1、.A-NASA TECHNICAl.MEMORANDUMNASA TM X-72843!X,ecZEFFECTS OF THICKNESS ON THEAERODYNAMIC CHARACTERISTICS OF ANINITIAL LOW-SPEED FAMILY OF AIRFOILSFOR GENERAL AVIATION APPLICATIONSBy Robert J. McGhee and William D. Beasley(NASA-TM-X-728_3) EFFECTS OF THICKNESS ONTHE AERODYNAMIC CHARACTERISTICS OF ANIN
2、ITIAL LOW-S_ED FAMILY OF AIRFOILS FO_GENERAL AVIATION APPLICATIONS (NASA) 51 pHC AO4/MF A01 CSCi 01A G3/02N79-13000Unclas234 2a._wREPRODUCEDBYNATIONAL TECHNICALINFORMATION SERVICEU. S. DEPARTMENT OF COMMERCESPRINGFIELD. VA. 22161NATIONALAERONAUTICSANDSPACEADMINISTRATIONLANGLEYRESEARCHCENTER,.HAMTON,
3、VIRGINIA23665 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I:l“IT-_lr-4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-P1. Report No. 2. Government Acclmon No.!NASA TM X-7_84_4 Title and SubtitleEf
4、fects of Thickness on the Aerodynamic Character-istics of an Initial Low-Speed Family of Airfoilsfor General Aviation Applications7. Author(s)Robert J. McGhee and William D. Beasley12.Performing Orpni_tion Name and AdamNASA Langley Research CenterHampton, VA 23665S_mtiori_ AglmCy _mt and AddrmNation
5、al Aeronautics and Space AdministrationWashington, DC 205463. Recilm_nts Cat_log _W.5. Relict DateJune g766. Puforming Orpnizatioo Code8. Plrforming Orgmnizaticm Report No.10. Work Unit No.505-06-31-0211. Contract or Grant No.13. Type of Rq)ort and Period CoveredTechnical Memorandum14. Sponsoring Ag
6、ency Code15. _o_tarv Not_Special technicalat a later date.information release, planned for formal NASA publication16. AbClTa_Wind-tunnel tests have been conducted to determine the effects of airfoilthickness-ratio on the low-speed aerodynamic characteristics of an initialfamily of airfoils.; The fam
7、ily of airfoils are designated as NASA LS(1)-0413,0417, and 0421 airfoils. The results were compared with theoretical predictions Iobtained from a subsonic viscous method. The tests were conducted over a Machnumber range from O.lO to 0.28. Chord Reynolds numbers varied from about2.0 x lO6 to 9.0 x l
8、O6.17. Key W_ (Sug_sted by Author(t) ) (STAR _tegor y underliGeneral Aviation AircraftLow-Speed Airfoil SectionsReynolds Number EffectsThickness Ratio EffectsExperimental-Theoretical Comparisdntg. Secmitv Oauif. (of this report)Unclassified_. Secmitv C,kmf. (of this laMP)Unclassified“Available from
9、t The Nali_l Technical Infrmatin S_rvica Springfield21. NO. of Plges 22. Price-Virginia 22151 _ -Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,i,J_Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EF
10、FECTS OF THICKNESS ON THEAERODYNAMIC CHARACTERISTICS OF ANINITIAL LOW-SPEED FAMILY OF AIRFOILSFOR GENERAL AVIATION APPLICATIONSBy Robert J. McGhee and William D. BeasleyLangley Research CenterSUMMARYAn investigation was conducted in the Langley low-turbulence pressuretunnel to determine the effects
11、of airfoil thickness ratio on the aerodynamiccharacteristics of an initial family of airfoils. The results are comparedwith theoretical predictions obtained from a subsonic viscous method. Thetests were conducted over a Mach number range from about 0.I0 to 0.28 and aReynolds number range from about
12、2.0 x 106 to 9.0 x 106 . The geometric angleof attack varied from about -I0 to 22 o .The results of the investigation indicate that the 13-percent airfoilprovided the best performance for this thickness family of airfoils. At aReynolds number of 4.0 x 106 with fixed transition near the leading edge,
13、 themaximum lift-drag ratios were about I00, 80, and 60 for the 13, 17, and 21-percent airfoils. Increasing the airfoil thickness ratio resulted in an averageincrease in drag coefficient of about three counts (0.0003) for each percentincrease in thickness ratio at the design lift coefficient with fi
14、xed transi-tion near the leading edge. Maximum lift coefficients at a Mach number of0.15 and a Reynolds number of 6.0 x 106 decreased from about 2.0 to 1.8 asthe airfoil thickness ratio increased from 0.13 to 0.21. Stall character-istics were of the trailing-edge type for the airfoil family. Maximum
15、 liftcoefficient was generally insensitive to roughness, just sufficient to tripfProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-the boundary-layer, for the 13-percent airfoil but was progressively moresensitive with increasing thickness ratio. Maxim
16、um lift coefficients for thisthickness family were substantially greater than the older NACA airfoils ofcomparable thickness ratios. Comparisons of experimental section data withthe theoretical viscous method of NASA CR-2523 were good for the 13- and 17-percent airfoils, but were poor for the 21-per
17、cent airfoil.INTRODUCTIONResearch on advanced technology airfoils has received considerableattention over the last several years at the Langley Research Center. Refer-ences 1 and 2 report the results of 17- and 13-percent-thick airfoils designedfor light General Aviation airplanes. References 3 and
18、4 report the resultsof a Fowler flap system and spoiler effectiveness for the 17-percent-thickairfoil. This report presents the basic low-speed aerodynamic characteristicsof a 21-percent-thick airfoil derived from the 17-percent-thick airfoil ofreference I. In addition, this report discusses the eff
19、ects of varying airfoilthickness ratio for this initial family and indicates some of the limitationsin present analytical performance prediction methods.The investigation was performed in the Langley low-turbulence pressuretunnel over a Mach number range from 0.I0 to 0.28. The chord Reynolds numberv
20、aried from about 2.0 x 106 to 9.0 x 106 . The geometrical angle of attackvaried from about -I0 to 22 o .SYMBOLSValues are given in both SI and U.S. Customary Units.and calculations were made in the U.S. Customary Units.The measurements2Provided by IHSNot for ResaleNo reproduction or networking permi
21、tted without license from IHS-,-,-CpCCcc dc dC ICCmCnhI/dMPqRtXZZ cz tpressure coefficient, PL - P_%airfoil chord, centimeters (inches)section chord-force coefficient, SCp d(_)section profile-drag coefficient, ._d d(_)wakepoint drag coefficientsection lift coefficient, c n cos _ - c c sinliftcurve s
22、lope per degreesection pitching-moment coefficient about quarter-chord point,section normal-force coefficient, -/Cp d(_)vertical distance in wake profile, centimeters (inches)section lift-drag ratio, cl/c dfree-stream Mach numberstatic pressure, N/m 2 (Ib/ft 2)dynamic pressure, N/m 2 (Ib/ft 2)Reynol
23、ds number based on free-stream conditions and airfoil chordairfoil thickness, centimeters (inches)airfoil abscissa, centimeters (inches)airfoil ordinate, centimeters (inches)mean line ordinate, centimeters (inches)mean thickness, centimeters (inches)geometric angle of attack, degrees3Provided by IHS
24、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-Subscripts:Lmax0Oolocal point on airfoilmaximumconditions at _ = 0free-stream conditionsAIRFOIL DESIGN AND DESIGNATIONThis airfoil family was obtained by linearly scaling the mean thicknessdistribution of the 17-perce
25、nt-computer-designed airfoil of reference I. Thus,all three airfoils have the same camber distribution and the design liftcoefficient is 0.40. This method of obtaining the airfoil family was selectedfor two reasons: to determine the performance of a scaled family of airfoilsand to validate the subso
26、nic viscous method of reference 5 for a range ofthickness ratios for aft cambered airfoils. The airfoil section shapes areshown in figure l, and figure 2 shows the mean camber line and mean thicknessdistributions. Tables I, II, and III present the airfoil coordinates.This initial family of airfoils
27、are designated in the form LS(1)-XXXX.LS(1) indicates low-speed (Ist series), the next two digits are equal to theairfoil design lift coefficient in tenths, and the last two digits are equalto the airfoil thickness in percent chord. Thus the GA(W)-I airfoil (ref. l)becomes LS(1)-0417 and the GA(W)-2
28、 airfoil (ref. 2) becomes LS(1)-0413. The21-percent-thick airfoil of this family is designated as LS(1)-0421.MODELS, APPARATUS, AND PROCEDUREModelsThe airfoil models were constructed utilizing a metal core around whichplastic fill and two thin layers of fiberglass were used to form the contour4Provi
29、ded by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-of the airfoils. The models had chords of 61 cm (24 in.) and spans of91.44 cm (36 in.). The models were equipped with both upper and lower surfaceorifices located 5.08 cm (2 in.) off themdspan. The airfoil s
30、urface wassanded in the chordwise direction with number 400 dry silicon carbide paperto provide a smooth aerodynamic finish. The model contour accuracy was generallywithin +.lO mm (.004 in.).Wind TunnelThe Langley low-turbulence pressure tunnel (ref. 6) is a closed-throat,single-return tunnel which
31、can be operated at stagnation pressures from l tolO atmospheres with tunnel-empty test section Mach numbers up to 0.42 and 0.22,respectively. The maximum unit Reynolds number is about 49 x lO6 per meter(15 x lO6 per foot) at a Mach number of about 0.22. The tunnel test sectionis 91.44 cm (3 ft) wide
32、 by 228.6 (7.5 ft) high.Hydraulically actuated circular plates provided positioning and attach-ment for the two-dimensional model. The plates are lOl.60 cm (40 in.) indiameter, rotate with the airfoil, and are flush with the tunnel wall. Theairfoil ends were attached to rectangular model attachment
33、plates (fig. 3) andthe airfoil was mounted so that the center of rotation of the circular plateswas at 0.25c on the model reference line. The air gaps at the tunnel wallsbetween the rectangular plates and the circular plates were sealed with flex-ible sliding metal seals, shown in figure 3.Wake Surv
34、ey RakeA fixed wake survey rake (fig. 4) at the model midspan was cantilevermounted from the tunnel sidewall and located one chord length behind thetrailing edge of the airfoil. The wake rake utilized total-pressure tubes,5Provided by IHSNot for ResaleNo reproduction or networking permitted without
35、license from IHS-,-,-0.1524 cm (0.060 in.) in diameter, and static-pressure tubes, 0.3175 cm(0.125 in.) in diameter. The total-pressure tubes were flattened to O.lOl6 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 a
36、part and located 8tube diameters from the tip of the tube and in the measurement plane of thetotal-pressure tubes.InstrumentationMeasurements of the static pressures on the airfoil surfaces and the wakerake pressures were made by an automatic pressure-scanning system utilizingvariable-capacitance-ty
37、pe precision transducers. Basic tunnel pressures weremeasured with precision quartz manometers. Angle of attack was measured witha calibrated digital shaft encoder operated by a pinion gear and rack attachedto the circular model attachment plates. Data were obtained by a high-speedacquisition system
38、 and recorded on magnetic tape.TESTS AND METHODSThe 0421 airfoil was tested at Mach numbers from O.lO to 0.28 over anangle-of-attack range from about -lO to 220 . Reynolds number based on theairfoil chord was varied from about 2.0 x lO6 to 9.0 x lO6. The airfoil wastested both smooth (natural transi
39、tion) and with roughness located on bothupper and lower surfaces at 0.075c. The roughness was sized for each Reynoldsnumber according to reference 7. The roughness consisted of granular-typestrips 0.127 cm (0.05 in.) wide, sparely distributed, and attached to theairfoil surface with clear lacquer.Th
40、e static-pressure measurements at the airfoil surface were reduced tostandard pressure coefficients and machine integrated to obtain-section normal-force and chord-force coefficients and section pitching-mement coefficients6Provided by IHSNot for ResaleNo reproduction or networking permitted without
41、 license from IHS-,-,-about the quarter chord. Section profile-drag coefficient was computed fromthe wake-rake total and static pressures by the method reported in reference 8.An estimate of the standard low-speed wind-tunnel boundary corrections(ref. 9) amounted to a maximum of about 2 percent of t
42、he measured coefficientsand these corrections have not been applied to the data.Testing of airfoil 0421 utilized high precision lower-range transducersto measure the wake-rake total pressures compared to the earlier testing ofairfoils 0417 (ref. l) and 0413 (ref. 2). These new transducers indicated
43、asmall difference in total pressure outside of the wake compared to the tunneltotal pressure. Accounting for this pressure difference resulted in a minorchange in the drag data reported in references l and 2. This drag adjustmenthas been applied to the most pertinent data of references l and 2 and i
44、ncludedin this report.PRESENTATION OF DATAFigureSection characteristics for LS(1)-0413 airfoil 5Section characteristics for LS(1)-0417 airfoil 6Effect of Reynolds number on section characteristics forLS(1)-0421 airfoil . 7Effect of Mach number on section characteristics forLS(1)-0421 airfoil . 8Effe
45、ct of Reynolds number on the chordwise pressuredistributions for LS(1)-0421 airfoil 9Comparison of chordwise pressure distributions forLS(1) thickness family of airfoils . lO7Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FigureEffect of thickness r
46、atio on section characteristics forLS(1) thickness family of airfoils . IIVariation of maximum lift coefficient with Reynolds numberfor LS(1) thickness family of airfoils . 12Variation of maximum lift coefficient with Mach number forLS(1) thickness family of airfoils . 13Variation of drag coefficien
47、t with Reynolds number forLS(1) thickness family of airfoils . 14Variation of lift-drag ratio with lift coefficient forLS(1) thickness family of airfoils . 15Comparison of maximum lift coefficient for present LS(1)thickness family with HACA airfoils 16Comparison of experimental and theoretical secti
48、oncharacteristics for LS(1) thickness family of airfoils . 17DISCUSSIONLift.- The effects of thickness ratio on the lift characteristics aresummarized in figure II for H = 0.15, R = 4.0 x lO6, and transition fixed atx/c = 0.075. The thickness ratio appears to have little effect on ci,o andthe angle of zero lift for thickness ratios up to 0.17. However, for athickness ratio of 0.21 the angle of zero lift increases by about 0.40o. Theangle of zero lift is largely determined by the airfoil camber, and thisincrease is attributed to viscous decambering
