1、NASA TECHNICALMEMORANDUMo,c_J!X=E(:,rrSA_14_X_72697) LOW-dP.IED _t:O,YN._NIC:I_FOIL STCTION DF._IG_;_,D _O._ G%P_L_V!lTlu_ AP_LICATiON5 (NaSn) 57 p HC 0_/M._01 CSCL 01_. ,.;3/0 ZLOW-SPEED AERODYNAMIC CH_-K_CTr_RISTICSOF A I3-PE_CENT-q!HIC_ AIRFOILSECTION DESIGNED FOR GE_“_AL AVIATIOLAPPLICATIONSBy R
2、obert J. McGhee, Willism D. Beasley, ar_ Dan M.NASA TM X-72697COPYHU_-_77-230_9_o_ersIThis informal documentation medium is used to provide accelerated orspecial release of technical information to selected users. The contentsmay not meet NASA forma_ editing and publication standards, may be re-vise
3、d, or may be incorporated in another pub!ication.“%,0, ,0, “I,“-_ ,_?, “ _ LANGLEYRESEARCHCFNTER,HAMPTGN,VIRGINIA23665 : ;.,j“Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LOW-SPEEDAERODYNAMIC CHARACTERISTICSOF A 13-PERCENT-THICK AIRFOILSECTION DES
4、IGNED FOR GENERAL AVIATIONAPPLICATIONSBy Robert J. McGhee, William D. Beasley, and Dan M. SomersLangley Research CenterSU_ARYAn investigation was conducted in the Langley low-turbulence pressuretunnel to determine the low-speed section cha_act_ristlcs of a 13-percent-thickairfoil designed for genera
5、l aviation applications. The results are comparedwith older NACA 12-percent-thick sections and with the 17-p_rcent-thick NASAGA(W)-l airfoil. The tests were conducted over a Mach number range from 0.0to 0.35 and an angle-of-attack range from -10 to 22. Chord Reynolds numberswere varied from about 2.
6、0 x l06 to 9.0 x l06.The results of the investigation indicate that maximum section lift co-efficients at a Mach number of 0.15 increased from about 1.7 to 2.1 as theReynolds number was increased from about 2.0 x l06 to 9.0 x l06. Stall chaT-acterlstlcs were generally gradual and of the trailing-edg
7、e type. The applica-tion of a n_rrow roughness strip near the l_adlng edge resulted in only _:ualleffects on the llft characteristics at a Reynolds number oz about 6.0 x l06,whereas extensive roughness wrapped aromd the leading-edge resulted in a de-crease in -_x_mum section lift coefficient of abou
8、t 19 percent. Increasing theMaeh number from 0.10 to 0.35 at _.constan _ Reynolds number of about 6.0 x l06decreased !_hemaximum section lift coefficient about 16 percent, with most ofthe decr,_sf ccurrlng above a Mach numb_r of abo_it 0.28. The 13-percent-thickProvided by IHSNot for ResaleNo reprod
9、uction or networking permitted without license from IHS-,-,-airfoil comparedto the 17-percent-thlck GA(W)-l airfoil, provided about a0.10 increase in m_xim_nsection lift coefficient, reduced the section profiledrag coefficients at all lift coefficients, and increased the section lift-drag ratio abou
10、t 22 percent at cruise and about 14 percent at climb. Maximumsection lift coefficient at a Reynolds numberof about 6.0 x l06 was about 16percent greater than the NACA23012airfoil section.INTRODUCTIONResearchon advancedtechnology airfoils has received considerable atten-tion over the last several yea
11、rs at the Langley Research Center. Reference 1reports the results of the NASAGA(W)-l airfoil, which was specifically de-signed for a twin-engine propeller driven light airplane. The achievement ofhigh performance of the GA(W)-l airfoil has _2omptedthe development of afamily of airfoila of differing
12、thickness and camber. This report presentsthe basic lov-speed aerodynamic characteristics of a 13-p_rcent-thick airfoilderived from the GA(W)-l airfoil. This airfoil has been d_signated as GeneralAviation (_nitcomb)-number two airfoil. (GA(W)-2).The investigation was performed in the Langley low-tur
13、bulence pressuretunnel over a Machnumberrsnge from 0.10 to 0.35. The chord Reynolds nu_obervaried from about 2.0 x 106 to 9.0 x l06. The g_ometrical augle of attackvaried from about -10 to 22.SYMBOLSValues are given in both SI and the U.S. CustomaryUnits. Themeasure-ments and calculations ?_eremade
14、in the U.S_ CustomaryUnlta.C pressure coefficient, PL “ P_P qc airfoil chord, centimeters (inches_2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Cnh“,/d% sectiono,or_/_)D (_)cd section profile-drag coefficient, . d dwake!c d point drag coefficient
15、(_-ef. 2)c section lift coefficient, c cos a - c sinI n cc section pitching-moment coefficient about quarter-chord point,ms%section normal-force coefficient, Pvertical distance in wake profile, centimeters (inches)section llft-drag ratio, ci/ cdN free-stream Mach numberp static pressure, N/m 2 _lb/f
16、t 2)q dynamic pressure, N/m 2 (ib/ft 2)R Reynolds number based on free-stream conditions_and airfoil chordt airfoil thickness, centimeters (inches)x airfoil abscissa, centimeters (inches)z airfoil ordinate, centimeters (inches)z mean line ordinate, centimeters (inches)Czt mean thickness, centimeter_
17、 (inches)a geometric angle of attack, degreesSubscripts:L local point on airfoilmaximumfree-strewn conditions3Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MODEL, .MPPARATUS, AND PROCEDUREModelThe 13-percent airfoil section (fig. i) was obtained by
18、 linearly decreas-ing the mean thickness distribution of the 17-percent GA(W_-l airfoil by 0.765(l_) and this thickness distribution with the camber line of thecombining meanGA(W)-l airfoil. This method of obtaining the airfoil family was selectedafter theoretical analysis, using a subsonic viscous
19、method, showed that theresulting airfoils should have similar characteristics to the GA(_I)-l airfoil.The mean camber and thickness distribution are shown in figure 2 and table Ipresents the measured airfoil coordinates.The airfoil model was constructed utilizing a metal core around whichplastic fil
20、l and two thin layers of fiberglass was used to form the contour ofthe airfoil. The model had a chord of 61.01 cm (24.02 in.) and a span of91.44 cm (36 in.). The model was equipped with both upper and lover surfaceorifices located 5.08 cm (2 in.) off the midspan and at the chord stationsindicated in
21、 table II. The airfoil surface was sanded in the chordwise direc-tion with number 400 dry silicon carbide paper to provide a smooth aerodynamicfinish. Figure 3 shows a photograph of the model.Wind _nnelThe Langley low-turbulence pressure tunnel (ref. 3) is a closed-throat,single-return tunnel which
22、can b_operated at stagnation pressures from 1 tol0 atmospheres with tunnel-empty test section Mach numbers up to 0.42 and 0.22,respectively. The maximum un_t Reynolds number is about 49 x l06 per meter(15 x 106 per foot) at a Math number of about 0.22. The tunnel test sectionis 91.44 cm (3 ft) wide
23、by 228.6 (7.5 ft) high.4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Hydr_alically actuated circular plates provided positioning and attach-ment for the two-dimensional model. The plates are 101.60 cm (_0 in.) indiameter, rotate with the airfoil,
24、and are flush w_th the tunnel wall. Theairfoil e_Is were attached to rectangular _odel attachment plates (fig. _) andthe airfoil was mounted so that the center of rotstion of the circular plateswas at 0.25c on the model reference line. Tho ai_“ ga_ at the tunnel wallsbetween the rectangular plates a
25、nd the circular plates were sealed with flax-ible sliding metal seals, shown in figure _Wake Survey RakeA fixed wake survey rake (fig. 5) at the model mldspan was cantilevermounted from the tunnel sidewall and located one chord length behind thetrailing edge of the airfoil. The wake rake utilized 91
26、 total-pressure tubes,0.152h cm (0.060 in.) in diameter, and six static-pressure tubes, 0.3175 cm(0.125 in.) in diameter. The total-pressure tubes were flattened to 0.1016 cm(0.0h0 in.) for 0.6096 cm (0.2h in.) from the tip of the tube. The static-pressure tubes each had four flush _,r_fices drilled
27、 90 apart and located 8tube diameters from the tip of the tube and in the measura_nent plane of thetotal-pressure tubes.InstrumentationMeasurements of the static presm_es on the airfoil surfaces and the wakerake pressures were made by an automatic pressure-scanning sy_t_mutillzingvariable-capacitanc
28、e-t_rpe precision transducers. Basic tunnel pressures _eremeasured with precision quartz manometers. Angle of attack _as measured witha calibrated digital shaft encoder operated by a pinion gear and rack attachedto the circular model attachment plates. Data were obtained by a hlgh-speedacquisition s
29、ystem and recorded on magnetic tape.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TESTSANDMETHODSThe airfoil was tested at Machntm_ers from 0.i0 to 0.35 over an angle-of-att_ck_ange from about -10e to 22. Reynolds numberbased on the airfoilchord _s
30、.s varied from about 2.0 x l06 to 9.0 x l06. The airfoil was testedboth smooth _natural transition) and with roughness located on both upper andlower surfaces at 0.075c. The roughness was sized for each Reynolds numberaccording_tn_r_fmr_nca_ The roughness consisted of granular-type strips0_127cm (0.
31、05 in.) wide, sparely distributed, and attached to the airfoilsurface with clear lacquer. At a Reynolds numberof 6.0 x l0 and a Nachnumberof 0.15_the NACAstandard roughness (number60 grains wrapped aroundleading edge on both surfaces back to 0.08c) was employedso that comparisonswith older NACAairfo
32、il data could be made. Eor_savezal_te_truns oil wasspread over the airfoil upper surface to determine if any local flow separationwas present. Tufts were attached to the airfoil and tunnel sidewalls withplastic tape to determine stall patterns.The static-pressure measurementsat the airfoil surface w
33、ere reduced tostandard pressure coefficients and machine integrated to obtain section normal-force and chord-force coefficients and section pitching-moment coefficientsabout the quarter chord. Section profile-drag coefficient was computedfromthe wake-rake total and static pressures by the method rep
34、orted in reference 2.An estimate of the standard low-speed wind-tunnel boundary corrections(ref. 5) amounted to a maximum of about 2 percent of the measured coefficientsand these corrections have not been applied to the data, except for the datashown in figure 18(a).6Provided by IHSNot for ResaleNo
35、reproduction or networking permitted without license from IHS-,-,-PRESENTATION OF DATA_igureEffect of Reynolds number on airfoil section characteristics.M = 0.15 6Effect of roughness configuration on airfoil section character-istics. M = 0.15_ R _ 6.0 x l06 . 7Effect of Mach number on airfoil sectio
36、n characteristics.R _ 6.0 x 106; airfoil smooth . 8Effect of Mach number on airfoil section characteristics.R _ 6.0 x 106; transition fixed at x/c = 0.075 . 9Comparison of the section characteristics for the GA(W)-l andGA(W)-2 airfoils. M = 0.15; transition fixed at x/c = 0.075 l0Typical chordwise p
37、ressure distributions for GA(W)-2 airfoil.M = 0.15; R = 3.0 x 106; airfoil smooth . llComparison of the chordwlse pressure distributions for the GA_WI-1and GA_W)-2 airfoils. M = 0.15; R = 4.3 x 106; transition fixedat x/c = 0.075 . 12Variation of maximum section lift coefficient with Reynolds number
38、for GA_W)-l and GA(W)-2 airfoils. M = 0.15 13Variation of maximum section lift coefficient w_th Mach number forGA(W)-l and GA_W)-2 airfoils. R _ 6.0 x 106; airfoils smooth . . . l_Variation of drag coefficient with Reynolds number for GA(W)-l andGA(W)-2 airfoils. M = 0.15; transition fixed at x/c =
39、0.075 15Variation of lift-drag ratio with Reynolds number for GA_W)-l andGA(W)-2 airfoils. M = 0.15_ transition fixed at x/c = 0.075 167Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FigureVariation of maximum section llft coefficient with Reynolds
40、numberfor various airfoils. N = 0.15; airfoils smooth 17Comparison of section characteristics of NASA GA_W)-2 airfoil andNACA 4_12, 23012, and 651-_12 airfoils. M = 0.15; R _ 6.0 x 106;wraparound roughness to 0.08c surface length (no. 60 grit) . 18Comparison of section characteristics of NASA GA(W)-
41、2 and NACA651-213 airfoils. M = 0.15; R _ 6.0 x 106; strip roughness 19Comparison of experimental and theoretical section character-istics for the GA(W_-2 airfoil. M = 0.15; R = B.0 x 106;transition fixed at x/c = 0.075 20DISCUSSIONLift.- Figure 6 shows that with the GA(W)-2 airfoil smooth _naturalb
42、oundary-layer transition) a lift-curve slope of about 0.11 per degree (un-corrected for wall boundary effects) and a lift coefficient of about 0.49 ata = 0 was obtained for the Reynolds numbers investigated, CM = 0.15). Maxi-mum lift coefficients (fig. 17) increased almost linearly with increasingRe
43、ynolds number and obtained values of about 1.7 at R = 2.1 x l06 and about2.1 at R = 9.0 x l06. The airfoil section exhibits a gradual type stall(fig. 6), particularly e,t the lower Reynolds numbers. Tuft pictures (notshown) and the pressure data of figure ll indicated that the stall is of theturbule
44、nt or trailing-edge type.The addition of a roughness strip at .075c (fig. 6) altered the llftcharacteristics because of changes in boundary-layer thickness, particularlyat the lower test Reynolds numbers. For example, at R = 2.1 x l0 g (fig. 6(a)the angle of attack for zero llft coefficient changed
45、from about -4.1 to -3,8 ,8Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-and the li_ coefficient at a = 0 decreased from about 0.49 to 0._. Theseeffects on the llft cl_racterlstics decreased as the Reynolds number was In-creased and were essentially
46、 eliminated at R = 9_h x 106 (fig. 6_e)l. FigureiB shows that the-roughness strip had only minor effects on the airfoilsmaximum llft coefficients for the Reynolds number range tested. A comparisonof the lift data obtaln_d with a roughness strip (.number 100 grit; sized forR = 6.0 x l06) and with _xt
47、enslve roughness (no. 60 grit) wrapped around theleading-edge is shown in figure 7(a) for R _ 6.0 x l06. A decrease in theangle of attack for Cl,m_ x of about 5 and a decrease of about 19 percent inc is shown for the wraparound roughness.l ,m_XThe effects of Mach number on the airfoil lift character
48、istics at aReynolds number of R _ 6.0 x l06 are shown in figure 8(a) for the smooth air-foil and in figure 9(a) for the airfoil with a roughness strip located atx/c = 0.075. The expected Prandtl-Gauert increase in lift-curve slope isindicated by increasing the Mach number from 0.10 to 0.35. This same Machnumber increase, however (figs. 8(a) and 9(a) resulted in a decrease in thestall angle of attack of about 6 and about a 16 percent decrease in Cl,max.Figure l_ shows that most of this decrease in Cl,ma x (about 12 percent)occurred above M = 0.28. These Mach number effects are
