NASA NACA-RM-L50E26-1950 Damping in roll of rectangular wings of several aspect ratios and NACA 65A-series airfoil sections of several thickness ratios at transonic and supersonic ls《在.pdf

1、6 RESEARCH MEMORANDUM DAMPING IN ROLL OF RECTANGULAR WINGS OF SEYERAL ASPECT RATIOS AND NACA 65A-SERIES AIRFOIL SECTIONS OF SEVEIUL THICKNESS RATIOS AT TRANSONIC AMD SUPERSONIC SPEEDS AS DETERMINED WITH ROCKET-POWERED MODELS By James L. Edmondson ., Langley Aeronautical Laboratory Langley Air Force

2、Base, Va. t. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTPN August 24, 1950 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. NATIONAL ADVISORY C0M“TEE FOR AERONAUTICS I DAMPING IN ROLL OF RECTANGWAR WINGS OF SENERAL ASPECT RATIOS AND NACA 65

3、A;SERIES AIRFOIL SECTIONS OF SEVERAL THICKNESS RATIOS AT TRAMSONIC AEhD SUPERSONIC SPEEDS AS DETERMTNED WITH R0CKEX“POWERED MODELS By James L. Edmondson Rocket-powered flight tests have been conducted to determine the damping in roll of rectangular wings of various aspect ratios and thickness ratios

4、 with the use of the RACA A-series airfoil sections. The Mach number range of these tests was from approximately 0.8 to 1.4. The experimental damping in roll was consistently lower than that $re- dieted by linear theory, and-this difference increased with aspect ratiog The experimental damping in ro

5、ll decreased as wing thichess ratio was increased. * - angular wings of finite operation for this factor. thickness ratio. Further data are needed to determine the limits of , - INI3ODUCTION A damping-inlroll investigstXon has been conducted for a eeries of wings of several aspect ratios and airfoil

6、 thickness ratios using . the NACA 65A-series airfoil section. Previous damping-in-roll tests of rectangular wings by this technique (reference 1) indicated that experimental damping would vary with airfoil thickness ratio; therefore, the present series of tests were conducted to determine the relat

7、ionship between the damping in roll for both thickness ratio and aspect ratio. UNCLASStFlED ! I I I i ! I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 - NACA RM L50E26 The testainge were mounted on identical fuselages incorporating canted nozzle

8、s, as described in reference 1. The damping-la-roll coefficient and the total-drag coefficient were obtained for each configuration at zero lift through a Mach nmiber range of approx - mately 0.8 to 1.4, corresponding to Reynolds numbers from 3 X 10 to 8 X lo6. The models were tested in ?Light at th

9、e Pilotlees Aircrae Research Station at Wallops Island, Va. 2 CZ C 2P CD D L Lp LO T v q M A R -: b SYMBOLS damping-in-roll coefficient (2) total-drag coefficient (D/s) total drag, pounds roll damping moment, foot-pounds rate of change of damping moment with rolling velocity, foot-pounds per radian

10、per second out-of-trim rolling moment, foot-pounds torque, pound-foot rolling angular velocity; radians per second rolling angular acceleration, radians per second 2 forward velocity, feet per second dynamic pressure, pounds per square foot Mach number aepect ratio (b2/S I) Reynolds number, based on

11、 wing chord airfoil-section thicknees ratio Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5026 3 b a b S S =X Subscripts: 1 2 body diameter, feet wing span, feet (diameter, of circle generated by wing tips) total wing area of two wings, s q

12、uare feet (wing panel assumed to extend to model center line) total wing area of three wings, square feet (wing panel assumed to extend to model center line) moment of inertia about longitudinal axis, slug-feet- 2 sustainer-on flight coasting flight MODEL AND APPARATUS The models used in this invest

13、igation were identical to those reported in reference I except for wing design. The basic body consisted of a wooden fuselage containing EL spinsonde nose sectLon (reference 2) and using a sustaining rocket motor with canted nozzles. The test wings were attached near the rear of this basic fuselage

14、in a three-wlng arrangement. Wing aspect ratios of 2.5, 3.0, 3.7, and 4.9 using the NACA 63A009 airfoil sectfon and atrfoil thickness ratios of 0.06, 0.09, and 0.12 on wings of aspect ratio 3.7 were tested. A sketch of the model configmation and pertinent- Lng geometry are . given in figure 1. The m

15、odels were boosted from a rail launcher to a Mach.n models 3 and 6 were reported in reference 1 and the results are repeated herein for comparison. The rate of roll for models 1, 2, 4, and 5 is plotted against Mach nuniber in figures 2(-a), 2(b), 2(c), and 2(d), respectively. For models 2 and 5, for

16、 which records of duplicate models are shown, the difference in rate of roll with sustainer on was caused by a difference in torque produced by the canted nozzles. The rate-of-roll variation or lateral trim change through the transonic speeds during coasting flight has been discussed in reference 3.

17、 The severity of this lateral trim change seems to vary directly with wing thickness. An apparent discrepancy in rate of roll at M X 0.93 during sustainer-on flight is noted in figure 2(b). This discrepancy is the result of (I) the, short time to record data and (2) the lateral trim change which is

18、caused by local flow conditions dependent upon airfoil section and surface conditions (reference 3). The variations of experhental Czp with Mach nuniber are shown in figure 3. Also shown are supersonic theoretical curves of from reference 4. This theory was derived for an isolatedwlng; however, the

19、interference effects of a body and three wings are considered small through the body diameter-to-wing span ratios and Mach nmbers of these tests. This has been shown by unpublished Czp data of wing alone, body plus two wings, and body plus three wings using wing plan form and body of model 3. Figure

20、s 3(a), 3(b), and 3(c) present the damping for the aspect-ratio series and show that the experimental curves are con- sistently lower in magnitude than theory. This difference between experiment and the-qy, however, varied directly with aspect ratio; the larger aspect ratios show a greater differenc

21、e. Figure 3(b) shows the effect of the lateral trim change to be an apparent increase or decrease in damping, depending upon the relative magnitudes and direction of this trim change. QP Figure 3(d) shows the damping for one of the thickness-ratio series; the other thiche.ss-ratio models were report

22、ed in reference 1. A comparison of these thickness-ratio tests showed that the damping .in roll varied inversely with wing thickness ratio; the greater thicbesses showed the less damping in roll. : I I ! t I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IH

23、S-,-,-The effects of aflpect ratio and thickness ratio on damping in roll are summarized in the following table: I I I I r lA NACA airfoil section -CzP at M = 1.15 -c zp at M = 1.30 65A009 65AOO9 65AOO9 65AOO9 65A006 65012 0 9 255 3a 385 .440 354 .418 . An empirical correction factor which relates t

24、hese experimental data with theory was derived to be used with existing supersonic linear- theory to allow prediction of Czp for wings of various thickness ratios. This factor, to be multiplied by values from supersonic linear theory, was found to be dependent upon wing aspect ratio as well as airfo

25、il thickness ratio and is expressed as (1 - ;y3. The comparison of experimental C with corrected theoretical C2 for the thickness-ratio series is shown imigure 4(a). The solid curve8 are the correclxd theory for the various thickness ratios. The thickness ratio of zero makes the correction factor eq

26、ual to unity; therefore, this curve is the same as uncorrected linear theory. Experimental CZp is shown as dashed lines. These curves are the faired values from fig- ure 3(d) of this paper and figure 9 of reference 1. Heretofore, the experimental curves for all these thickness ratios were compared t

27、o uncorrected theory shown as zero thickness ratio. It-can readily be seen that the use of this empirical correction factor allows a much closer theoretical prediction of CzP for these test wings. 2P P The agreement-of the corrected theory with experimental data for the aspect-ratio series is shown

28、in figure 4(b). Again, the corrected theory is shown as a solid curve for each aspect ratlo, and experimental data fbr these aspect ratios are shown as dashed lines. The comparison of experimental CzP with uncorrected theory has previously been shown in figures 3(a), 3(b), and 3(c) of this paper and

29、 figure 91.a) of refer- ence 1. The use of the empirical correction factor allows- a much closer prediction of Cz for this range of aspect ratio. P + Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I . NACA RM 5026 - 7 Tests of a rectangular wing of

30、aspect ratio 45 and NACA 65-006 airfoil section (reference 5) .also showed good agreement with corrected theory. However, tests of double-wedge airfoil-section wing from refer- ences 5 and 6 show good agreement with supersonic linear theory above M Z 1.25 without using a thiclmess correction factor.

31、 It is thus indicated that this factor will apply to a rounded-nose, moth-contour airfoil of these wing-body coribinations; however, additional data will be needed to determine the limits of operation. . The total-drag coefficients of theee configurations were also obtained from these tests. The CD

32、are directly comparable because the wing area was constant .in all cases. TIE relative effects of thickness .ratio and aspect ratio on total drag are shown in figure 5. All configurations had approximately the same drag at aubsonic speeds. At. supersonic speeds the effect of increasing aspect ratio

33、was a small increase in drag. However, as would be expected, the effect of increasing wing thicbess ratio was to cause an earlier transonic drag rise and an appreciable increase in supersonic drag. CONCLUSIONS ! I ! I ! The following conclusions were drawn from teats of rectangular wings having NACA

34、 65A-series airfoil sections, aspect ratios from 2.5 to 4.5, and thickness ratios from 0.06 to 0.12: 1. Damping in roll increased with increasing aspect ratio but at a slower rate than predicted by linear supersonic theory. 2. Damping in roll decreased with an increase in thickness ratio. 3. An empi

35、rical relationship factor was eetablished which, when applied to linear.theory, allows 831 accurate prediction of the damping in roll at supersohic speeds for the caBes investigated. Ikngley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Afr Force Base, Va. I I I I Provi

36、ded by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 REFERENCES MACA RM 5026 1. Edmondson, James L., and Sanders, E. Claude, Jr: A Free-Flight Technique for Measuring Damping in Roll by Use of Rocket-Powered Models and Some Initial Results for Rectangular Wi

37、ngs. NACA RM LgI01, 1949. 2. Harris, Orville R.: Determination of the Rate of Roll of Pilotless Aircraft Research Models by Means of. Polarized Radio Waves. NACA TN 2023, 1950. 3. Stone, David G.: Wing-Dropping Characteristics of Some Strafght and Swept Wings at Transonic Speeds as Determined with R

38、ocket-Powered Models. MACA RM L5OC01, 1950. 4. Harmon, Sidney M.: Stability Derivatives at Supersonic Speeds of Thin Rectangular Wings with Diagonals ahead of Tip Mach Lines. NACA Rep. 925, 1949. 5. Dietz, Albert E., and Edmondson, James L.: The Damping in Roll of Rocket-Powered Test Vehicles Having

39、 Rectangular Wings with NACA 65-006 and Symmetrical Double-Wedge Airfoil Sections of Aspect Ratio 4.5. NACA RM L5OB10, 1950. 6. Brown, Clinton E., and Heinke, Harry S., Jr.: Preliminary Wind- Tunnel Tests of Triangular and Rectangular Wings in Steady Roll at Mach Nuuibere .of 1.62 and 1.92. NACA RM

40、L8L30, 19497 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . I 4 I 4.5 I 8.343 6 I 3.7 1 9.192 I 5 I- 7*4 1 I I I I 17.0391 ./9/ 165AOO613.8 to 7.8 Figure 1.- Sketch and wing gemtry for the model tested. All dlimensions In inches. . - Provided by

41、 IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 - NACA RM 5026 C (a Model 1; A = 2.5; t =-0.09. Figure 2. - Variation of rolling velocity with Mach number. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. (b

42、) Model 2; *A = 3.0; - t = 0.09. C Fiwe 2.- Continued. ! i I I I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 - NACA RM 5026 rad/an 5ec 0 (d) Model 5; A = 3.7; - t = 0x2. C Figure 2.- Concluded. Provided by IHSNot for ResaleNo reproduction or n

43、etworking permitted without license from IHS-,-,-. c NACA RM 5026 .I_ (a) Model 1; A = 2.5; - = 0.09. t C .6 .4 -c+ .2 0 I 13 (b) Model 2; A = 3.0; Figure 3.- Comprison of: experimental reference 4. t - = 0.09. C czP with theory from I : ! I ! I i I I ! I Provided by IHSNot for ResaleNo reproduction

44、 or networking permitted without license from IHS-,-,-14 .6 .4 .2 NACA RM 5026 .6 (c) Model 4; A = 4.5; 2 = 0.09. .4 0 (a) Model 5; A = 3.7; - t = 0.12. Figure 3. - Concluded. C Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5026 - .6 .4 0 (

45、a) Thickness-ratio series; A = 3.7. /-2 L3 M (b) Aspect-ratio eeries; - = 0.09. t C Figure 4.- Comparison of experimental C with empirical Cz - P I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . . . . . . . . . . . . . . . . . . .

46、C D 1 07 .8 09 /* 3 I ! I Figure 5.- Variation of total-drag coefficient with Mach number showing the effect of aspect ratio and thiclmess. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-f Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-