NASA NACA-RM-L53K09-1954 Measured and estimated lateral static and rotary derivatives of a 1 12-scale model of a high-speed fighter airplane with unswept wings《带有非扫掠机翼的高速战斗机1 12比例模.pdf

《NASA NACA-RM-L53K09-1954 Measured and estimated lateral static and rotary derivatives of a 1 12-scale model of a high-speed fighter airplane with unswept wings《带有非扫掠机翼的高速战斗机1 12比例模.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L53K09-1954 Measured and estimated lateral static and rotary derivatives of a 1 12-scale model of a high-speed fighter airplane with unswept wings《带有非扫掠机翼的高速战斗机1 12比例模.pdf（25页珍藏版）》请在麦多课文档分享上搜索。

1、 . , s RESEARCH MEMORANDUM HIGH-SPEED FIGHTER AIRPLflNE WITH UNSWEPT WINGS By James L. Williams Langley Aeronautical Laboratory “ LangLey Field, Va. NATIONAL ADVISORY COMMITTEE - FOR AERONAUTICS WASHINGTON Jaw 11, 1954 CON F I DE-NTt AL Provided by IHSNot for ResaleNo reproduction or networking perm

2、itted without license from IHS-,-,-IL . c WA RM L53K09 NATIONAL ADVISORY COMMITTFtE FOR AERONAUTICS DERIVATIVES OF A 1/124CALE MODEL OF A HIGH-SmD FIGHTER AlRpLAwE By Jmes L. Willlams SUMMARY A law-speed investigation was made in the Langley stability tunnel in order to determine the lateral static

3、and rotary derivatives of a l/U-scale model of a high-speed fighter airplane. The experimental results obtained through the complete angle-of-attack range are pre- sented primarily for reference purposes. However, a detailed compari- son at three angles of attack of the lateral static and rotary der

4、iva- tives estimated by currently available methods with the experbnental lateral static and rotary derivatives is made. In general, the vertical-tail contributions to the static and rotary derivatives could be estimated with a good degree of accuracy. The estimated wing- fuselage-combination deriva

5、tives, however, were not in good agreement with the measured vdues. The lack of better agreement of the esti- mated and measured derivatives of the wing-fuselage combination may be caused by the interference of the thick wing roots at the wing-fuselage juncture whlch could not be accounted for by th

6、e methods employed and the inability to calculate readily the fuselage-alone contribution to certain of the stability derivatives. Several methods are available for estimating stability derivatives of airplanes (for example, see ref. I); however, these methods do not account well for the effect of u

7、nusual airplane geometry on the sta- bility derivatives. This deficiency often results in a poor predic- tion of the dynamic stability characteristics of the airplane. A similax situation appears to exist for the high-speed fighter air- plane employed in this investigation since the damping of the l

8、ateral Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-oscillation of this airplane could not be calculated in one investiga- tion with the accuracy desired by using estimated stability derivatives (ref. 2) although better agreement was obtained in a

9、nother investigation (ref. 3). The purpose of the present investigation, which was made in the Langley stability tunnel, was to obtain the low-speed lateral static and rotary stability derivatives of a 1/12-scale model of a high-speed fighter airplane with unswept wings and ta compare the experiment

10、al stability derivatives with the derivatives estimated by current methods for the wing-fuselage canbination, the vertical- and horizontal-tail combination, and the complete model. In addition, since a large dif- ference existed between the static lateral stability derivatives pre- sented herein and

11、 the unpublished derivatives obtained in previous tests of a sting-supported model, a few tests were made to determine the effects on the static lateral stability derivatives of a fuselage modification sFmilar to that necessitated for sting-mounting. This modification consisted of an increase in the

12、 cross-sectional sea Of the rear portion of the flrselage under the vertical tail. SYMBOLS AND com1cIENTs The data presented herein are in the form of standard NACA coeffi- cients of forces and moments which are referred to the stability system of axes (fig. 1) with the the wing mean aerodynamic dir

13、ections of the forces, in figure 1. The symbols origin at the projection of the 0.23 point of chord on the plane of symmetry. The positive moments, and however, R detailed comparison at three angles of attack of the lateral static and rotary derivatives eathated by currently available mthods with th

14、e experhentd lateral static and rotary derivatives is presented in figure 9. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53K09 7 Effect of Fuselage Modification on Static Lateral Derfvatives A comparison of the static lateral stability

15、 derivatives. obtained in the present Investigation (fig. 6) with some unpublished results obtained at a Mach number of 0.4 indicates larger differences than would be expected to be caused by Mach rider effects alone. The model employed Fn the tests at a Mach number of 0.4 was sting supported with t

16、he sting enter- the rear portion of the fuselage. This axrangement necessitated a revision to the fuselage aftersection because of the fuselage shape (see fig. 2). In order to determine the inportance of this modification on the static lateral characteristics, the fuselage of the mdel used in the st

17、ability-tunnel investigation was modified (see fig. 4) to slmlate this sting-swported model. he derivatives resat- from tests of this arrangement are presented in figure 6. The values of the modified-fuselage derivatives are in good agreement with the unpublished derivatives obtained at a Mach numbe

18、r of 0.4. The fuselage modification produced a large increase in and c (see fig. 6). These changes are believed to result fromthe increase in end-plate effect and the induced sidewash of the fuselage on the vertical tail as the fuselage size under the tail is increased. The use of values of Cy from

19、the sting-supported-model tests in esti- mating c would give erroneous results, of course. cypv nPV BV nrv It appears, therefore, that in testing models similar to the model of the present investigation an effort should be made to -1nimlze fuse- lage modifications. If the effect of fuselage modLfica

20、tion on the test results cannot be evaluated by experimental or theoretical methods, then it may be necessazy to mount the model on whg-tip stings which would require, of course, the determfnation of tares. Estimation of Derivatives and Caqarison With Experiment Wing-fuselage contribution.- The proc

21、edure employed for estimating the -fuselage combination derivatives except as noted for Cnr and C was to estimate the wing and fuselage derivatives separately and to add them algebraically. The derivatives of the basic wing plan form and fuselage were obtained from the following sources: “p Provided

22、 by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM L53K09 i- Derivative “ Reference 8 9 10 II 12 13 he lift and drag data of the wing-fuselage combination (fig. 5) were used with the methods of references 8 and 10 to estate Cnr and Cnp and no addition

23、al increments were added for the fuselage since it is indirectly accounted for in this manner. The effect of wing dihe- dral on Cyp was determined from reference 14 and on C2 and C 2, from references 15and 16, respectively. The effect of wing position on the sideslip derivatives was determined from

24、reference 17 assuming a low-wing position. The mutual-interference effects of the wing-fuselage combination have not been accounted for in these calcula6ions since all the currently available interference data have been determfned for sfmple bodies of revolution only (refs. U, 12, and 18). There was

25、 no air flow through the wing ducts. It is believed that for this case the flaw through the ducts has no appreciable effect on the stability derivatives. P In general the esthated derivatives of the wing-fuselage combination are only in fair agreement with the measured derivatives (see fig. 9) . It

26、appears that this lack of better agreement could be caused by a large interference effect of the thick wing roots at the wing-fuselage juncture which cannot be accounted for by the currently available methods, and the inability to calculate readily 821 accurate fuselage-alone contribution to some of

27、 the stability derivatives. Evfdently, more information on the mutual-interference effects for wing-fuselage combinations other than simple bodies of revolution is needed. Vertical-tail contribution.- The vertical-tail increments to the stability derivatives were calculated by means of the equations

28、 given Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-*. . in reference 19. The lift-curve slope was determined from ref - I erence 20 for an effective aspect ratio determined fran references IL i and 17. In the estimation of the yawing and rollfrtg

29、 derivatives of the “- vertical hence this effect was not accounted for in this paper. Complete model.- The estimated derivatives for the wing-fuselage combination and tail group were summed to obtain the complete model derivatives. The agreement between the estfmated and measured deriva- tives was

30、generally good. The poor weement between certain estimated and measured complete-model derivatives is obtafned as a directconse- quence of the Inability to estbte the wing-fuselage contribution to the derivatives. A low-speed investigation was made in the Langley stability tunnel in order to determi

31、ne the lateral static and rotary derivatives of a 1/12-scale model of a high-speed ffghter airplane with unswept wings. The experimentally determined derivatives “Ough the complete angle-of- attack range are presented primarily for reference purposes. However, a detailed comparison at three angles o

32、f attack of the lateral static and rotrary derivatives estimated by currently avaflable methods with the experimental derivatives is presented. In using current methds to estimate the derivatives of the airplane it was found that in general the tail contribution to the lateral static and rotary deri

33、vatives could be estimated with a good degree of accuracy. The estimated wing-fuselage-combination derivatives, however, were not in good agreement with the measured values. This lack of better agreement may be caused by the Interference of the thick wing roots at the Provided by IHSNot for ResaleNo

34、 reproduction or networking permitted without license from IHS-,-,-10 wing-fuselage juncture which could not be accounted for by the methods employed, and the inability to calculate readily the fuselage-alone con- tribution to certain of the stability derivatives. Langley Aeronautical Laboratory, Na

35、tional Advieory Ccsmnittee for Aeronautics, Langley Field, Va., October 26, 1953. c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L53K09 .: - REFERENCES fi. Campbell, John P., and McKinney, Marion 0. : Sumnary of Methods for Cdculating Dyna

36、mic Lateral Stability and Response and for Estimating Lateral Stability Derivatives. NACA Rep. 1098, 1952. (Supersedes NACA TN 24-09. ) 2. Crane, H. L., Beckhardt, A. R., and Matheny, C. E.: Flight Measure- ments of the Lateral Stability and Control Chrscteristics of a High-speed Fighter Airplase. N

37、ACA RM L52Bl4, 1952. 3. Heinle, Donovan R., and McNeill, Walter E.: Correlation of Predicted and Experimental Lateral OscKIhtion Characteristics for Several Airplanes. WA RM 5206, 1952. 4. MacLachlan, Robert, and Letko, William: Correlation of Two Experi- mental Methods of Determining the ROUT Chara

38、cteristics of Unswept Wings. NACA TN 1309, 1947. 4 Bird, John D., Jaquet, Byron M., and Cowan, John W. : Effect of Fuse- lage and Tail Surfaces on Low-Speed Yawing Characteristics of a Swept-Wing Model As Determined in Curved-Flaw Test Section of the Langley Stability Tunnel. NACA TN 2483, 1.951. (S

39、upersedes WA IIM L8G13.) i$ Silverstein, Abe, and White, James A. : Wind-Tunnel Interference With Particular Reference to Off-Center Positions of the Wing and to the Damwash at the Tail. NACA Rep. 547, 1.936. YJ. Gillis, Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.: Charts for De-bermi

40、bLn.13. Jet-Boundary Corrections for Cmrplete Models in the 7- by 10-Foot Closed Rectangular Wind Tunnels. NACA WR L-123, 1945. (Formerly NACA ARR L5G3l. ) 8. Toll, Thomas A=, and Qwijo, M- J. : Approxhate Relations end Charts for Lm-Speed Stability Derivatives-of Swept Wings. NACA TN 1581, 1948. Go

41、odman, Alex, and Adair, Glenn E.: Estbtion of the Demging in Roll 0 of Wings Through the Normal Flight Range of Lift Coeff fcient . NACA TET 1924, 1949. 10. Goodman, Alex, and Fisher, Lewis R.: Investigation at Low Speeds of the Effect of Aspect Ratio and Sweep on the Rolling Stability Derivatives o

42、f Untapered Wings. WLCA Rep. 968, 1950. (Supersedes NACA TN 1835.) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 - NACA RM L53K09 %E I . . DlMENsIONS AWD CHARACTERISTICS OF MODEL L . Wing: Airfoil section at fold (fold at 0.427 b/2) . Airfoil se

43、ction at theoretical tip . TOW area. +. sq ft Mean aeroaynamic chord. E. ft Root chord (trailing edge extended). ft . TLp chord. ft Sweep of leading edge. deg . Incidence. deg span. bW. ft . Aspect ratio At theoretical root chord . At theoretical tip chord . At wing fold root chord Dihedral. deg NAC

44、A 651-u2 . NACA 63-209 2.04 3.47 0.613 0.762 0.361 5.90 0 42 42 42 3 Horizontal tail : Airfoil section U-percent-thick WA 63-series kea. SH. eq ft . 0. 485 span. bH. ft 1.51 Root chord. ft . 0.402 Tipchord. ft 0.243 Sweep. leading edge. deg 8.45 Area ratio. SH/SW 0.29 Vertical tail: Airfoil section

45、U-percent-thick WA 65-series Total area. %. sq ft 0.328 Root chord. Ft . 0.673 Tip chord. ft 0.268 Spas. h. ft 0.694 Sweepback. leading edge. deg 24.50 Tail length. distance from kenter of gravity to 2. ft 1-42 Tail height. perpendicular distance from center of gravity Area ratio. +/% 0.161 Fuselage

46、 length. 2% . 3.33 E 4 . to EV. ft . 0.434 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 C FT .r . 0 4 8 0 4 8 0 4 8 Angle of uftuck, E, deg (c) Yawing derivatives. Figure 9. - Concluded. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-