1、L2CASE FILE_ f_ !“_ VNATIONAL ADVISORY COMMITTEEFOR AERONAUTICSTECHNICAL NOTENo. 1546AERODYNAMIC CHAI_CTER._TICS OF 24 NACA 16-SERIES AIRFOILSAT MACH NUMBERS BETWEEN 0.3 AND 0.8By W. F. Lindsey, D. B. Stevenson, and Bernard N. DaleyLangley Aeronautical LaboratoryLangley Field, Va.WashingtonSeptember
2、 1948Provided 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-,-,-NATIONAL ADVISORY C0_ITYEE FOR AERONAUTICSTECHNICAL N_fE NO. 1546AERODYNAMIC CHARACTERISTICS OF 24
3、NACA 16-SERIES AIRFOILSAT MACH NUMBERS BETWEEN 0.3 AND 0.8By W. F. Lindsey, D. B. Stevenson, and Bernard N. DaleySUMMARYAn investigation has been conducted to determine the aerodynamiccharacteristics of a group of NACA 16-series airfoils releted in camberand thickness over a Mach number range from 0
4、.3 to approximately 0.8.The results obtained from the present Invest_gation were combined withthe data of 12 NACA 16-serles airfoils obtained under the same con-ditions and previously reported in NACA Rep. No. 76-_. All thecurrently available force-test data for NACA 16-serles airfoilsobtained under
5、 the same test conditions in the Langley 2h-inch high-speed tunnel are presented.IBTRODUCTIONThe NACA 16-series airfoils were derived (referen0e I) for useat high speeds, particularly for propeller applications. The variationsin design camber and thickness ratio, covered in referenze l, were notof s
6、ufficient scope to meet all the requiremsnts of propeller design.A test program was formulated, therefore, whereby the aerodynamiccharacteristics would be obtained for some of the airfoils ofreference 1 over an extended angular range, as well as for 12 addi-tional airfoils of the same series. The re
7、sults of this investig_tloncombined with the data of ref3rencs 1 are presented herein uncorrectedfor tunnel-wall constriction effects. The magnitude of the constrictioneffect on Mach number at supercritical speeds is approximately 2 percentof the uncorrected value and does not affect the validity of
8、 the conclusions.SYMBOLSc_c_cz ianglo of atSack, degreessection llft coefficientdesign section lift coefficient (incompressible potential flow)Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACATN No. 1546c Zocdc c/4Z/dMMcrPoma xt/cxYsection lift
9、coefficient at M = 0 (experimental valueswere obtained by extrapolating from M = 0.3 to M = 0by Glauerts method)section drag coefficientsection pltchlng-moment coefficient about quarter-chordaxislift-drag ratiostream Mach numbercritical Mach number; Mach number at which speed of soundis attained loc
10、ally as on airfoilmaximum incompressible pressure coefficientthlckness-chord ratio, percentairfoil station, fractions of chordairfoil ordinate, fractions of chord measured normal tocamber lineAPPARAI“UB AND TESTSForce measurements of lift, drag, and pitching moment were madein the Langley 24-inch hi
11、gh-speed tunnel (described in reference 2)on a series of airfoils having NACA 16-series profiles. The thickness-chord ratios of the airfoils tested ranged from 6 to 30 percent and thedesign lift coefficients ranged from 0 to 1.0. The specific airfoilsfor which force meaeurem_nte were made in this in
12、vestigation are givenin table I and are differentiated from those airfoils reported inreference 1.The models were made of duralumin and had a chord Of 5 inches.Each model spanned the 24-inch test section and passed through holescut in flexible brass end plates that preserved the contour of thetunnel
13、 walls. The holes were the same shape as, but elightly largerthan, the model. The ends of the model were secured in a balance ofthe type described in reference 3.The llft, drag, and pitching-moment coefficients were measured atangles of attack corresponding at low Mach numbers to a llft-coefficientr
14、ange from 0 to approximately 1.O. These data were obtained for a MachProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-_ACA TN _o. 1!96 3number range from 0.3 to approximately 0.8. The c?rresponding Reynoldsnumber range extended approximately from 0.85
15、 lO6 to 2 x lO6. Dragcoefficients for several of the airfoils were obtained by the wake-survey method. The wake-survey measurements were generally limitedto an angle of attack of 0 or the design angle.Critical Mach numbers at low angles of attack were estimated bymeans of small total-pressure tubes
16、mounted on the upper surface ofthe airfoils. The tubes were generally located at the 75-percent-chordstation and 2 to 3 percent chord above the airfoil surface. The Machnumber at which the measured total pressure decreased approximately0.02 percent was taken as an estimate of the critical Mach numbe
17、r.NACA 16-SERIES AIRF01I_The NACA 16-series airfoils are designated by a five-digitnumber (except for the case in which the design lift coefficient isequal to or more than 1.0). The first digit represents the seriesclassification. The second digit indicates at design conditions thedistance in tenths
18、 of chord from the leading edge to the position ofmlnmum pressure. The third digit, first digit following the dash,indicates the amount of camber expressed in terms of design llftcoefficient in tenths. The last two digits together express thethickness in percent chord.The thickness distribution of t
19、he NACA 16-series airfoils wasdeveloped (reference i) to produce a shape having very low inducedvelocities and thus having high critical Mach numbers. The ordinatesfor the basic or symmetrical profile of the NACA 16-serles airfoilscan be obtained from the following equations:_I = OOl-tc(0“989665xlI/
20、2 - 0“239290xi - OO_lO00x12 - 0“559400x13_s_ndwhere y is the ordinate in fractions of the chord measured normalto the camber line and x is the station in fractions of the chord.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 NACATN No. 1546Subscrip
21、ts I and 2 pertain to the region ahead of and behind themaximumthickness location, respectively Ifor example, xI _ 0.5and r_ 5)The leading-edge radius expressed in percentage of the chord isL.E. radius = 0.0048972C_) 2The ordinates for a 9-percent-thick airfoil are presented in table II.The camber l
22、lne for the NACA 16-series airfoils was derived(reference l) to have essentially a uniform chordwlse loading. Thiscamber line, designated the a = 1 mean line in reference 4, can beexpressed in equation form asdYo _0.079577czi_ Idx = OgeX - loge(l - x_where Yc is the mean-llne ordinate in fractions o
23、f chord and x isthe station in fractions of the chord.The ordinates and slopes of the camber line for NACA 16-seriesairfoils are presented in table II. It may be noted that the slopeof the leading-edge radius as given in table II differs from that givenin the corresponding table of reference 1. Sinc
24、e the slope of theleadlng-edge radius is determined by the slope of the camber line, thevalue specified for the leading-edge radius depends upon the chordwisestation x at which the camber-line slope is obtained. The slopespecified in reference 1 corresponded to the 0-percent-chord stationand fairing
25、 difficulties were later encountered at the leading edgeof highly cambered thick airfoils. The slope specified in table IIcorresponds to the 0._-percent-chord station. The leading-edge radiusmust be as specified but the slope of the leading-edge radius does notappear to be rigidly fixed, probably as
26、 a result of the approximationsmade in determining the extremities of the camber line. (Seereference 1.)Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA _W No. 1_6 5PRECISIONThe errors to which the data were subject can be classified asaccidental
27、 andsystematic. The accidntal errors arose frominaccuracies in model installation, in calibrations of alr stream andbalance, and in the reduction of the test records. The accidentalerrors evaluated from an inspection of the test data are as follows:c Z +0.005cd . /9.0005Cmc/_ . +0.oo2_, deg . +0.iTh
28、e largest systematic error to whlch the data were subject arosefrom leakage of air through the (3/64 inch) gaps at the Junctures ofairfoils and tun_lel wall. The corrections were first determined bytests on an NACA 0012 airfoil with v_r_ous end gaps; these correctionswere applied to the data of refe
29、rence 1. Tests to determine the cor-rection were extended for the present Investigation to include not onlyairfoils of various thlckness-chord ratios but also a large angle-of-atteck range. The tests of the NACA 16-series airfoils cambered for adesign l_ft coefflc_ent of 0.3 and having thickness-cho
30、rd ratios from6 to 21 percent were made. The angle-of-attack r_nge corresponded tothe range presented herein for those airfoils.Investigation of the end-leakage effects showed that the correctionto drag coeff_clent depended not only on Much number and lift coeffi-cient but also on thlckmess-chord ra
31、tio. The investigation furtherindicated that camber mdght have an effect on the angle-of-attackcorrection since the tests on the cambered airfoils showed only ashift in angle of zero lift (see also fig. l0 of reference 2); whereasthe tests on a sy_mmtrical airfoil (reference l) indicated that thecor
32、rection was a function of lift coeff_clent. The maximum differencebetween the two corrections d_d not exceed 0.3 .The corrections for end leakage as determined by the tests onNACA 16-series airfoils were considered to be more reliable than thoseused in reference 1 because of the greater renge of the
33、 test. Thecorrections were applied not only to the 12 _ddltion_l eIrfoils butalso to the 12 airfoils reported _n reference 1. Differences in theaerodynamic characteristics for airfoils _n th_s paper emd _nreference 1 are primarily a result of the change in the end-leakagecorrections.A comparison of
34、dr_ coefficients obtained from wake surveys andfrom force-test data corrected for end leakage is sho_1 in figure 1.At low design llft coefficients the d_fferences are approximately ofProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACATN No. 1546th
35、e samemagnitude as the accidental errors. At high design llftcoefficients the differences increase and thereby indicate that theapplied leakage correction mlght be too small in that lift-coefflcientrange. The wake-surveydrag coefficients, however, do not increaseas rapidly with design lift coefficie
36、nt as would be expected,especially at high values of czi.The remaining systematic errors associated with these data arisefrom wlnd-tunnel wall interference. The correction for this effectas determined by the method of reference 5 maYbe subject to error atsupercritlcal Machnumbers. The msthod, howeve
37、r, should give anestimate of the magnitude of the errors Involved. An estimate ofthe error can be obtained from figure 2 which shows a comparisonbetween basic aerodynamic data for the NACA16-309 airfoil correctedand uncorrected for tunnel-wall effects. An examination of figure 2Indicates that the co
38、rrection _s generally small and, therefore, hasnot been applied to the results presented herein. A further exami-nation _ndicates that at the highest Machnumbers the correctionappears to have the greatest effect on Machnumber. Thus, in theapplication of the data in a design problem, the approximatel
39、y2-percent correction for Machnumber could have a large effect.The choking phenomenonIs an additional effect of tunnel wallsthat enters into the problem of w_nd-tunnel testing at high subsonicMach numbers. At the choking Mach numbers sonic velocities extendfrom model to tunnel wall, and the static p
40、ressure is lower behlnd themodel than In the undisturbed region upstream of the model; thus, largegradients in static pressure cen be produced. The resulting flow pastthe model is unlike any free-air condition, and data obtained at andnear choking Mach munbers are, therefore, of questionable value.
41、Datanear the choking Mach number _re, consequently, omitted from this paperwith the exception of a few conditions shown In flgure 3. The arrowsat the zero-lift axes (fig. 3) show the one-dimenslonal theoreticalchoklngMach number at design condltlons for each of the airfoils _n_dindicate how closely
42、these basic date approach the choking speed. Thedata at the higher angles of attack generally do not approach theirrespective choking Mach numbers as closely as do the data near designconditions.RESULTSThe variation _n aerodyn_zHc characteristics with Mach numbersat constant angles of attack for eac
43、h of the NACA 16-series airfoilstested constitutes the basic data for the investigation _nd is presentedin figure 3.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACATN No. 1546 7The data of figure 3 were cross-plotted to show the effects ofvariati
44、on of thickness-chord ratio and camber or design lift coeffi-cient on the aerodynamic characteristics. The cross-plotted resultsare showrl _n figures h to 9.The effect of csmberon the variation of llft-drag ratio withlift coefficient for airfoils of 6-percent, 9-percent, 12-percent,and l_-percent th
45、ickness is presented In figures i_ ii, 12, a_ 13,respectively. The effect of thickness on lift-drag ratio for cP_mberedairfoil having a design lift coefficient of 0.3 is presented infigure 14.The theoretically derived maxlmnmnegative pressure coefficients(reference 4) for NACA16-series airfoils are
46、presented in figure 15.Figtue 16 permits the max_mumnegativepressure coefficients in an_ncompresslble or low-speed flow to be transcribed to critical Ms_chnumbers in accordance with yon Ka_rm_n-Tsiensreletion (reference 6).A comparison between the theoretically derived and the experln_ntallydetermin
47、ed critical Machnumbers is given in figure 17. The n_suredcritical Machnumbersat several stations on an tultwlsted propellertip are comparedin figure 18 with the values estimated from two-dimensional flow.A comparison is madein figure 19 between the critical Maehnumber and the Machnumber for llft br
48、eak and maximumllft-drag ratio.The effects of thickness and camberon the Machnumbers for llft breskand drag rise are presented in figure 20.A tabulation of the airfoils investigated and the figure numberscontaining the pertinent test data are given in table I.DISCUSSIONSubcriticalLift.-The NACA16-se
49、ries airfoils at the design angle of attackof 0 and at low Machnumbersdo not produce experimental lift coeffi-cients of the samevalue as the design lift coefficients. The dif-ference can be attributed primarily to viscosity in the real flow andits modifying effect on the theoretically predicted influence of thecamber line, especially over the rear
