NASA NACA-RM-L8L29-1949 Low-speed static-stability and rolling characteristics of low-aspect-ratio wings of triangular and modified triangular plan forms《三角形和改良三角形平面低展弦比机翼的低速静态稳定性和.pdf

《NASA NACA-RM-L8L29-1949 Low-speed static-stability and rolling characteristics of low-aspect-ratio wings of triangular and modified triangular plan forms《三角形和改良三角形平面低展弦比机翼的低速静态稳定性和.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L8L29-1949 Low-speed static-stability and rolling characteristics of low-aspect-ratio wings of triangular and modified triangular plan forms《三角形和改良三角形平面低展弦比机翼的低速静态稳定性和.pdf（45页珍藏版）》请在麦多课文档分享上搜索。

1、0 -3 f3 e- 9 copy No. RM No. L8L29 J “ . :. r -“ - I . “ RESEARCH MEMORANDUM LOW-SPEED STATIC-STABII;ITP AND ROLLING CIULRACTEmTICS OF JLO-ASPECT-RATIO WINGS OF TRIANGULAR AND 5-cl: MODIFIED TRIANGULAR PLAN FORhE Byron M. Jquet and Jack D. Brewer Langley Aeronautical Laboratory Langley Ar Force Base

2、, Va. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON March 29,1949 .- “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. I . By Byron M. Jaquet end Jack D. Brewer A low-speed investigation was made in the Langler stabilitp tunnel to determin

3、e experimentally the effects of change8 in profile and aspect ratio on the low-speed static-stability asd rolUng characteristics of triangular wings. The invemic center position, and the dqing in roll, the agreement was poor except at the lowest test aspect ratio (A = 1.07)- The vertical ffna teated

4、 provided good directia stability through- out the entire Uft-coefficient range. The fins increased the vazious portions from the tip6 of a basic triangular wing genemy had good longitudinal and directional stability but very hi hereinafter, each model will be referred to by the nunibsr designated i

5、n the table. All profiles referred to are parallel to the plme of .sprmtry. All the tests were mad.e on a six-coqonent strain-gage balance strut with the models mounted at a point two-thirae of the root chord from the apex of the triangles Figure 2 presents the profiles of the series of dele having

6、re not presented for sngles of attack greater than approximately 160. The measuremsnts taken are believed to be accwate within the following amounts which axe based on the values of the forces and mments for model 6 : a, deg -10.1 $, deg . 50.2 cL . 0.0029 Cy . iO.0046 Presentation of Resulta The st

7、atic and rolling ch increased longitudinal stability is noted at the 8m.e lift coefficient. An opposite trend is noted for model 7 at CL = 0.6 where a decrease occurs In the %lax LCG Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-0 longitudinal stab

8、ility and in the lift-curve slope. Increased longitudinal stability is noted throughout the lift-coefficient range however, the aerodpamic center is in a more forwarrd position than the results of the present investigation indicate. 1% should be noted that all the triangular wings tested in referenc

9、e 8 had flat-plate airfoil sections, whereas those tested herein had an NACA 0012 profile with a larger tra3ling-edge angle and a blunt 1ead-g edge Large-scale teste of a triangular wing Xith a double-wedge airfoil (5 percent thick at 20 percent chord) indicate about the sam position of the aerodync

10、 center as does the present investigation for ELII aspect ratio of 2 .O . (he reference 9 .) The trend of C4pl%x in figure 20 with the trends of reference 8 in that %e peak vahe of C was reached at about the sane aspect ratio. As the aspect ratio is decreased, the lift-curve slope is decreased as ca

11、n be seen in figure 20. The swept-wing theory of references 3 and 5 shows fair apamsrit with the experimental data for the aapect-ratio range considered. The theory of reference 2 approaches the experimental values only.= the aspect ratio approaches zero as would be erpected. It should be remembered

12、 thak a.i the profiles for the models for which the data are presented in figure 12 are of NACA OOl2 sections parallel. to the plane of symmetry- With a highly wept model (as model 4) there is 8 very large area in the plane of sgmmstry forward of the quarter chord of the mean aerodynamic chord which

13、, when the model is yawed or rolled, acts in the manner of a fin. Model 4, because of this mea, has positive values of the directional-stabilftp parameter C below CL = 0 -73. The model does have increasing directional stability at the stall whfle models 2 and 7 do not. If mdel 4 was equipped with a

14、high-a8pect-ratio fin, the objectionable characteristics below CL = 0.73 might be overcome, resulting in a model having better over-all character- istics than mdels 2 or 7. The values of however, the experimental values of C are negative only at moderate lift coefficients (See fig. 13. ) The availab

15、le theory is, therefore, extremely limited in the range of applicability to triangular plan forma AB the aspect ratio is decreaseicms increaee thmughout the entire Uft; range. The addition of either fin causes a positive increment of C at CL = 0 which decreases aa the lift coefficient incrreases. %

16、for z* rate of change of C with CL for the rlng.fin 2J than for the wlng alone (See fig 15 ) Addition of the large fin cauaed negative displacements of at cyP CL = 0 ; the addttion of either fin caused an increase in the rate of change of Cy with CLg (See fig. 16 .) n almost constant positive increm

17、ent in hp is the reat of adding the large fin to model 2 . (see fig 16 . ) P Both the lmge and amall fins increwed the wing in roll tlwoughout the Jlft-coefficient range; the increase for the lcarge fin mounts to about 30 percent of the damping in roll of the wing alone. Effect of hpect Ratio of Mod

18、ified Triangular PLan Born The model6 o$ the present grog were fonnsd by cutting various portloll from the tips of a baeic trt.angular wing (model 7) parallel to the plane of symmstry to obtain alspect ratio 3 (model 8), aspect ratio 2 (mdel 9)y and aspect ratio I (model 10 ) including tips of revot

19、ion. For the lowest test aspect ratio (A = 1.U) a definite nonlinear variation of CL with a is noted in figure 17 which agrees with the rests of sim11a.r models tested in reference 8. whereas the theow of reference 5 indicates a emall reduction in this parameter wfth decreasing aspect ratio and the

20、triwar theory of reference 2 iucatee a large increase in aC wfth decreasing aspect ratio. In general, better agreement was obtained with reference 5 The expemntal curve of figure 25 indicates a decrease in the variation of c”p with CL with decreasing Reducing the aspect ratio results in a decrease i

21、n damping in roll C of reference 5 whfle the theory of reference 2 fs in fair agreement with experiment nly at the lowest test aspect ratio (A = 1.0). me theory of reference 4 is in fair agreement at lo aspect ratios and sxceent agrement at A = 3 and A = 4. P “.P YPP, -. aepect ratio which is in fai

22、r agreement xith the theory of reference 5. - (See fig 25 1 GOOted with fair accuracy by available swept-wing theory. Available low-aspect-ratio triangular-wing theory was found to be reliable for Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 NA

23、CA RM Ho- L8L29 certain characteristic8 but for othere, particularly the lift-curve slope, damping In roll, and aerodynsmic-center position, the agreement was poor except at the lowest teat aspect ratio (A = 1.07) . 3- The vertical fin8 tested provided good directional atability for the wing-fin cop

24、lbinationa throughout the lift-coefficient range. The fins increased the ming in roll and decreased the Tarlation of effective dihedral parameter with lift coefficient. 4. The eerie8 of wings obtained by cutting various portions from the tips of 8 basic triangular wing generally had good 1ongLtudina

25、l and directional stability but very high effective dihedral. Most of the characterietice of these wings, at low lift coefficients, could be predicted wlth fair accuracy by means of available ewept-wiq theory Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Air For

26、ce Base, Va. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-XACA RM NO L-9 13 1. Jones, Robert T. : Properties of Low-Aqect-Ratio Pointed Wings at Speeds below and above the Speed of Sound. NACA Rep. Eo. 835, 1946. 2 Ribner, Herbert S . : The Stabil

27、lty Derivatives of Low-Aspect-Ratio Triangular Wings at Subsonic and Supersonic Speeds. NACA TN No. 1423, 1947. 3 - DeYoung, John: Theoratical Additional Span Loading Chasacterfstics of Wings with Arbitrary Sweep, Mpect Ratio , end Taper Ratio. XACA TN NO 1491, 1947 4. Bird, Joh D.: Same Theoretical

28、 Low-Sped Span Loading Characteriotics of Swept wings in RO and sideslip. NACA Tpa No. 1839, 1949. 5. Toll , Tho- A. , and Quei 30, M. J . : Approxfmate Relations and Charts for Low-Speed Sta;biuty Derivatives of Swept wings. NACA TN NO 1581, 1948 6 MacLachlan, Robert, and Leth , William: Correlatio

29、n of Two EQerimsntal Methods of Determining t?le Rollinn Characteristics of Unmept Wings. XACA TN No - 1309 , 1947. 7 Letko , William, and Jquet, Byron M. : Effect of moil Profile of Symmetrical Sections on the Low-Speed Static-Stability and Yawing Derivatives of 45O Sweptback Wing Models of Aspect

30、Ratio 2.61. NACA RM No L-, 1948. 8. Tosti , Louis P . : Law-Speed Static Sta.bili$y and DqFng-Fn-Roll Characteristics of 45 0 - 0 c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“ . Figure 1.- System of stability axes. Positive forces, moments, and

31、 anglee are indicated. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. “ . . , . . . . . . . . . . - . . - -. . . . . . - . . . . .- t . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. I. b

32、. r . . I 1 L Provided 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-,-,-Figure 4.- Model 4 mounted in tunnel. A = 1.07; Ac/4 = 70.4“; profile, WA 0012. Provided b

33、y IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. I . . . t a . . . . . . . . I L Provided by IHSNot for ResaleNo reproduction or networking permitted without lic

34、ense from IHS-,-,-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. W Iu Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without

35、license from IHS-,-,-I I . . . I . . . 312. . . Provided 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-,-,-. Figure 8.- EfYect of profile on aerodymmlc characteris

36、tics of a triangulm wing of aspect ratio 2.31. A44 = 59.2. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACA RM No. L8L29 P Figure Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. L9 29 t . :Z O .2 -4 6 .8 LO 27 mefficxmf, 2P0 Figure 10.- Effect of profile of a triangulsr wing of aspect ratio 2.31 n cyp, cnp, and c 4 = 52.2O. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-