NASA NACA-TN-730-1939 Wind-tunnel investigation of effect of yaw on lateral-stability characteristics II - rectangular N A C A 23012 wing with a circular fuselage and a fin《偏航对带有圆形.pdf

1、I- 7- I FILE COPY L-l NO. 5 TECZTICAL ETOTES NATIOlrAL ADVISORY COMMITTEE ZOR AEROXAUTICS IJO l 730 1fIBD-TUMXEL IMVESTIGATIOX OF EFEKECT OF YAW OW LATERAL-STA3ILITY CEARACTERISTICS II - RECTAUGULAR B.A.C.A. 23012 WIKG WIT!y A CIRCULAR FUSELAGE.AWD A FIEt 3y M. J. Sanbor and R. 0. Eouse . Langley Me

2、morial Aeronautical Laboratory THIS DOCUMENT 014 i.OA,Y FROM THE FILES CF NATlON ADVISORY COMMITTEE FOR AERONAUTICS LANGLEY AERONAUTICAL LAGCRATORY IANGLEY FIELD, HAMPTON, VIRGINIA RETURN TO THE ABOVE AgDi?ESS ,.-t-.- - _. REQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED Washington As fOLLoWS: -mi- Sep

3、tember 1939 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1724 F STREET, N.W., WASHINGTON 25, D.C. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS TECHNICAL NOTR NO. 730 WIND-TUNNEL INVESTIGATION OF EFFECT

4、 OF YAW ON LATERAL-STADILITY CHARACTERISTICS II - RECTANGULAR N.A.C.A. 23012 WING WITH A CIRCULAR FUSELAGE AND A FIN By M. J. Danber and R. 0. House SUMMARY An N.A.C.A. 23012 rectangular wing with rounded tfps was tested fn combination with a fuselage of circular cross section at several angles of y

5、aw in the N.A.C.A. 7- by lo-foot wind tunnol. The model was tested as a high- wfng, a midwing, and a low-wing monoplane-; for each wing location, tests were nade wfth two amounts of dihedral and with partial-span split flaps. For each combination of wing and fuselage, tests were made with and withou

6、t a fin. Sample charts of the uoofficients of rolling moment cat, yawing nonent Cnl, and lateral force Cyt are given for some of the conbinations tested. The rate of change in the coefficients with angle of yaw Q* is given for stability caiculation. The value of the effect on dCl/d$ of wing-fuselage

7、 intcrferenco change sign as the wing was moved downward on the fuselage from the high to the low position, being zero at some intermediate position. In goneral, the change in dG/dQ for a given change in the dihedral was only slightly affected by wfng-fuselage interference. Moving the wing from the

8、low to tho high position generally tondod to fncreaso the value of dCa/d#/ and dCn/dQ and to decrease the effectiveness of the fin on dCa +b. dC,*/d$l, and dCyf/d;r. INTRODUCTION Estimation of the stability characteristics of air- planes can be nade only if the stability derivatives of Provided by I

9、HSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 N.A.C.A. Technical Note Bo. 730 . the particular airplane are known. Mathonatical equations are available and convenient charts are given in rcferenco 1 for predicting the lateral-stability characteristics of airp

10、laaos. Some of the factors affecting the values of those.derivatives arc wing, fuselage, and fin forns and the aerodynamic interference between these parts. =* The effects on the lateral-stability characteristics that depend on yaw of several wing ferns, changes in tip shape, dihedral, taper, wing s

11、ection, sweep angle, and split flaps are given in references 2 and 3. A theoroti- cal predfction.of sone of the lateral-stability charac- teristics for wings is given in reference 4. The purpose of the present investigation was to obtain infornation relative to the interference effects on lateral-st

12、ability characteristics of wing-fuselage-fin conbinatfone, -Tests were nade with a fuselage of circular cross section having a re-movable fin; the fuselage was tested separately and in combination with.the K.A.C.A. 23012 rectangular wing, for which results of a sinflar investigation are given in ref

13、 ereilce 3. The wing variables were dihedral and flap de- flection for e.ach of three locations representing a low- wing, a midwing, and a high-wing monoplane. The effect of the fin was obtained for each variation of each fuse- lage-wing combination. This paper &;ives the lateral-stability characte

14、ris- tics of the wing-fuselage-fin qomb.inations tested and the variations with yaw in rolling-nonent, yawing-moment, and side-force coefficients for the combinations, APPARATUS AND MODELS The tests were nnde in the N.A.C.A. 7- by lo-foot wind tunnel with the regular-6-component balance. The closed-

15、throat tunnel is described in reference 5 and the bnlaace is described in- reference 6. Figure 1 is .a drawing of the laninated mahogany rCodc1, The rectangular wing was usedf.or the tests re- ported in reference 3. The tip plan forn of the wing is composed of two quadrants of similar ellipses. The

16、N.R.C.A. 23012 profile is maintained to the end of the wing and, in elevation, the naxinun upper-surface section ordinates are in one plane. The wing was sot at 0 in- cidonce in all positions. The fuselage is circular,in Provided by IHSNot for ResaleNo reproduction or networking permitted without li

17、cense from IHS-,-,-N.A.C.A .Technical Note ,No. 730 3 - cross section and was made from the dinonsions given in reference 7 for the circular fuselage. The fin was nade to the N.A.C,.A. 0009 section and, in plan forn, fs.represontative of the fins used on the average airplane. The arca of the fin is

18、45 square inches; the ratio of fin area to wing area is 0.08; the aspect ratio of tha fin is 2; and the distance from the assumed position of the center of gravity of the mo.del to the trailing edge of the fin is 0.455 times the w.Jng * span. The 20-percent-chord split flap, nade of l/16-inch steel

19、plate,is attached to the wing at an angle of 60. and extends over 60 percent of the span at the center section. For the nidwing and the high-wing positions, the center section of the flap was cut away to allow for the fuselage. The gap between the flap and the fuselage _- was sealed for all conditio

20、ns. TESTS The fuselage was tested alone and in combination with the wing as a high-wing, a midwing, and a low-wing mono- plane. For each wing position, the combination was tested with Oo and 50 dihedral and with the split flap deflected 0 and 60. Yor all fuselage-wing conbinations, tests were made b

21、oth with and without the fin. Ivery model conbination was tested,at angles of yaw of O“, So, *2, f3O, f5O, 7O, *loo, and 15O. At each angle of yaw, tests were nade at angles of attack of -5, 5O, and 15O for a flap deflection of Oo and at angles of attack of -loo, O“, and loo for a flap deflection 60

22、 .- Tests were also made for all model conbinations at 2O in- - tervals of angle .of attack from -loo to the stall at yaw angles of -50 and 50. The tcsts.were nade at a dynamic pressure of 16.37 pounds per square foot, which corresponds to an air speed of about 80 miles per hour under standard cond

23、itions. The testReynolds Number was about 609,000 based on the chord of the wing. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-N.ACd. Tedhnical Rote no. .730 RESULTS The data, with prinps toindicato w.in,d axes, are giv- en in standard nondinons

24、ional. hb-efficient form. The coef- ficients for the fuselago are based on the wing dimonsioqs. cy 9 qs C, where Yt Lr, IS, cl.1 v, P* s, b, an d a C I w , lateral-forco*,c.oefficient (Yl/qS). rolling-moment coefficient (Ll/qSb). yawing-moment coefficient (Nl/qSb), is lateral force rolling moment. y

25、awing monont. dynamic pressure (l/2 pV2). tunnel-air velocity. air density. wing area. wing.span. is angle of attack. chord. angle of yaw, degrees (positive when the is yawed to the right). dihedral angle of plane of section chord exclusive of tip portion. flap deflection , modei lines The forces an

26、d th,e.mononts have been gfven with re- spect to the wind-tunnel system of axes that intersectin the made1 at the cent-or-.of-gravity locption ehown in fikr ure 1. The lateral.for.ce, the ro1,lfn.g moment, and the yaw- ing non figures 2 to 7 are sample plots. The stability characteristics, and (z, g

27、 , : *g“ were obtained by peas- . . - uring at zero yaw the slope of the curves of the coeffi- cients against angle of yaw. These values are-given as tailed points in figures 8 to 20. The values given as untailed points were obtained fron data ncasured-at ql = *5O (a variable) by assuning that the c

28、oefficients had a straight-line variation for the range from 5 to b5O yawa The values obtained by this nethod are within the practical linits of accuracy for the range of low angles of attack. For sore cases at the higher angles of attack, the values are considerably different when there are large d

29、opartures from the straight-line variation of the coeffi- cients with yaw and when breaks occur in the curves as shown in figures 2to7. The breaks in the curves sople=- tines ilidicate that-the wing is partly stalled for sone angles of yaw For angles of attack near the stall, the variation of the st

30、ability characteristics with angle of attack is expected to be discontinuous depending upon the _ critical nanner in which the wing stalls. The variation of the stability characteristics with wing location-is shown for a condition of uedium angles of attack in figures 21 and 224 The values for the w

31、ing alone given in figures 8 to 22 were obtained fron reference 3 and were converted to the center-of gravity of the aonplete nodel. DISCUSSIOB The tes no rudder; deflections were possible. Wingand fuselage.- *The value of the change of dC/dqt with r varied considerably for difPerent wing- fuselage

32、combinations, although the averago was about the sane as that obtained fur wings alone! 0.00021 per degree (reference 3). In general, i the mount, however, was mall. No regular change in dCy/dJI? with r was shown. The values of dC/dQt usually decreased algebrai- cally as the wing was moved fron the

33、high to the low posi- tion, as is shown in figures 21 and 22. With the wing in the nidposition, the effect on dC1 fd of wing-fuselage interference was snail, being equivalent to an increaee in dihedral of loss.than lo for nost of the angle-of-attack range. With the wing in the high position, the int

34、orfer- ence was equivalent to an increase in effective dihedral of about 30 to 50. In tho low position, the interference affect changed considerably with angle of attack. The effect on dCn/d$ of wing-fuselage interference was small and acted to reduce the instability of the fuse- lage. The interfere

35、nce generally increased as the wing was novad from the high to the low,position. The values of dCyl/dl/r were quite appreciably af- fected by flap deflection. With the wing in the nidgosi- tion with zero dihedral and flaps undoflocted, the wing- fuselage interference was negligible but increased whe

36、n the wing was noved to either the high or the-low position and also when the dihedral was increased. With. the flaps deflected, the interference effects were nogative, tending to reduce the lateral forcei. The nagnitude of tho inter- fcronce was greatest when the wing was in the nidpositioa, the re

37、sultant lateral force being approxinafely zero. This characteristic cannot be explained, but it nay have some effect on the sideslipping characteristics with dif- ferent flap deflections.,. , Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-* +J N.A.C

38、A. Tochnicnl Note NO. 730 7 Fib and fuselam?.- The characteristics of the fuse- lage-fin conbination are shown in figure 8. The value of the increase in dCrl is slightly larger than would - be expected of an airfoil of the smc aspect ratio as the fin but without the fuselage interference (reference

39、 8). The change in dCt/dQl with angle of attack is of the order expected fron dCy./dv and tho vertical-tail po- sition; the change in dC,t/df produced by thc,fin is somewhat less than would be expected fron the change in dC$/d$ and the tail length. Wing, fuselage, und fin.- The contribution of the f

40、in to dC*/d$ is equivalent to an increase in dihedral of fron lo to 4O (figs. 8, 21, and 22). Although of interest in relation to inherent stability, this contribution should not be counted when dihedral for naneuvering is considered because a novenent of the rudder to obtain sideslip will act to co

41、unteract this effective dihedral. The effoctivecess of the fia in changing dC,/dq and dC,*/dV was dependent upon wing position, angle of attack, and flap angle. Moving the wing downward on the fuselage increasod the effectiveness of the fin,rthe naxf- nun effectivoncss being greater than that of the

42、 fin and the fuselage without the wfng. No explanation of thisre- sult is forthconing at present. A nunber of possibilities were considcred,such as angle of attack, wing wake, influ- ence of the wing-tip vortices, and turbulence effects. These effects, however, should vary with angle of attack and n

43、o such large variation was shown by the results. CONCLUSIONS t 1. Tho value of the wing-fuselage fnterfcrence affect on dC VW, the slope of the curve of rolling-nonent - . coefficient against yaw, was mall whert the wing was nount- ed in the nidposition, being equivalent to an increase in. effective

44、 dihedral of about lo or less. . 2. When the wing was shifted to the high position, the wing-fuselage interference on dC2 */a* increased, befng equivalent to an increase in dihedral of the order of 40; with the wing.in the low position, the interference varied considerably with allele of attack. Pro

45、vided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 M.A.C.A. Technical MotcBo. 730 3. The change in dQ/d$* for a given change in wing dihedral was gmersJ.ly not greatly affoctad by the wing-fuselage interference. . h- 4. With flaps-deflected., tho wing-fu

46、selage interfer- enco appreciably reduced the lateral force; without flaps, this iatorferonce was negligible. 5. The offectivencss of the fin on dCn/d$*, the slope of the curve of ynwing-nonent t against Yaw I increased as the wing was novcd downward on the fu- solage, the nnximun offectfveness bein

47、g greator than that of the fuselage-fin conbination and occurring with flaps deflected. 6. Moving the wing downward on tha fuselage generally tended to increase the effaotiveness of the fin on dCl/d$ and on dCy*/dbf. the slope of the curve of lateral-force coefficient against- yaw. Langley MenoriaL.

48、Aeronautical Laboratory, Wational Advisory Connittea for Aeronautics, Lmgley Pield, Va., Augus.t 22,.1939. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c 3 .J W.A.C.R. Technical Uote Ho. 730 9 RCFEEEBCE s 1. Zinneman, Char106 II.: An Analysis of L

49、ateral Stabil- it;- in Power-Off Elic;ht with Charts for Use in De- sign. E.R. Ho. 589, N.A.C.B., 1937. 2. Shortsl, Joseph A.: Zffect of Tip Shape and Dihedral on Lateral-Stability Characteristics. T.B. No. 548, B.R.C.A, 1935. 3. Danbcr, M. J., and Rouse, 11. 0.: Wind-Tunnel Investi- gation of Effect of