NASA NACA-RM-L56D05-1956 Static lateral stability and control characteristics of a model of a 45 degrees swept-wing fighter airplane with various vertical tails at Mach numbers of .pdf

1、.,. * ,?.b ?A- .a. : I, .*; -.-i,“. “;N;h “j _ _ -. .i; - Fq 1 r-“- fJ f-y; :;i , :; j ( ; -. I f ,j ;I 6 . 1-L i.!, i i 1 -.y., Ill,?. ,:,$?gy iAeronwtlca1 Laboratory :, b p* h i” t /i I * ,: - 1 ; I( I .- 1 ;- f”r Langley Field, Va. I; -.I I i 4 j s: I j ,_. , -. t i,:. 9 P 4 +z-,- j i- ,x ,:,- (7

2、 : , 3 T, , _ -3 .r .:i iei :, ! ! 3: ;” 1 T, 71 - . . ; / _ ” - :, _ y / .,/.*.- : y : r. . ,. r _ ,. . . . _. I . “. ,_ , _ “, : . ?. i, !,; i ,. ,“,# : / : ,:,; CLASSIFIED WCUMEW ,* (6 This material contains information affecting the National Defense Of the United states within the mealdng of the

3、 eapiomge laws, Title 18, U.S.C., Sets. 793 and 794, the transmission 01 revektioo ofwhich in my manner to an unauthorized person is prohibited by law. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON June 19, 1956 : , . .,:z ;,I.: : 4, (.d_ I_ Provided by IHSNot for ResaleNo reproduction or n

4、etworking permitted without license from IHS-,-,-NACA RM 5605 slrinnsriiiiiiiiiiiii ,i! 3 1176014372172 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS -5 ,h RESEARCHMEMORANDUM STATIC LATERAL STABILITY AND CONTROL CHARACTERISTICS OF A MODEL OF A 45 SWEPT-WING FIGHTER AIRPLANE WITH VARIOUS VERTICAL TAILS

5、 AT MACH NUMBERS I OF 1.41, 1.61, AND 2.01 By M. Leroy Spearman and Ross B. Rob.inson SUMMARY An investigation has been made in the Langley 4- by b-foot super- sonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 of a model of a 45 swept-wing fighter airplane. The wing had an aspect ratio o

6、f 3.86, a taper ratio of 0.262, and NACA 64(,6)AOO7 airfoil sections in a streamwise direction. Static lateral stability and control charac- teristics were obtained through an angle-of-attack and sideslip range for various combinations of component parts and for the complete model with three differe

7、nt vertical tails of varying sizes and aspect ratios. The majority of the tests were conducted at a Mach number of 1.61, and only limited sideslip results were obtained at Mach numbers of 1.41 and 2.01. Aileron- and rudder-control characteristics were obtained for the complete model at a Mach number

8、 of 1.61 only. The directional stability derivative C, P for the complete config- uration progressively decreased with increasing Mach number and angle of attack until regions of directional instability occurred. Increasing the size of the vertical tail provided increases in C I-9 so that the onset

9、of directional instability was delayed to higher Mach numbers or angles of attack. The lateral and directional control characteristics were essentially constant throughout the angle-of-attack and sideslip ranges. ; Provided by IHSNot for ResaleNo reproduction or networking permitted without license

10、from IHS-,-,-2 INTRODUCTION NACA RM 561x15 A research program has been undertaken in the pngley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a model of a 45O swept-wing fighter airplane in the Mach number range from 1.41 to 2.01. The static longitudinal sta

11、bility and control char- acteristics at Mach numbers of 1.41, 1.61, and 2.01 are presented in reference 1. Effects of various external stores on the longitudinal and lateral characteristics at Mach numbers of 1.61 and 2.01 have been determined but the results are unpublished. Flight-test results of

12、a similar configuration are presented in reference 2. The present paper contains the static lateral and directional sta- bility and control characteristics at Mach numbers of 1.41, 1.61 and 2.01. The Reynolds numbers of the tests based on the wing mean geometric chord varied from 1.40 x 106 to 1.16

13、x 106. Results were obtained for the model equipped with three different vertical tails of varying area and aspect ratio. c0EFF1cIENTS AND SYMBOLS The lift, drag, and pitching-moment coefficients are referred to the stability axis system (fig. l(a). The lateral-force, yawing-moment, and rolling-mome

14、nt coefficients are referred to the body axis system except where noted (fig. l(b). The center of moments of the model was at a longitudinal position corresponding to the X.5-percent station of the wing mean geometric chord. The coefficients and symbols are defined as follows: CL lift coefficient, -

15、F whereas for the tests at M = 1.41, a manually adjustable sting was employed. TESTS, CORFUETIONS, AND ACCURACY The conditions for the tests were as follows: Machnumber 1.41 1.61 2.01 Stagnation temperature, OF . . . . 100 100 100 Stagnation pressure, lb/sq in. abs. 6 6 6 Stagnation dewpoint, 9i . .

16、 . . . . -20 -20 -25 Reynolds number, based on E . . . _ 1.40 x lo6 1.34 x 106 1.16 x lo6 Tests were made through the following approximate angle ranges: I ti I 1.41 1.61 2.01 Variable angle range, deglConstant angle, degl P = -8 to 15 a = 5.1 a = -8 to 16 P = -4.8, o = -20 to 20 a= 0 to 15 f.1, 8.3

17、, P = 0 to 12 a = 20.9 ; = 0 to 20 a= * 0 to 15 a = i.1, 8.2 The model angle was corrected for the deflection of the balance and sting under load. Base pressure was measured in the plane of the model base. By equating the base pressure to free-stream static pressure, the drag values have been adjust

18、ed so that the base drag was zero for all configurations. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM 5605 Maximum probable errors in the individual measured quantities are as follows: CL Q. Cm Cn c2. % a,p,deg it, 6aL, deg . M . - M =

19、1.41 and 1.61 M = 2.01 *0.0044 ;to.o051 *0.0005 +0.0007 0.0017 f0.002l *0.0003 *0.0003 f0.0002 f0.0002 f0.0020 f0.0020 f0.2 20.2 fO.l fO.l fO.O1 *0.015 RESULTS AND DISCUSSION As seen in table II, the basic data are presented in figures 4 to 11; the summary data, in figures 12 to 18; the aileronYcont

20、rol data, in figures 19 t0 2lj and the rudder-control data, in figures 22 to 24. Static Stability Characteristics Directional stability.- The directional stability Cn P for the basic configuration decreases progressively both with increasing angle of attack and increasing Mach number until regions o

21、f undesirably low stability are encountered (see fig. 15). The directional characteristics for the tail- off configuration (fig. 16) are essentially invariant with Mach number and angle of attack and indicate a relatively large unstable moment. This large unstable moment results primarily from the l

22、arge fuselage and the far-rearward moment center. The far-rearward moment center also results in a short tail moment arm and, hence, lessens the ability of the verti- cal tail to provide a stabilizing moment. Consequently, the condition exists where a large percentage of the tail contribution is con

23、sumed in overcoming the instability of the wing-body combination and relatively little tail effectiveness is available to provide a stability margin. Under such conditions, factors that affect the tail contribution, even to a slight degree, begin to assume greater importance. For example, r; the rap

24、id decrease in Cn P with increasing Mach number for the complete configuration is a direct result of the decrease to be expected in the vertical-tail lift-curve slope. In addition, as pointed out inrefer- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,

25、-,-i 3- d i ,I , L NACA RM 561105 7 ence 2, the losses in tail contribution resulting from aeroelasticity might be significant for a full-scale airplane. Increasing the tail contribution through increases in the tail area and aspect ratio, although having little effect on the variations of CnP with

26、Mach number or angle of attack, does increase themagnitude of CnD in such a way that the imminence of directional instability is delayed to higher angles of attack or to high Mach numbers. (See figs. 12, 13, and 17.) The variation of Cn with p for the complete model is rather nonlinear and does, in

27、fact, indicate a reversal in direction which results in the occurrence of unstable yawing moments (fig. 7, for example). This trend is influenced to some extent by the increasing instability of the wing-body conibination and by a nonlinear vertical-tail contribution, and occurs even though the tail

28、contribution continues to increase with increasing sideslip. Increasing the tail size does not remove this non- linear variation of C, with p but does delay the occurrence of the unstable yawing moments to higher angles of sideslip. The presence of the horizontal tail provides a slight increase in t

29、he directional stability at a = 0 either with or without the vertical tail (figs. 6 and lo), but at higher angles of attack this effect reverses. Negative deflections of the horizontal tail provided an increase in the directional stability for the basic configuration at M = 1.61 (fig. p), apparently

30、 because of a transmittal of positive pressures from the upper surface of the horizontal tail to the windward side of the body and vertical tail. The effect of tail deflection is evident at M = 2.01 (fig. 10) but to a lesser degree since a smaller portion of the body and vertical tail are influenced

31、 by the flow field of the horizontal tail as the Mach number increases. Results from other investigations involving configurations having high horizontal tails (ref. 3, for exsmple) indicate an opposite effect in that negative deflections of the horizontal tail cause a decrease in the directional st

32、ability. An interesting feature concerning the effects of the axis system on the interpretation of the data is illustrated in figure 18 where the variation of C 33 with a for the basic configuration at M = 1.61 is presented for both the stability and the body axis systems. The results computed for t

33、he stability axis system indicate less deterioration of directional stability with increasing angle of attack and, in fact, do not indicate any directional instability for the tail-on case, whereas the results computed for the body axis system indicate directional insta- bility above a = 16O. This e

34、ffect results from the transfer of rolling Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM 561105 moment into yawing moment for the stability axis system and can cause an a-ppreciable difference in C, P at the higher angles of attack if the

35、 rolling moments are large and the yawing moments are small. Thus, it is possible that some configuration changes that have a large effect on roll but little effecton yaw (such as wing dihedral) may, if computed for a stability axis system, show an effect on yaw. Effective dihedral.- The variation o

36、f CZR with a for the basic configuration is particularly nonlinear at Mach numbers of 1.41 and 1.61 (fig. 15): it varies from small negative values to small positive values at low angles of attack and increases to relatively large negative values at higher angles of attack. The results at M = 2.01 a

37、re for only a limited angle-of-attack range up to about 8 , but within this range the variation of C2 with a is fairly linear. P For Mach numbers of 1.41 and 1.61, the variation of C2 with a P at low angles of attack is generally positive either with or without the vertical tail (figs. 15 and 16); w

38、hereas for M = 2.01, the variation is negative. This trend toward negative variations of C2 with a for P increasing Mach number is in general agreement with the linear-theory prediction for swept wings having supersonic leading edges (ref. 4). The presence of the vertical tail, of course, provides a

39、 negative increment of C2 P that progressively increases as the tail size increases and progressively decreases as the Mach number increases (fig. 17). Effects of sideslip on longitudinal characteristics.- The lift, longiGZina1 force, and pitching moment vary only slightly with angle of sideslip for

40、 angles of attack up to about 8O (figs. 4, 8, and 11). At a = 15.7O and 20.p“ (fig. S), however, a rapid positive increase of pitching moment with increasing sideslip indicates the possibility of cross coupling of the lateral, directional, and longitudinal motions. This cross-coupling tendency, comb

41、ined with the greatly reduced direc- tional stability, might be the source of undesirable stability character- istics at the high angles of attack. Lateral and Directional Control Aileron characteristics.- The effects of aileron deflection on the lateral aerodynamic characteristics at M = 1.61 for t

42、he basic config- uration are presented in figure lg. The aileron remains effective in producing roll throughout the angle-of-attack and angle-of-sideslip Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-y NACA RM 5605 .- 9 ranges investigated. The res

43、ults at a = O“ indicate that deflection of the left aileron provides larger increments of rolling moment and smaller increments of yawing moment at positive sideslip angles than at negative sideslip angles. This probably occurs because the flow over the left wing tends to become more subsonic at pos

44、itive sideslip angles and less subsonic at negative sideslip angles. These increments of rolling and yawing moments may also be associated with interference effects at the tail; however, no aileron deflection tests were made with the tails removed. Although the linearity of rolling moment with ailer

45、on deflection was not determined for deflections above about loo, it appears that sufficient rolling power would be available to neutralize the maximum rolling moments encountered throughout the a and p ranges investi- gated with the possible exception of some combinations of a and p above a = 12O w

46、here C2 P becomes large (fig. 15). The aileron effectiveness at j3 = 0 appears to increase slightly with increasing angle of attack (fig. 2l). Upward deflections of the left aileron caused a negative yawing- moment increment at low angles of attack, whereas downward deflections caused negative yawin

47、g-moment increments at high angles of attack. Although these increments were small, they may, under the conditions of initially low directional stability and for greater aileron deflections, assume greater importance. Deflection of the aileron does not appear to alter significantly the variation of

48、CL, Cx, and Cm with p for angles of attack of 0 and 8.3O (fig. 20). At a= 20.9, negative deflection of one aileron appears to result in a more rapid increase of Cm with p than for zero deflection. However, opposite deflection of the other aileron should reduce this effect. As expected, deflection of

49、 the left aileron produces slightly greater increments of lift and pitching moment at positive sideslip angles than at negative sideslip angles. The differ- ences in drag increments due to aileron deflection at positive and neg- ative sideslip angles (a = O“) were small. Rudder characteristics.- A rudder deflection of loo for the basic con