1、b RESEARCH MEMORANDUM WIND-TUNNEL INVESmGATION AT HIGH SUBSOMC SPEEDS OF THE LATERAL-CONTROL Cl3X3WTE3Eus?cS OF AN AILERON AND A STEPPED SPOILER ON A WING WITH L;EADING EDGE SWEPT BACK 51.3 BY Leslie E. Schneiter ami John R. Hagerman Langley Aeronautical Laboratory Langley Air Force Base, Ta.“ “ “ “
2、 “ I . . “ . “ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON - June 7,1949 !JNCLASSI F1 ED Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c . . A wind-the1 investigation ha8 been made through a sped range frm a Mach number of 0.30 to a Mach
3、 number of approximately 0.90 to determine the Lateral-control characterietice-of a 20-percent-chord by 39-percent- semispan aileron and a 60-percent-semispan stepped spoiler on a semispan- wing model with aspect ratio of 3.06 having 31.3O .sweepback of the wing leading edge. In addition, the aerody
4、namc characteristice of the plain wing were determFned through the speed range. The aileron rolling effectiveness decreased BB the Mach number increased; wherejectlan through the upeed range and at wing angles of attack frcm appror-Itely -4O to 4. The charac- IxriHtics of the mare satisfactory of th
5、e two spoiler canfigurations were determined through a range of ppoiler projectiom. “ “ Tho teats were mado in the Langley high-speed 7- by 10-foot tunnel. llEFINITIONS AND SYMBOLS The forces and moments on the wing are presented about the wind axe8, which far the conditions of these tosts (zoro yaw
6、) correspond to the atability axes. (See fig. 1.) The axes intersect at a polnt 26.6 inches rearward of the lea- eQe of the wing root on the line of interection of the plane of syrmnotry and the chord plane of the mo.del, as shawn in figure 2. This corresponds to a point 26.2-percent chord rearward
7、of the loading edge of the wing mean aerodynamic chord, ae also sham in figure 2. -. “ “ The rolling-mcanent and yawing-mcanent coefficients determined on the semispan wing represent the aerodynermic effecte that occur on a camflete wing a8 a roault of deflection of the aileron or projection of the
8、apoilor on only one sdopan of the cmplete wing. The lift, drag, and pitching- mwnt coofficients dotormined on the semispan wing (with the ailoron or upoilor neutral) roprooent those that occur on a cnmplete wing. - CL CD cm c2 The oymbols used in the presentation of results are as folloml: lift coof
9、ficiont drag. coof f ?c.i?nt rolling-mamont coefficient ( these two spoiler arrangements are hereinafter referred to as spoiler configuratian 1 and configuration 2, respectively. A sketch of spoiler configuration 1 is shm in figure 4. b b b 2 The Mach number range for the tests wae fram about M = 0.
10、30 to about M = 0-91, which corresponds to a Re;gnolds nlrmber range *can +q = 4.22 X lo6 to Rm = 9-34 X 106 based on a mean aerodm-c chord length of 2.087 feet. The variation of Rqnolds nmber with Mach number is shorn in figure 5. Wing angle-of-attack tests with the aileron at Sq = Oo were made at
11、various constant Mach numbers through an angle-of-attack range Frau approximately -bo to wing stall at M = 0.30 and to approximately 80 at all other Mach numbers. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 - NACA RM 906 constant projectiom. Sp
12、oiler canfiguration 1 was investigated at projections -of -E, -1, -2, -3, -5, aTld -7 percent chord; whereas spoiler configuration 2 w-aa investigated only at a projection of -7 percent chord. (Spoiler projection is negative when spoiler projects above wing upper surface. ) 1 The remlts of the inves
13、tigation are presented In figures 6 to ll. Wing Aerodynamic Characteristics Lift characteristics.- The om88 of lift coefficient against wing angle of attack for all Mach numbers (shown in fig. 6 were linear through the low angle-of -attack range. At a Mach number of 0.30 (the only Mach number .at wh
14、ich data wer0 obtained at hi) BB shown in figure 6. For the angle-of-attack range . wherein Q increased slightly, the stability of the wing, as indfcated by the slope of the curve of pitchinglncanant coefficient against lift coefficient aC whereas, at any constant value of total equal up-and-dam ail
15、eron deflection, the roUlng-mmnt coefficient decreased with increasing Mach ntndber. This effect. of Mach number an the rolling effectiveness of the controls was of such -tude that with the wing at a = 00 the total aileron deflection required to produce a rolling-moruent coefficient equal to that pr
16、oduced by the spoiler at its maximum projection (-0.07) Increased from 1 at a Mach nmber of 0.30 to 300 at a Mach number of 0.85. The spoiler produced favorable yanlng mamSnts as canpared to the generally unfavmable yawhg manents produced by the aileron. This effect, in conjunction with the nw large
17、 negative values of the stability parameter cz (rolling maanent due to sideslip) for a swept wing, will increase the rolling effectiveness of the spoiler and decrease the rolling effectiveness of the aileron. B coNcuTsIoNs The results of an inveatiga-l;ion at high speeb of a semispan-wing model with
18、 an aspect ratio 3 and a leading edge swept back 31-30 to determine the wing aerodynamic characteristics and also the lateral- control characteristics of a partial-span aileron eLnd of a stepped spoiler lead to the follaring conclusions: 1- The slope of the curve of lift coefficie against wlng angle
19、 of attack % increased with increasing Mach nlDdber and the variation was in excellent agreament with the theoretical variation. 2. The wing longitudinal stability, as indicated by the slope of the curve of pitching-moment coefficient against lift coefficient m, increased slightly with increasing Ma
20、ch nmiber. - 3. The wing drag coefficient increased slightly with increasing Mach nmiber but the critlcal speed of the wing was not exceeded for any ccanbi- number investigated (0.91) . - nation of lift coefficient and Mach number, even at the highest Mach Provided by IHSNot for ResaleNo reproductio
21、n or networking permitted without license from IHS-,-,-10 - MACA RM LgDo6 4. The aileron rolling effectiveness decreaeed as the wing angle of attack and Mach number increased. 5- The total yarwingplaansnt coefficient resulting frm equal up-and- down deflection of the aileran was essentially unaffect
22、ed by variation of the Mach number, and wa8 generslly adverse (sign of yawing mcnnent opposite to sign of r-ciUiq mamant). 6. The hingeplmnant paramsters Ch and Ch (slope of the curvee of hingeplament coefficient with wing angle of attack and aileroq deflec- tion, respectively) were negative and var
23、ied sfmost negligibly with variation of the Mach number. % 7. The Spoiler rollingPMnent coefficient increased with increasing Mach number and wing angle of attack wit- the amll angle-of-attack range (-bo to bo) investigated. The increase of rollinglncnnent coeffi- cient with spoiler wojectian was ne
24、arly linear at wing angles of attack - of 00 and 40. 8. The yarwinglnamsnt coefficient resulting fram spoiler projection w88 favarable (ai of yawing maslent 8- 88 eign of rolling mament) , incremed with increasing spoiler projection, and tended to increase with increasing Mach number. Langley Aerona
25、utical Labaratmy Batianal Adviamy Canrmittee for Aeronautics Langley Air Force Baae, Va. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM 906 11 1. Schneiter, Lealie E., and Watson, James M. : Low-Speed Wina-Tunnel Jnvestigation of Varioua P
26、lain-Spoiler Configurations for Lateral Control on a Q0 Sweptback Wing. NACA TN 1646, 1948. 3. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, with Cmideration of the Effect of Cmpxmibility. NACA RM Ap28, 1947: 5. Lowry, John G., and Schneiter, Leslie E.
27、 : Estimation of Effectiveness of Flap-Type Controls on Sweptback Wings. NACA 1674, 1948 6. Langley Research Staff (Compiled by Thomas A. Toll) : Summary of Lateral-Control Research. NACA Rep. 868, 1947. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-
28、,-,-e Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, . . . I . . . . . . . . . . . . . . . . . W r . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c Provided by IHSNot for ResaleNo reproduction o
29、r networking permitted without license from IHS-,-,-. . . . . “. . . . . . . . . . . . . . . . . . . . . . . . . I I . . . . . . . . Y , I c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking pe
30、rmitted without license from IHS-,-,-I . r Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4.0 .3 .4 .5 .7 .8 .9 Mach number, # Figure 5.- Variation of ReynolU n-er with Mach nmber Repolder numbex+ is based m wing mean aerodynamic chord of 2.087 feet
31、 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-32 -8 ./ 0 Figure 6.- Wing aercdynamic Characteristics and. aileron hinge-mament characteristics with 6 = 0 for various Mach numbers. %ve Provided by IHSNot for ResaleNo reproduction or networking pe
32、rmitted without license from IHS-,-,-20 NACA RM igD06 -8 0 6. 16 24 32 Angle of affack,c, deg (a) Concluded. Figure 6.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 906 21 (b) Mach numbers from M = 0.829 to M = 0.913. Figure 6.-
33、 Continued. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 NACA RM 96 -8 0 9 (b) Concluded. FQure 6.- Concluded. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RACA RM gm6 .3 .5 .6 .7 .8 .9 F
34、igure 7.- Pxperlmsntal and theoretical variation of the wing llf“ve slope Q with Mach number. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 NACA RM 936 .3 T3 (a) a = lC.oo. Figure 8.- Variation of aileron lateral-control characteristics with Mac
35、h number for various. aileron deflect io-. L11 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM I Figure 8.- Continued. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-26 NACA RM 906 -3 Figur
36、e 8.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 906 (a) u = 8.2O. Figure 8.- Continued. “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACA RM gn06 I ? . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . .3 .4 .5 .B A“cA flurnber, A (f) a = 16.4 . Figure 8.- Concluded. 0 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
