NASA NACA-TR-1224-1955 Effects of wing position and fuselage size on the low-speed static rolling stability characteristics of a delta-wing model《机翼位置和机身尺寸对三角形机翼模型低速静态旋转稳定特性的影响》.pdf

《NASA NACA-TR-1224-1955 Effects of wing position and fuselage size on the low-speed static rolling stability characteristics of a delta-wing model《机翼位置和机身尺寸对三角形机翼模型低速静态旋转稳定特性的影响》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-TR-1224-1955 Effects of wing position and fuselage size on the low-speed static rolling stability characteristics of a delta-wing model《机翼位置和机身尺寸对三角形机翼模型低速静态旋转稳定特性的影响》.pdf（34页珍藏版）》请在麦多课文档分享上搜索。

1、:. ,. 3 ,- I, i By ALEX GOOQMAN ah D ,. I . : 1% I. / ,. .- I i I “, , I _ ; , , I .,- . . ._ r I -. . , . ., I / -: . * . Provided by IHSNot for Resale-,-,-i TECH LIBBARY KAFB, NM REPORT 1224 EFFECTS OF WING POSITION AND FUSELAGE SIZE ON THE LOW-SPEED STATIC AND ROLLING STABILITY CHARACTERISTICS OF

2、 A DELTA-WING MODEL By ALEX GOODMAN and DAVID F. THOMAS, JR. Langley Aeronautical Laboratory Langley Field, Va. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSNational Advisory Committee for Aeronautics ,Headquarters, 1512 H Street NW., Washington 26,

3、D. 0. Created by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. ode, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25,1948. The members a

4、re appointed by the President, and serve as such without compensation. JEROME C. HUNSARER, SC. D., Massachusetts Institute of Technology, Chairman LEONARD CARMICHAEL, PH. D., Secretary, Smithsonian Institution, Vice Chairman JOSEPH P. ADAMS, LL. B., Vice Chairman, Civil Aeronautics Board. DONALD L.

5、PUTT, Lieutenant General, United States Air Force, ALLEN V. ASTIN, PH. D., Director, National Bureau of Standards. Deputy Chief of Staff (Development). PRIGSTON R BASSETT, M. A., Vice President, Sperry Rand Corp. DONALD A. QTJARLES, D. Eng., Secretary of the Air Force. DETLEV W. BRONK, PH. D., Presi

6、dent, Rockefeller Institute for ARTHUR E. RAYMOND, SC. D., Vice President-Engineering, Medical Research. Douglas Aircraft Co., Inc. THOMAS S. COMBS, Vice Admiral, United States Navy, Deputy Chief of Naval Operations (Air). FREDERICK C. CRAWFORD, SC. D., Chairman of the Board, Thompson Products, Inc.

7、 FRANCIS W. REICHELDERFER, SC. D., Chief, United States Weather Bureau. LOUIS S. ROTHSCHILD, PH. B., Under Secretary of Commerce for Transportation. RALPH S. DAMON, D. Eng., President, Trans World Airlines, Inc. JAMES H. DOOLITTLE, SC. D., Vice President, Shell Oil Co. CARL J. PFINGSTAO, Rear Admira

8、l, United States Navy, Assistant Chief for Field Activities, Bureau of Aeronautics, NATHAN F. TWINING, General, United States Air Force, Chief of Staff. HUGH L. DRYDEN, PH. D., Director JOHN W. CROWLIY, JR., B. S., Associate Director for Research JOHN F. VICTORY, LL. D., Executive Secretary EDWARD H

9、. CHABERLIN, Executive Ojicer HENRY J. E. REID, D. Eng., Director, Langley Aeronautical Laboratory, Langley Field, Va. SMITH J. DEFRANCE, D. Eng., Director, Ames Aeronautical Laboratory, Moffett Field, Calif. EDWARD R. SHARP, SC. D., Director, Lewis Flight Propulsion Laboratory, Cleveland, Ohio WALT

10、ER C. WILLIAMS, B. S., Chief, High-Speed Flight Station, Edwards, Calif. II Provided by IHSNot for Resale-,-,-REPORT 1224 EFFECTS OF WING POSITION AND FUSELAGE SIZE ON THE LOW-SPEED STATIC AND ROLLING STABILITY CHARACTERISTICS OF A DELTA-WING MODEL . . By ALEX GOODMAN and DAVID F. THOMAS,JR. SUMMARY

11、 An investigation was made to determine the eye2 2 per radian 217 Czp = 25 per radian ab 217 WY, cypv=- a pb per radian 217 A,C, A,C, increments of coefficients caused by Al G- that is, AIGp, A that is, A2CyB= (Cy,w+F+,- %lr.-+F) - (%+9YBF) (increments of coefficients caused by mutual interference o

12、f fuselage and vertical tail; that is, A.Bv Subscripts: W isolated wing 2 isolated fuselage or body T7 isolated vertical tail WF wing-fuselage combination r root CT component due to sidewash APPARATUS AND MODELS The tests of the present investigation were made in the 6-foot-diameter rolling-flow tes

13、t section of the Langley sta- bility tunnel. This section is equipped with a motor-driven rotor which may be used to impart a twist to the airstream so that a model mounted in the tunnel is in a field of flow ,.4.50 max diam. k-27.00 -21.50 - I kc,= 31.6 I I V-54.00 FIGURE 2.-Dimensions of the compl

14、ete models. All dimensions are in inches. similar to that which exists about an airplane in rolling flight (ref. 3). Details of the wing, fuselages, and vertical tail surfaces and the relative locations of the wing and vertical tails with respect to the fuselages are given in figure 2. The various w

15、ing positions, fuselage sizes, and vertical-tail sizes will be referred to herein by the following designations: We-_,-_-_-_-_-_- Midwing -_-_-_-_-_- Highwing IV,- _ -_-_- _ -_-_- Lowwing F1_-_-_- Small fuselage FZ- _-_-_-_-_- Medium fuselage F3 _-_ -_-_-_- _ -_- Largefuselage VI _ -_-_- _ Small ver

16、tical tail V, _ -_-_- _ - _ Mediumvertical tail V, _ -_- _ - _ Large vertical tail A list of the pertinent geometric characteristics of the various component parts is given in table I. TABLE I.-PERTINENT GEOMETRIC CHARACTERISTICS OF MODELS Fuselage: Fl F3 F3 Length,in. -_- ._ - 54. 0 54. 0 54. 0 Max

17、imum diameter, iu. _ . _ - - - - _ 4. 5 6. 0 9. 0 Finenrssratio-_-_._- 12. 0 9. 0 6. 0 Body-size ratio, d/b,v . . . - _ 0. 123 0. 165 0. 246 Volume, cuin. -.-_- 545 990 2,200 Side area, sq in.-.-.- 186 252 370 Wing: Aspectratio-.-. 2. 31 Taperratio-.-.-. 0 Leading-edge sweep angle, dtlg._ - - _ - _.

18、 _ _. _ _ 60 Dihedral angle, deg _._ - _. - _ _ 0 Twist,deg_-.-.-. 0 NACAairfoilsection _.- _- _ -. 65A003 Area,sqin.-.-. 576. 7 Span, in. _ -. . - . . _. 36. 5 Mean aerodynamic chord, ill. _. . .-. -. . . 21. 1 Rootchord,in_. . - -_.-. . 31. 6 Wing-height ratio for all wing- fuselage combinations,

19、zw/tl_ - - _ _ 0, f 0. 333 Vertical tail: 171 1,; 173 Aspect ratio_-.- _ 2. 18 2. 18 2. 18 Taper ratio-.- _-_ 0 0 0 Leading-edge sweep angle, deg-_ . . 42. 5 42. 5 42. 5 NACA airfoil section _ _ -. _ _. _. 65-006 65-006 65-006 Area, sq in. - _ -_ 39. 2 48. 3 66. 0 Span, in. -_ 9. 25 10. 25 12.00 Roo

20、t chord, in. -_- 8. 50 9. 40 11.00 Mean aerodynamic chord, in. _ _ _ _ 5. 67 6. 25 7. 35 Tail length, in. - _ -_ 21. 5 21. 5 21. 5 Arearatio, Sv/Sw- 0. 068 0. 084 0. 115 Tail-length ratio, Iv/b,“- _- - - - _ _ _ _ 0. 59 0. 59 0. 59 The complete models used for the present investigation were designed

21、 to permit tests of the wing alone, the fuselages alone, the wing-fuselage combinations (with the wing at three different vertical positions rela.tive to the fuselage), or the fuselage in combination with any of the three vertical tails with or without the wing. The fuselages used in the investi- ga

22、tion had fineness ratios of 6, 9, and 12 and were bodies of revolution having parabolic-arc profiles and blunt-tail ends. The wing was a 60” delta wing of aspect ratio 2.31 and had an NACA 658003 profile in sections parallel to the plane of symmetry. All the triangular vertical tails had an aspect -

23、 - _ - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS4 REPORT 1224-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ratio of 2.18,42.5 sweepback of the leading edge, and NACA lateral force of the isolated vertical tails as well as the tails in 65-006 profile

24、s in planes parallel to the fuselage center line the presence of the fuselages were obtained by means of an and differed only in area, (See table I.) Ordinates for the electrical strain gage. Photographs of two of the configura- NACA 658003 and 65-006 sections and for the fuselages are tions tested

25、are presented as figure 3. The wing was set at 0 given in tables II and III, respectively. All parts were incidence with respect to the fuselage center line in all constructed of mahogany. positions. TABLE II.-ORDINATES FOR NACA 658003 AND 65-006 AIRFOILS Station and ordinates in percent airfoil cho

26、rd NACA 65A003 NACA 6546 0 3: 1. 25 2.50 5.00 7. 50 10.00 15.00 20.00 25.00 30.04 35.00 40.00 45.00 50.00 Ei 65.00 70.00 75.00 80.04 2:E 95.00 100.00 Ordinates ” .234 ,284 .362 .493 :X ,912 1.097 1.23G 1.342 1.420 1.472 1.498 1.497 1.465 1.402 1.309 1.191 1.053 .897 727 : 549 ,369 :% L. E. radius: 0

27、.057 L. E. radius: 0.240 Station Ordinates 0 :% 1.25 2.50 5.00 7.50 10.00 15.00 20.00 E:Z 35. oil 40. cm 45.00 50.00 55.00 60.00 it: 75:OG 80.00 85.00 90.00 95.00 100.00 0 .476 ,574 :6i 1.310 1.589 1.824 2.197 2.482 2.697 2.852 2.952 2.998 2. 983 2.930 2.741 2.518 2.246 1.935 1.594 1.233 .865 .510 0

28、 195 TABLE III.-FUSELAGE ORDINATES I Station, s/l 0 .006 .oQ9 .015 .030 :E .12n :G .3Oil .360 .400 ,420 .480 .540 :E .72O .780 .840 :% 1. ooo Fl 0 .0013 .OQ19 .W32 .0059 .0115 .0167 .0213 .0291 : %i .0413 .0417 .0417 .0413 .0406 :%I : E .0304 .0270 .02.33 .0208 Ordinate, z,ll F2 0 .0017 .0024 .0041

29、0080 .0154 .0222 .0284 .0387 .0467 .052O .0550 .0556 : L% .0541 2% .0476 .0443 .0406 : E .0276 1 F8 0 .0024 .0037 .C%l .Ol!m .0232 .0333 .0426 : -Vertical tail V, Strain-gagedB balance Single-strut support- LATERAL-STABILITY CASE Interference increments.-By using a method analogous to the one employ

30、ed for the longitudinal-stability case, the static-lateral-stability derivatives of the present complete configurations can be expressed as (see ref. 2) The interference increments can be obtained from the test results in a manner analogous to that used for the longitudinal-stability case. For examp

31、le: FIGURE 4.-Sketch of vertical-tail mounting for determining isolated- vertical-tail results. CORRECTIONS Approximate corrections, based on unswept-wing theory, for the effects of jet boundaries (ref. 5) have been applied to the angle of attack and drag coefficient. The data are not corrected for

32、blocking, turbulence, or support-strut inter- ference. METHODS OF ANALYSIS A5c%= (cy8F,-c%F) - (%,F 00) where (Cy8v)F is the vertical-tail contribution to Cy, in the presence of the fuselage. Equations (9) and (10) when added together result in equation (8). Vertical-tail efficiency factors-The vert

33、ical-tail contribu- tion to the lateral-stability derivatives as affected by the wing-fuselage interference can, for example, be expressed as The results of the present investigation are analyzed in terms of the individual contributions of the various parts of the models to the aerodynamic character

34、istics and to the more important interference effects. (11) LONGITUDINAL-STABILITY CASE where (CY8”) WF is the vertical-tail contribution to CyO in the presence of the wing-fuselage combination. Similarly, the contribution of the vertical tail to Cy8 as affected by the fuselage interference can be e

35、xpressed as Solving equations (11) and (12) for the efficiency factors gives, for wing-fuselage interference, In accordance with conventional procedures (for exam- ple, as outlined in ref. 6), the lift and pitching-moment co- efficients for the present complete configurations can be ex- pressed as C

36、L= L,+L+A,L 0) cm= %z,+ (%ql-An=Gn,+p- (cm,+mJ (4) I . ./ _ ,-:.L . _ ,mmm.I;. -:-1.:e- -L .I: .- ; . _,. , , ,. . , . . ,-_. . . . . . . . . . _ -. (14) A3cysE (Cy how- ever, the increase becomes less as the wing is moved from the low to the high positions. The interference is, therefore, a functio

37、n of the body-size ratio and decreases with a decrease in the ratio. The variation of CL, with body- size ratio and wing-height ratio is presented in figure 9 and illustrates this effect. Also, as can be seen in figure 9, the high-wing configurations attained the highest CL, 351683-56-2 The addition

38、 of a vertical tail to the wing-fuselage com- binations had little effect on the longitudinal stability characteristics. (Compare figs. 10 and 11 with fig. 7.) STATIC LATERAL STABILITY CHARACTERISTICS Wing characteristics-The variations of Cl-, Cn8, and C!, with angle of attack for the 60 delta. win

39、g are presented in figure 12. The derivative C, and C, are generally small for most of the angle-of-attack range. The value of the slope bCZJbCL through a=O” of 0.0047 for this wing is in good agreement with the value of 0.0050 calculated by the method of reference 11. Fuselage Characteristics-The m

40、ain contribution of the isolated fuselages to the static lateral stability characteristics is an unstable yawing moment throughout the angle-of- attack range (see fig. 13). The magnitude of the unstable yawing moment at low angles of attack is apparently a direct function of the fuselage size. The f

41、uselage characteristics at r=0 are summarized in figure 14. In order that the results obtained may be applied conveniently to arbitrary airplane configurations, coefhcients in terms of fuselage dimensions are needed, This end is accomplished by plotting the quan- tities (C,), $ and (C,), F against f

42、uselage fineness ratio. The quantities plotted, therefore, are effectively a lateral-force coefficient based on fuselage side area S, and a yawing-moment coefficient based on fuselage volume v,. The results presented in figure 14 are compared with the results of reference 12 and the theory of refere

43、nces 9 and 13. . - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS8 REPORT 1224-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS presented in figure 15 as ( CYBV)p where ( CYBV)F is the verti- cal-tail lift-curve slope based on the wing area. The varia- tion

44、of the efficiency factor (ql8)F ,as determined by the procedures explained in the section entitled “Methods of Analysis” is presented in figure 16 with angle of attack. This factor is a direct measure of the induced sidewash at the tail for a=O. The effects of fuselage size on the efficiency factor

45、(7)8)p and the tail lift-curve slope CLaV (based on tail area) are summ.arized in figure 17 for (Y=OO and show an increase in (T) and CLolV as the fuselage diam- eter is increased. The effect of fuselage size could be cal- culated with good accuracy by using a finite-step method such as discussed in

46、 reference 14 and by accounting for the effects of the fuselage by using a method similar to that of reference 15. This method also yields the span loading on the tail. The calculated values are also in good agreement with the experimental results and indicate an increase in stabilizing sidewash at

47、the vertical tail with an increase in fuselage size. (See fig. 15.) The experimental results show a negative lateral force which increases as the fineness ratio is decreased; this result is in good agreement with the results of reference 12. The theory of reference 9, which is based on potential-flow consideration for closed bodies, predicts no lateral force. The theory of reference 13 results in a fair estimation of the fuselage lateral-force coefficient. The experimental results obtained for the directional-stability paramete