1、NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS TECHNICAL NOTE No. 1677 EXFEEXMENTAL AND CALCULATED CXARACTERJSTICS bF SEVERAL HIGH-ASPECT-RATIO TAPERED WINGS INCORPORATING NACA 44-SERIES, 230-SERIES, AND LOW-DRAG 64-SEFUES AIRFOIL SECTIONS :; I By Thomas V. Bollech Langley Aeronautical LaboratQry Langl
2、ey Field, Va. . Washington September 1948 . L c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c-l NACA TN No. 1677 -fiAmDCmmCmCTERWICSOFSm G pressure tunnel to demonstrate the accuracy of the lifting-line theory in predicting the aerodynamic charac
3、teristics of unswept wixs with moderate to high aspect ratios and to determine the effects of vsria- tions in the geometric parameters of the wings on their aerodynamic characteristics. In the first phase of the investigation, reported in reference 1, seven unswept wings having NACA beries sections,
4、 aspect ratios of 8, 10, and 12, and taper ratios of?.5 and 3.5 were investi- gated to determine the effects of aspect ratio, taper ratio, and chord thickness ratio. In the final phase ofthe investigation, reported herein, two wings of NACA 2304eries and low4lrag 6keries sections were tested and the
5、 results are compsred with those of a wing of NACA kkeries sections and of the same plan form reported in refertince 1 to determine the effects of airfoil profile on the wing aerodynemic characteristics. AU. wings had sn aspect ratio of 10, taper ratio of2.5, and a root-chord thickness ratio of 0.20
6、. The experimental lift, drag, pitching-mame nt, and stalling charac- teristics of wings with aqd without leadingedge roughness for the flamutral and flap-deflected conditions are presented along with the calculated lift, drag, and pitching-moment characteristics of wings without leading-edge roughn
7、ess for the flap-neutral condition. The wing characteristics were calculated by the generalized method of the lifting-line theory, which allows the use of nonlinear section-lift curves. (See reference 2.) SYMBOIS The coefficients and symbols used herein are defined as follows: % lift-coefficient W 1
8、 EL increment of lift coefficient due to flaps . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1677 . 3 (Wmax R M Mach nmiber (V/a) a a P P S 9 . E drag coefficient (D/qS) Profile-drag coefficient (Do/) pitch-ant mean aerodynsmic chord
9、max3mm lift-drag ratio Reynolds number PVC 0 -r angle of attack of the wing root chord, degrees flap deflection, degrees slope of lift curve slope of pitching-mament curve lift, pounds wing profIle drag, pounds d- that is, the quertemhord line was perpendi- culartothe plase 8 was 35. The root-sectic
10、m and-tip-section thickness ratios were 20 and 12 percent; respectively, for all wings. The gecmetric charac- teristics of the test wings are presented in table I. The designatim for the wings is formed from nur6bers representing, consecutively, the taper ratio, aspect-ratio, NACA airfoil series, an
11、d root-section thick- ness in percent of wing chord. (See reference 1.) Et preps a lew roughness established by the Imgley tw spau, . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1677 5 respectively. me flaps were oonstructed of $-
12、Inch sheet steel which was attached to wooden blocks cut to the des 60, 80, and 90 percent of the chord and spaced 6 inches on the upper surface of the wing. Corrections for support tare smd interference have been applied to sll force-test data. Jet-boundary and air-flaw-misalinement correc- _ tions
13、 have been applfed to the angle of attack and drag coefficients. An additional tare drag correction has been applied to all drag data, which causes the drag characteristics of wing 2.510- thie effect- resulted in a divergence of the two drag polers. (See figs. 3 to 5.) For wing8 2.5-l-,20 and 2.104Z
14、30r20 this divergence occurred at=a lift coefficient ofapproximately 0.2, whereas for wing 2.+X3- however, after considera- tion that the discrepancy representa an increment in drag coefficient of approximately 0.0010, the correlation appears to be reasonable. Lift.- The calculated lift curves predi
15、cted quite accurately the angle of attack for maximum lift and the general shape of the experimental- lift curves throughout-the range from zero lift to beyond the stall. In . . - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RACA TR MO. 1677 7 no
16、case did the calculated maximum lift coefficients very more than 0.05 from the experimental values, with the average discrepancy being 0.03. (See table II.) The calculated and experimental values of the lift-Curve slopes for wings 2.X)- h this conditicm may be due to less adverse compressibility eff
17、ects than those which were encountered for the wing of EACA 23Gseries secticms. The application of leadwdge roughness (fig. 13) greatly reduced the value of maimm lift coefficient for all wings. The de em8n-b in trail-dge split flaps are thus indicated to be more effective for wings 2.51tack range.
18、In the vicinity of %?lsxJ separation occurred rather abruptly over the wing center secticm. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 CONCLUSIONS NACA TN NO. 1677 The aerodynamic chmacteristics of 88veral unswept tapered wings were determine
19、d by calculations using the method of NACA TX No. 1269 and by wind-tunnel tests to demonstrate the accuracy of calculations and to show the effect of airfoil section on the aerodynamic characteristics of unswept tapered wings. . The wings investigated were similar in plan form, had aspect ratio 10 e
20、nd taper ratio 2.5, and incorporated NACA 44-series, 23GSeries, and low-drag keries airfoil profiles. On the basis of compariecm at equal values of Reynolds nunber the following conclusicms were made: 1. The agreement obtained between the calculated and experimental cheracteristics was in most cases
21、 excellent. No definite trend existed within the scope of the investigation which would indicate that the degree of cornelation depends on airfoil Section. 2. The maximm lift coefficients obtained for the smooth wings with flaps neutral.were approximately equal. With flaps deflected md Smooth surfac
22、es, the highest value of maximum lift coefficient was obtained for the wing of NACA 23C-series sections. Because of the low flap effectiveness for the wing of NACA herlee sections, the maximum lift coefficient obtained for this wing was larer than that obtained for the wing of NACA of NACA 23a-Serie
23、s sections with the flaps neutral exhibited an abrupt stall, which may be unsatisfactory when stall warning or lateral stability at the stall is considered. The stall of the wings with F -tea - t -0.W -0.096 -.Mh -.OfB -.w -.OlO oulaalated zKpd=la o;ol% 0 t -*ceg -.COl .w -030 2.5404,20 3.0 -3.2 2.m
24、6-4,20 42.1 -2.0 e+loQp,20 -0.3 -0.5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Wiag 2.!3-lo-44,20 -3.2 4.8 31.4 . 23.7 0.00% o.ol.33 -0.096 -0.og2 0 0.02l 2.51064,20 -1.9 -1.8 P-1 24.0 .0070 .Q125 -.070 -.065 -.oll -.Ol.l 2.5-m30,20 -.s -.s 3.4
25、 23.3 .og8 .m8 -AC8 -.008 .WS -0% . =_ . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c I. . . , Figure l.- Layout of typical tapered wbg. (All dimensions in inches.) Provided by IHSNot for ResaleNo reproduction or networking permitted without l
26、icense from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. 1 * : . Figure 2.- High-aspect-ratio tapered wing mounted in the Langley 19-foot pressure tunnel. P v1 - Provided by IHSNot for ResaleNo reproduction or networking permitted withou
27、t license from IHS-,-,-. .* Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.4 .2 -.2 -.4 I?! ! 0 -02 .W .M .08 -8 -4 0 4 8 I2 I6 20 ./ 0 :I GT a cm (U CL agp1b CD a, and c rn Figure 3.- Experimental and calculated characteristics of wing 2.5-lo-44,2
28、0 with smooth leading edge. Flaps neutral. R = 3.49 x 106. E Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. Experimental (force tests)-+J 1 1 1 Experimental (wake survey) I, I Calculated A*, I I 1 I I I I I I I I I I IIIIIIIII -2 0 .2 .4 .6 08. LO
29、 M CL (b) C, against 0 CL. Figure 3.- Concluded, Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.6 Cl .4 . - 0 02 .04 .W -08 -8 -4 0 4 B 12 I6 20 24 ./ 0 -.I CD cp cm b) CL agrlDst CD, a, and c,. Figure 4.- Experimental and calculated characteristic
30、s of wing 2.5-10-64 R = 3.49 x 108. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.016 5 = 3.49 x 106. B Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,024 .0/6 cob .0/Z to ch Experimental (force
31、teds) Experimental (wake survey) :4 0 2 .4 -6 .8 I.0 f R = 3.49 X 106. Y Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . .6 I i t 1 I I t LPI I I I I I I I t t t t t 1 .08 J2 .I6 .20 -24 .28 A? -4 0 4 8 I2 I m -J 32 *3 6 a: Cm Figure 7.- Effect of airfoil section on the characteristics of wings with smooth leading edge, Partial-span flaps deflected COO; R = 3.49 x 106. L . . 9 . * Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
