) high. Operational characteristics and results of a new calibration ofthe tunnel are included in the appendix.4Provided by IHSNot for ResaleNo reproduction or netwo

33、rking permitted without license from IHS-,-,-Hydraulically powered circular plates provided positioning and attachment for thetwo-dimensional model. The plates are 101.60 cm (40 in.) in diameter, rotate with theairfoil, and are flush with the tunnel wall. The airfoil ends were attached to rectangula

34、rmodel attachment plates (fig. 2) and the airfoil was mounted so that the center of rotationof the circular plates was at 0.25c on the model chord line. The air gaps at the tunnelwalls between the rectangular plates and circular plates were sealed with flexible slidingmetal seals, shown in figure 2.

35、Wake Survey RakeA fixed wake survey rake (fig. 3) at the model midspan was cantilever mountedfrom the tunnel sidewall and located 1 chord length behind the trailing edge of the airfoil.The wake rake utilized 91 total-pressure tubes, 0.1524 cm (0.060 in.) in diameter, andfive static-pressure tubes, 0

36、.3175 cm (0.125 in.) in diameter. The total-pressure tubeswere flattened to 0.1016 cm (0.040 in.) for 0.6096 cm (0.24 in.) from the tip of the tube.The static-pressure tubes each had four flush orifices drilled 900 apart and located 8 tubediameters from the tip of the tube and in the measurement pla

37、ne of the total-pressuretubes.InstrumentationMeasurements of the static pressures on the airfoil surfaces and the wake rakepressures were made by an automatic pressure-scanning system utilizing variable-capacitance-type precision transducers. Basic tunnel pressures were measured withprecision quartz

38、 manometers. Angle of attack was measured with a calibrated digitalshaft encoder operated by a pinion gear and rack attached to the circular plates. Datawere obtained by a high-speed data-acquisition system and recorded on magnetic tape.TESTS AND METHODSThe airfoil was tested at Mach numbers from 0.

39、10 to 0.36 over an angle-of-attackrange from about -100 to 200. Reynolds number based on the airfoil chord was variedfrom about 3.0 x 106 to 23.0 x 106.Most of the roughness data were obtained with standard NASA type strips locatedat x/c = 0.05 on both upper and lower surfaces. The strips were 0.25

40、cm (0.10 in.)wide over the airfoil span and the carborundum grains were sparsely spaced and attachedto the airfoil surface with clear lacquer. The roughness was sized according to refer-ence 3 and the size required for each Reynolds number is listed in table III. A limitedcomparison of roughness tec

41、hniques was made for the selected test condition ofM = 0.15 and R = 5.9 x 106 by utilizing a sparse distribution of number 60 grains in5Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-the forms of a strip and extensive roughness wrapped around the le

42、ading edge. The stripswere 0.25 cm (0.10 in.) wide and located at x/c = 0.04, whereas the wraparound extendedfrom the leading edge to x/c = 0.04 over both surfaces.For several test runs oil was spread over the airfoil upper and lower surfaces todetermine whether any local flow separation was present

43、. Tufts were attached to theairfoil and tunnel sidewalls with plastic tape to determine stall patterns on both the air-foil and adjacent tunnel sidewalls.The static-pressure measurements at the airfoil surface were reduced to standardpressure coefficients and machine integrated (based on the trapezo

44、idal method) to obtainsection normal-force and chord-force coefficients and section pitching-moment coeffi-cients about the quarter chord. Section profile-drag coefficient was computed from thewake-rake total and static pressures by the method reported in reference 4 and machineintegrated by use of

45、the trapezoidal method.An estimate of the standard low-speed wind-tunnel boundary corrections (ref. 5)amounted to about 2 percent of the measured coefficients and these corrections have notbeen applied to the data.PRESENTATION OF RESULTSThe results of this investigation are presented in the followin

46、g figures:FigureFlow-visualization photographs for NACA 651-213 airfoil .4Effect of Reynolds number on section characteristics. M = 0.22 . 5Variation of airfoil minimum upper surface pressure coefficient withReynolds number. M = 0.22; transition fixed at x/c = 0.05 . . 6Effect of various grit sizes

47、on section characteristics. M = 0.22;R = 5.9 x 106 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Effect of strip and wraparound roughness on section characteristics.M =0.15; R = 5.9x 106; no. 60 grit . 8Variation of maximum lift coefficient with Reynolds number. M = 0.22 . . . . . 9Variat

48、ion of minimum drag coefficient with Reynolds number. M = 0.22 . . 10Effect of Mach number on section characteristics and chordwise pressuredistributions. R = 5.9 x 106; transition fixed at x/c = 0.05 . . 11Variation of maximum lift coefficient and airfoil minimum upper surfacepressure coefficient w

49、ith Mach number. R = 5.9 x 106; transitionfixed at x/c = 0.05 . . 12Comparison of section characteristics for NACA 651-212 (a = 0.60) andNACA 651-213 (a = 0.50) airfoils. Models smooth; M _5 0.22 . 136Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FigureComparison of experimental and theoretical chordwise pres