NASA NACA-RM-L51A10-1951 An investigation of the effect of vertical-fin location and area on low-speed lateral stability derivatives of a semitailless airplane model《垂直翼片位置和区域对半无尾飞.pdf

《NASA NACA-RM-L51A10-1951 An investigation of the effect of vertical-fin location and area on low-speed lateral stability derivatives of a semitailless airplane model《垂直翼片位置和区域对半无尾飞.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L51A10-1951 An investigation of the effect of vertical-fin location and area on low-speed lateral stability derivatives of a semitailless airplane model《垂直翼片位置和区域对半无尾飞.pdf（42页珍藏版）》请在麦多课文档分享上搜索。

1、RESEARCH MEMORANDUM AN INVESTIGATION OF Ti33 EFFECT OF VERTICAL-FIN LOCATION AMD AREA ON LOW-SPEED LATERAL STABILITY DERIVATIVES OF A SEMITAILLESS AIRPLANE MODEL By Lewis R. Fisher and William H. Michael, Jr. Langley Aeronautical Labratory Langley Field, Va NATIONAL ADVISORY COMMITTEE FOR AERONAUTIC

2、S WASHINGTON March 7, 1951 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 NACA RM L5IA10 NATIONAL ADVISORY COMMIlTEE FOR AERONAUTICS RESEARCH MEMORANDUM AN INVESTIGATION OF TEE EFFECT OF VEKEICAL-FIN LOCATION AND AREA ON LOW-SPEED LFlTERAL STABILI

3、TY DERIVATIVES OF A SEMITAILLESS AIRPLANE MODEL By Lewis R. Fisber and William E. Michael, Jr. SUMMARY The results of a low-speed wind-tunnel investigation to determine - the effects of vertical-fin location and area on the static and rotary lateral stability characteristics of a adtailless airplane

4、 model indicated that the contributions of the vertical fin to the stability derivatives could be estimated with reasonable accuracy by simple considerations in spite of the unusually ahort tail length ad large tail height of this configuration. - Although the differences in fin effectiveness are no

5、t large, for comparable fin areas and equal tail lengths, fins located at the 86- percent spanwise location are, in general, more satisfactory for providing directional stability and daurping in yaw than fine located at the ), V(bhS), and V(ck3). For the V(bk3) and V(c4.S) configurations, two additi

6、onal mall fins were placed on the bottom surface of the wing opposite to and in the same plane as fins (bz) and V(-), respective-. I V(b2S) , and V( c2S) all have the 8- area, whi-ch is one-half that for - The symbols refer to the complete model having the epecific fin arrangement tested. In figures

7、 illustrating the effects of the addition of various component pafts, the symbol W refers to the isolated wing and the symbol FTF refers to the wing-fuselage combination. The arrange- ment V(aL) is considered to be the basic confivation. Photographs of the model mounted in the Langley stability tunn

8、el with three different fin arrangements are shown in figure 3. The effects of the addition to the basic configuration of a cockpit canopy and wing-root engine nacelles (fig. 2) were determlned in the investigation. The basic model was also tested with the addition of full-span slats attached to the

9、 leading edge of the wing. These slats were shaped fram - inch aluminum sheet to -the contour of the wing leading edge. The V(b2L) fin arrangement W the latter effect probably remlted from the load carried by the slat itself, which effectiyely extended the wing leading edge fo-d. - was extended and

10、tk stabili%y generally reduced, particularly at the - Lateral characteristics.- The static-lateral derivatives of the basic model and all alternate configurations were obtained between side- slip angles of 530 and are presented in figure 5. Results for the wlng alone, the wing and fuselage, and the

11、complete basic model (fig. 5(a) show that the wing alone has a small amount of directional stability as is indicated by the small negative values of that the wing-fuselage in directional stability for the complete V(aL) configuration, combination is unstable, a) in producing the derivatives cy, Cnr,

12、 and -Czr (fig. 6(e). The addition of dorsal fins to V(b2L) resulted in no appreciable change in the damping in yaw (fig. 6(f). At the O.% position, the V(c2.L) arrangement is slightly more favor- b “ able .than the V(c4) arrangement for producing damping in yaw (fig. 6(g). The tests made with the V

13、(b at the 86-percent spanwise position, they contributed a sizable positive increment. This change in sign of the effective dihedral with lateral movement of the fins is believed to be the result of the nature of the span loadings induced on the wzhg by the lift on the fins. Langley Aeronautical Lab

14、oratory National Advisory Cammittee For Aeronautica Langley Field, Qa. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 “ . . . NACA RM L5IAl-0 . “ 1. Bird, John D., Jaquet, -on M., and Cowan, John W. : Effect of Fuse- lage and Tail Surfaces on Low

15、-Speed Yawing Characteristics of a Swept-Wing Model as Determined in Curved-Flow Teat Section of Langley Stability Tunnel. NACA RM L8G13, 1948. 2. MacLachlan, Robert, and Letko, William: Correlation of Two Experi- mental Methods of Determining the Rolling Characteristics of Unswept Wings . NACA TN 1

16、309, 1947. 3. DeYoung, John: Theoretical Additional Span Loading Characterietics Of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. mACA TN 1491, 1947. 4. QueiJo, M. J., and Wolhart, Walter D.: Experimental Investigation of the Effect of Vertical-Tail Size and Length and of Fuselage Shape

17、 and Length on the Static Lateral Stability Characteristics of a Model w&th 45 Sweptback Wing and Tail Surfacee. NACA TN 2168, 1950. - 5. Kateoff, S., and Mutterperl, William: The End-Plate Effect of a Horizontal-Tail Surface on a Vertical Tail-Surface. NACA TN 797, 1941. 6. Thompson, F. L., ana Gil

18、ruth, R. R.: Motes on the Stalling of Vertical-Tail Surfaces and on Fin Design. MCA Tm 778, 1940. 7. Goodman, Alex: Effect of Various Outboard and Central Fins on Low- Speed Yawing Stability Derivatives of a 60 Delta-Wing Model. NACA RM LWE12a, 1950. 8. Toll, Thomas A., and Queijo, M. J.: Approximat

19、e Relations and Charta for Low-Speed Stability Derivatives of Swept Wings. NACA TN 1581, 1948. 9. Campbell, John P., and Goodman, Alex: A Semiempirical Method for Estimating the Rolling Moment Due to Yawing of Airplanes. MACA TN 1984, 1949. 10. Goodman, Alex, and Fisher, Lewis R.: Investigation at L

20、ow Speeds of the Effect of Aspect Ratio and Sweep on Rolling Stability Derivative8 of Untapered Wings. NACA Rep. 968, 1950. 11. Goodman, Alex, and Adair, Glenn H. : Estimation of the Damping in Roll of Wings Through the Normal Flight Range of LiFt Coefficient. MACA TN 1924, 1949. Provided by IHSNot

21、for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5UO 15 12. Bamber, Millard J.: Effect of Some Present-Day Airplane Destgn Trends on Requirements for Lateral Stability. XACA 814, 1941. 13. Letko, William, and Riley, Donald R. : Effect of an Urnwept Wing on . th

22、e Contribution of Unswept-Tail Configurations to the Low-Speed Model. NACA TN 2175, 19%. - Static- and Rolling-Stability Derivatives of a Midwing Airplane 14. Murray, Harry E. : Wind-Tunnel Investigation of End-Plate Effects of Horizontal Tails on a Vertical Tail Compared with Available Theory. NACA

23、 TN 1050, 1946. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM 5110 pltctung moment 9 Roffmg moment - Relatwe wind Refatwe wind Yawmg moment v Figure 1.- The stability system of axes. Arrows indicate positive direc- tions of forces, mom

24、ents, and angular displacements. This system of axes is defined as an orthogonal system having the origin at the center of gravity and in which the Z-axis is in the plane of symmetry and perpendicular to the relative wind, the X-axis is in the plane of symmetry and perpendicular to the Z-axis, and t

25、he Y-axis is perpendic- ular to the plane of symmetry. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WIW fins Rn L Aspect rono 3.60 0.4# 0.455 Toper roflo 1.85 185 NACA 0010- 64 AwfcuJ sstm a296 0.420 0.976 lweon owodynomlc &d, ft 41“ 4. 41.5 Swp

26、, L. E, deg 0.53 0.75 536 Span, ft 0.15 630 3.13 &eo, sq ft 0.444 /- 2.10 ./qj 4 4.73 1 I -1538- / 7 r 7701 I , W Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5UO I I t. 7 3.94 (b) Additional fin arrangements tested. Figure 2.- Continued.

27、 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NMA RM 5110 Locuhon und exfreme drmensAons of eng/nenuce/es und cock campy (c) Details of cockpit canopy, engine nacelles, and dorsal fins. Ffgure 2.- Concluded. Provided by IHSNot for ResaleNo reprodu

28、ction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA m 5110 21 Figure 3.- semi (b) Vertical-fin configuration V(b2L). IC) Vertical-fln configuration V(b&). v L-68421 tailless airplane mdel mu

29、ted in the curved-flow test sectioa of the Langley stability tunnel. 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-,-,-NACA RM L5lAlO 72 0 2 4 6 .8 Lift c

30、oefficmf, CL (a) Component parts of basic model. Figure 4.- Aerodynamic characteristics of semitailless airplane model. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 ./ NACA RM 5110 . . . _. :2 .2 .4 .6 .8 10 LtfZ coeffrclent, CL (b) Effects of

31、canopy, nacelles, and slats Figure 4.- Concluded. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“. , I 1 I Do8 .ow 0 :P 0 .2 .4 .6 .8 LO 72 0 .2 .rp hft CmffWM, CL Llft meffwent, CL (a) Effect of component pts of (b) Effect of cockpit canopy, win

32、g- basic cmguration. rout nacelles, an8 lea-edge Bh%S M to basic configuration. Figure 5.- The variation wlth lift coeMcient of the static lateral stability derivatives for a semitailless airplane model Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,

33、-.w2 Dm m4 0 C Y P .a2 m3 rn 0 72 0 .2 .4 .6 .8 LO 32 0 .2 .4 .6 .8 LO Llfi meffmant, c L LIft &f&, c -537 1 - (c) Effect of longitudinal position (a) Effect of lateral position of of vertical fin. = = 1.20, at vertical fine for conetaat fin C area and tau length. 9, = 0, 0.85. b/2 Egure 5.- Ccmtixl

34、ued. 0.43, 0.86. L I . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a2 1708 rn 0 I I 1 # -u V(bZl), Dwsds 9 0 .2 .4 .6 .8 LO 72 0 .2 .4 .6 Llft &flmlt* CL Llft iweff&, 5 (e) Effect of J.ateml positicm and (f) Effect of dorsal addltion to increas

35、ed area for vertical. flns inboard wing fin and alternate th BIE tail le. JL = 0, arrangement of III am. 0.43. 0.43, 0.66. b/2 & Figure 5.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. .w6 .OM 5, m4 0 -902 P . “ V(b4S) -.2 0 .2 .4 .6

36、.8 LO 52 0 .2 .4 .6 Llft cm?ffmt, CL Lft canffic&, 5 (g) Effect of dternste amangemnt. (h) EfPect of lateral position for of fin area at wtbaard Vrsg the alternate fin arrangement. pOEitiOll. = 0.86. b/2 b7 = 0.43, 0.86. Figure 5.- Concluaed. . Provided by IHSNot for ResaleNo reproduction or network

37、ing permitted without license from IHS-,-,-1 .P .4 O CY .2 r r 52 0 “W .I co r n .I co r I -. I 0 C n r -.I .3 .2 C I J r 0 -.I (a) Effect of canrponent parts of (b) Effect of cockpit canopy, wbg- basic conflguration. root mcelles, and Leading-edge slate as addltiom to basic COnfigKfatiOn. Figure 6.- The varlatian with lift coefficient of the yawing stability derivatives for a semitailleas &plane mcdel. . . E 0 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-