REG NACA-RM-A9L01-1950 Wing-tunnel investigation at Mach numbers from 0 50 to 1 29 of an all-movable triangular wing of aspect ratio 4 alone and with a body.pdf

1、RESEARCH MEMORANDUM WIND-TUNNEL INVESTIGATION AT MACH NUMBERS FROM 0.50 TO 1.29 OF AN ALL-MOVABLE TRIANGULAR WING OF ASPECT RATIO 4 ALONE AND WITH A BODY By LoF: IXIAE- mG OF ASPECTRKLIO4ALCNEMDWITRABCfDY By Louis S. Stivers, Jr., and Alexander W. Malick SUMMARY The aerodynsmic characteristics of an

2、 al whereas the latter may prescribe lifting surfaces of higher aspect - ratio. As a consequence, it appears that a compromise would confront the designer contemplating the use of an all-movable wing or ccntrol surface on supersonic aircraft. It is expected that the design of all.ovable lifting surf

3、aces will be dictated largely by information available from experimental inveeti- gations, yet only a small amount of such data exists. Recourse to theory does not necessarily lead to satisfactory design data. Further- more, in the transonic Mach number range the applicability of existing theory wou

4、ld be generally questionable. Inordertoprovide experimental data in the transonic Mach number range applicable to the design of trielan-form,all-movable lifting surfaces, an investigation has been made of an all-movable wirg in the Ames l-by +l/Sfoot high-speed windtunnel. The results of this invest

5、igation are presented herein for the wing alone and with a body at Mach numbers from 0.50 to about 0.98 and from 1.09 to 1.29. In addition, calculations based on the theories of referencea 1 to 5 are presented for c-i 3 m CL lift coefficient based on the exposed wing area t .= - c, pitching- coeffic

6、ient of wing about quarterwbOay. Principaldimensions ofthemodels are showninfigure 2. ThexFngmodelxasone-hEtlfacletexFngxhichhadatrlangularplan form of aspect ratio 4. Section8 in a streamwise direction were doubly symmetrical doublwwedge profiles having a msxthickness of 8 percent of the chord. The

7、 wing w and the lea- and traiwdge radifwere approximately 0.002 inch. Thebodywas e of a 2-1/24nc?+dismeter body of revolution with identical pointed ends. (See f%g. 2.) The body was constructed of aluminumalloy andthe surface was polished. For the investigation, the model.8 were mounted on a balance

8、 plate which was held in an approximately l however, lift and dreg data were obtained for angles of attack from approximately 0 to 9O at Mach numbers frcwi 0.50 to about 0.98 and at 1.20 and 1.29. Lift data for the -body combination and pitcwcment data for the wing in the presence of the body were o

9、btained for wing angles of attack from about -3O to 10 at Mach numbers from 0.50 to about 0.98 and at 1.20 and 1.28. For the tests of the uiwbody combination, the body attitude was fixed at O“, and the gap between the wing root and body remained unsealed. The wing-induced lift on the body (body atti

10、tude O“) was obtained for wing angles of attack from about -3O to U“ at Mach numbers from 0.50 to about 0.96 and at 1.20 and 1.28. For this condition, the models were mounted in an identical manner to that for the wing and body tests except that the wing spindle was held independently of the balance

11、 plate. Choking conwtions in the tunnel test section precluded testfng of the wing alone between Mach nuu .OOl . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS6 NACA RM A9UX . The draoefficient tares for the balance plate alone and with the wing in the

12、presence of the plate are presented in figure 5 (2) shock-ye b -layer interaction at the high 8ubsonic and supersonic Mach nw and (3) shock waves resulting frcan secon whereas for the e and body there is an appreciable effect at the high subsonic Mach numbers. Also shown in figure Xl(b) sre calculat

13、ed lift-curve slopes for the same three cases. Insofar as known, there are no existing theories which are directly applicable to the wing and body configuration of this report; therefore, the following procedures were employed in the calculation of the slopes. _ . For the calculations at subsonic Ma

14、ch numbers, it was assumed that the body could be replaced by a flat surface with boundaries that are formed by extending the leading and trailing edges of the all- movable wing to the axis of symmetry of the body. The surface replacing the body remains at zero angle of attack and the all+uovable wi

15、ng fe thought of as a flat, full-chord, partial-pan, outboard control surface. The theory of reference 4 then provides a method for determining the total lift on the control surface and the fixed surface, and the dis- tribution of lift between the two. Liftiurve slope8 (rate of change of lift coeffi

16、cient with control-surface deflection) calculated by the methods of this reference, however, are not specificeUy applicable to the present configuration since the theory is valid only for lifting surfaces of very low aspect ratio. It was believed, nevertheless, that this restriction could be allevia

17、ted, at least for a configuration of aspect ratio 4, if ratios of lifkurve slopes were employed, that is, the ratios of the slopes given by the the whereas those for the wing in the presence of the body are about 17 percent less. The experi- mental and calculated slopes for the induced lift on the b

18、ody arein good agreant at all the slibsonic Maoh numbers. At stqersonic Mach nmibers the experimental slopes for the wing and body are about 16 percent lea6 than those calculated; those for the wing in the presence of the body, about 13 percent less; and those for the induced lift on the body, from

19、about 25 to 30 percent Less. These disagreements between the cal- culated and experimental slopes are considered to be generally small in viewofthe procedureandtheoryemployed, andof theneglect, incomurm withthe theory forthewing alone, of viscous asd secondSwder ccqprea- sibility effects. Because th

20、e agreementbetweenthe calculatedand Provided by IHSNot for Resale-,-,-10 NACA RM A9LOl experimental lifl+curve slopes is generally so much better for the wing in the presence of the body than for the wing alone, it is felt that the influence of the tunnel-wall boundary layer on the flow around the w

21、ing in the presence of the body was small. Although there may have been a significant effect of the tunnel-wall boundary layer on the lift induced on the body, suchan effect appears to have influenced the cam- bined lifts of the wing and body but very little. This is undoubtedly due to the fact that

22、 the body lift is only a small part-of the total. It is of interest to note the difference between the calculated and experimental lift-curve slopes for a semispan all+novable triangular wing tested in the presence of a body at a Mach number of X.9 and reported in reference 12. The wing of this refe

23、rence had an aspect ratio of 2.31 and tiularc-a;r c sections P-percent chord thick, and wrc-plan-form wing of aspect ratio 4 is shown in figure 14 at subsonic and supersonic Mach numbers. For this lower boundary, the resultant force vector is inclined at its marim= calculated forward position with r

24、espect to the wing-chord line. The subsonic values of this C8hXikted drarise factor were determined by the methods of reference 1, but for all practical purposes are equivalent to the constant l/3 whereas at supersonic Mach ntmibers 2t is identical to that Cuert, H.: Wind+Iunnel Interference on Wing

25、s, Bodies, and Airscrews. R, yer and Compression Shock aud Its Effect Upon Airfoil Pressure Distributions. NACA RM A7AO2, 1947. 10. Ferri, Antonio: Ezperiment8lResulte With Airfoils Tested In the Hi+SpeedTmnelatGuidoni8. NACAmg46, 194.0. IL Love, Eugene S. : Investigation at Supersonic Speeds of 22

26、Tri- rm Wiues multiplied by the following factors Wing alone Machnumber 0.50 1.000 1.000 to o.ggg 1.000 to 1.001 -70 1.000 to 1.002 -999 to -998 1.000 to 1.001 -80 1.001to 1.003 -999 to -997 1.001 to 1.003 -90 1.001to 1.006 -998 to .gg4 1.001 to 1.006 -95 1.003 to 1.012 -997 to -987 1.002 to 1.012 -

27、98 1.006 to 1.023 .gg5 to -972 1.008 to 1.024 Wing andbody -50 1.008 to 1.012 .gg1to -988 1.005 to 1.008 -70 1.013 to 1.018 .g87 to -982 1.010 to 1.012 .80 1.020 to 1.026 .g8lto -968 1.017 to 1.022 -go 1.030 to 1.040 -972 to .953 1.028 to 1.037 .g5 1.040 to 1.055 -961 to -938 1.039 to 1.052 .98 1.05

28、0 to 1.072 .sO to .929 1.049 to I.072 Provided by IHSNot for Resale-,-,-fWe adjustment - mechanism 1 1 Aiiv/o w Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS.A/ dimensions in inches 4 3mo . f ocafion of moment/ lo.03 GOP K-$iJ7 strain gage e eat- Axis

29、of rutafion Figr/re P.- Sk& uf off=movubh? trianguhr wihg und body. . . Provided by IHSNot for Resale-,-,-NACA RM A9LCl - - I_ - - (a) Wing alone. Provided by IHSNot for Resale-,-,-Provided by IHSNot for Resale-,-,-L , I.2 x .8 1 I f06 I .7 4 .5 .6 .7 .8 .9 LO l./ 12 /.3 13 Mach number, M Figure 4.-

30、 The variation of Reynolds number wifh Mach number for the tests of the all-movable triangular wing in the Ames l-by 34 -foot Leigh-speed whd funnel E Provided by IHSNot for Resale-,-,-. .06 I I I I I I I I I I 0 Balance plate done a Wing he/din presence of bohnce plate .5 .6 .7 .8 .9 Mach number, M

31、 Figure 5.- Effect of Much number on the tore drag toe fficlents. 12 Provided by IHSNot for Resale-,-,-, . .8 .6 Provided by IHSNot for Resale-,-,-.8 .6 3 I I I I I I I I I I I I I . 4 .5 .6 .7 .8 .9 LO /.I Much number, M Figure 6- &mc/uded . (61 Gap settled, . . . Provided by IHSNot for Resale-,-,-

32、 t I I I .8 .6 . ai deg F&ure Z- EffM of Mach number on fhe l/f coeff/cien# of fhe tihg and body fw various wing angles of attack. Body attitude 05 Provided by IHSNot for Resale-,-,-4 .5 .6 .7 .8 .9 10 /.I 12 13 64 15 Mach number, a! . . Provided by IHSNot for ResaleNo reproduction or networking pe

33、rmitted without license from IHSI 1 1 ./4 .lP JO 0” ,” I I I I I I I I I l/i I I 0 # .5 .6 .7 .8 9 I. If 12 13 /.4 /.5 ch number, M Figure &I- Hfecf of Mach number on the drag coeff/chnt of the whg o/one at variws angles of attack, gap unseated. Provided by IHSNot for Resale-,-,-.08- .4 .5 .6 .7 .8

34、9 l.0 1.1 1.2 13 1.4 1.5 Mach number, M (a) Wing done. Figure 9.- Effect of Mach number on the pitching-momenf coefficients of the wing both alone and in the presence of the body Body utfitude 00, gap unsealed Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS. * I .08 I -I - 4 .5 .6 .7 .8 .9 l.0 /.I L2 13 i.4 15 Mach number, A4 (b Wing in prssence ofbooy. Figure 9- Concluded. Provided by IHSNot for Resale-,-,-