RESEARCH MEMORANDUM NAT~NAL ADVISORY COMMITTEE I! 3 FOR AERONAUTICS CHARACT. WASHINGTON March 3, 1958

Transcription

1 RESEARCH MEMORANDUM STATIC LONGITUDINAL AND LATERAL STABILITY AND CONTROL CHARACT TICS OF A MODEL OF A SWEPT-WING FIGHTER- 3 A~RPLANE WITH A TOP INLET AT MACH B~MB+YPE 1 NUMBERS FROM 1.6 TO 2.35 NAT~NAL ADVISORY COMMITTEE I! 3 FOR AERONAUTICS WASHINGTON March 3, 1958

2 NACA RM A57JS2 NATIONAL STATIC LOIGITUDINAL AND LATERAL STABILITY AND CONTROL CHAR&CTERISTICS OF A MODEL OF A SWEPT-WING FIGFITER- BOMBER-TYPE AIRPLANE WITtr A TOP INLET AT MACE NUMBERS FROM 1.6 TO 2.35 By A. Vernon Gnos and Richard L. Kurkowski Static longitudinal and Lateral stabilfty characteristics of a swept -wing fighter-boniber airplane model with a top Inlet have been determined experimentally. In addition, vertical-tail effectiveness, spoiler effectiveness, effects of several store configurations, and effects of mass flow were investigated. Tests were made at Mach numbers of 1.6, 1.8, 2., 2.2, and 2.35 and Reynolds nmibers, based upon mean aerodynamic chord, of between 2.>C1O6 Ehnd 2.5x16. INTRODUCTION A swept-wing fighter-bomber-type airplane model with a top inlet was the subject of an investigation in the Ames 9- by 7-foot supersonic wind tunnel (ref. 1). The engfne air inlet of the test configuration was located on the top of the fmelage behind the cockpit canopy. This unusual inlet location has a number of advantages both aerodynamic and mechanical. Some of the more important advantages that accrue from this location include: freedom for carrying stores under the fuselage, short duct lines, and the possibility of goo& high angle-of -attack characteristics. Since inlet location can have important effects upon stability characteristics af an airplane, especially where the inlet is large, the subject investigation of the top-inlet modelwas conducted. Further, because the vertical tail and parts of the fuselage are immersed in the external flow field of the inlet, variation of internal flow conditions can affect the kteral characteristics of the airplane. Accordingly, the investigation included the effect on lateral characteristics of three mass-flow conditions: supercritical, engine matched (appro-te), and low subcritical where buzz was alwa s resent.

3 2 c - MACA RM A57K2 of store cpnfiguration-_u_ppq longitudin@ &a$eral and characteristics and the effects of sp6ller deflection upon lateral control parer were investigated "., -r NOTATION Force and moment coefficients are referred ta- the stability axes with the origin on the fuselage reference axis at the projection of the 35-percent point of the wing mean aerodynamic chord. The aystem of axes and the positive direction of forces, moments, and angles are shown in figure 1. drag coefficient, measured drag - base drag - internal drag ss lift coefficient Y * side-force coefficient, side force qs rolling-moment coefficient, rolling moment ssb pitching-moment coefficient, pitching. ss E yawing-moment coefficient, yawing moment qsb free-stream Mach nmiber basic wtng area, 3.46 sq ft wing span, ft wing chord, ft moment wing mean aeroa;ynamlc chord, M.A.C., 1.19 ft free-stream aynamic pressure, lb/sq ft mass-flow ratio, angle of attack, deg - compressor mas6 flow free-stream mass flow based on capture area

4 NACA RM A57K2 3 angle of sideslin, deg angular deflection of inboard deflector, deg angular deflection of intermediate inboard deflector, deg angular deflection of intermedlate outboard deflector, deg angular deflection of outboard deflector, deg angular deflection of inboard spoiler, deg angular deflection of intermediate inboard spoiler, deg angular deflection of intermediate outboard spoiler, deg angular deflection of outboard spoiler, deg angular deflection of vertical tail, deg rate of change of rolling-mment coefficient with sideslip angle, 3, per as rate of change of yam-moment coefficfent with sideslip angle, The model is illustrated in the photographs of figure 2. Sketches of configuration details are shown in figure 3. Geometric dimensions are listed in table I. - The test model was mounted on a hollow sting by a seven-component strain-gage balasce so that forces and moments were recorded simultaneously with the internal-flow measurements. me balance consisted of four normsl-force gages, two side-force gages, and a chord-force gage.

5 4 - NACA RM A57K2 Balance forces were read out and recorded by conventional wind-tunnel d equipment. The model had a vertical-wedge inlet, subsonic diffuser, wedge boundary-layer bleed system, fmelsge boundary-layer bleed system, and cooling es& ventihting intake scoops and exit. Internal flow was L reguhted with an iris valve'located at the exit of the hollow st-. Mass flow was measured with a calibrated flow meter which was built into the sting. Three store configurations were tested with the model.they included a semisubmerged tank without fins as shown in figure 3(a), a saddle tank, and a saddle tank plus small store with fins as shown in figure 3(b).... The model was provided with an all-movable vertical tail which was tested at deflection angles of g and -3 and an all-movable horizontal tail which was tested with no deflection. The left wing of the model was provided with an instrumented spoilerslot-deflector system. Each spoiler and deflector could be adjusted to fixed deflection Wles of between 5O and 7. Spoiler detalls are shown in figures 2( c), 2( a), 3( a), and 3( c). The spoilers and deflectors were replaced with blank plates for tests which required no spoiler deflecticm. " 'TEST PROCEDURE "?The investigation was conducted at Mach nwlibers of 1.6, 1.8, 2., 2.2, and 2.35 and Reynolds number6 based uponmean aerodynamic chord of between 2.~l-O~ and 2.5>ito6. Angle of attack. was varied from -6 to +2 and angle of sideslip was varied from -9 to +4O. Sideslip runs were made at nominal angles of attack of -5O, Oo, 7, and 14O. kss-flow ratio was set at the estimated matched value for the model attitude of 2O angle of attack &nd Oo sideslip. No further aaustrnent of the ma66 flow was made during pitch or sideslip rims, even though "buzz" (defined as unsteady flow in the inlet and subsonic diffuser at slibcritical massflow conditions) was sometimes encountered. BUZZ generally occurred above 4 sideslip at a Mach nwer of 1.8 and above 2 sideslip at a Mach nmiber of 2.2. Buzz was not encomtered during pitch rum. Measured angles of attack and sideslip were corrected for the tunnel stream angle and for sting and balance deflection unaer load. A buoyancy correction was applied to take into account tunnel static-pressure varlations. The data have been corrected by adjustjng the measured base pressure to free-stream static pressure. In addition- the internal drag, which was determined from the change - in momentum from free-stream conditions to measured conditions at the duct exit, -was subtracted from measured. drag. No corrections were made for inlet spillage drag or for internal drag of the cockpit cooling a& ventilating air flow. Deflections of the vertical. tail, spoilers, and deflectors under load were not known.

6 HACA RM ~57~2 5 k Precision of the data, which wa6 determined from scatter and repeatability for moderate angle-of -attack conditions, is indicated in the following table: CD m.5 The longitudinal characteristics of the basic configuration and the tail off configuration are presented in figure 4. In figure 5 longitudinal characteristics with two additional store configurations are shown. Figure 6 summarizes pitching moment at zero lift and aerodynamic center location as taken from figures 4 and 5. Angle-of-attack effects on lateral characteristics of the basic configuration are shown in figure 7 and summarized in figure 8. Angleof-attack effects on lateral characteristics of the vertical tail off configuration are presented in figure 9 and summarized in figure 1. Figure El presents the effect of vertical tail on lateral characteristics of the basic configuration. The effect of vertical tail on the lateral characteristics of the saddle tank and store configuration is presented in figure 12. The effect at Mach n-er 2.2 of internal flow conditions upon lateral characteristics of the basic configuration is shown in figure 13 and of the vertical tail off configuration is shown in figure 14.

7 6 - NACA RM A57K2 Rolling-moment effects of spoiler deflection are presented in figure 15. Figure 16 presents yawing-moment effects of spoiler deflec- tion. In figure 17 the effect of spoiler deflection on Longitudinal characteristics is shown. SUMMARY OF RESULTS General observations regarding the results are as follows: 1. Aerodynamic center did not shift significantly with Mach nuniber for any of the tail-on configurations. Store configuration had little effect upon the location of the aerodynamic center. 2. Directional stability was maintained to about W mgle of attack at a Mach nuniber of 2.2. The usual trend of a decrease in directional stability with an increase in either angle of attack or Mach nuniberwas evident. 3. Large changes in directional stability were encountered at low inlet mass-flow ratlos at a Mach nuiber of 2.2. The vertical-tail contribution to directional stability was not seriously affected by low inlet mass-flow conditions since sfmilar changes were obtained with the tail off. There were shifts in dtrectiona1 stability near angle of sideslip associated with unstable air flow in the twin-duct system at low inlet mass flaws. Ames Aeronautical Laboratory National Advisory Cormnlttee for Aeronautics Moffett Field, Calif., Nov. 2, Huntsberger, Ralph F., and Parsons, John F.: The Design of Large High-speed Wind Web. Papers presented at 5th meeting of Wind Tunnel and Model Testing Panel, AGARD, Fourth General Assenibly, S cheveningen, The Netherlands, Rep. AG 15/F6, May 3-7, 19%, p~

8 NACA RM A57K2 " 7 WEE I.. GEOMETRIC CEAFUC~RISTICS OF MODEL wing Total basic area. sq in Span. in Aspectratio Taper ratio....3 Dihedral angle. deg... Root chord. in Tip chord (equivalent). in Mean aerodynamic chord. in Wing station of mean geometric chord. in Sweepback of quarter chord. deg...: Incidence. deg... Thickness. percent... 5 Airfoil section... NACA 665 modified. Fuselage. Length. in Fertical tail Total area. sq in Span. in... ll.78 Aspect ratio Taper ratio Root chord. in u Tip chord (eqrdvdent) in Mean aerodynamic chord. in Sweepback of quarter chord. deg Ratio of vertical-tail area to wing area Area moment. cu in Thickness. percent Airfoil section... NACA 6W3.5 zorizontal tail Total area. sq in Span. in Aspectratio Taper ratio Dihedral. deg... Root chord. in ~ i chord p (equivalent). in Mean aerodynamic chord. in Sweepback of quarter chord. deg c Ratio of horizontal-tail area to wing area Thickness. percent Airfoil section... MCA 65AOO3.5."

9 8. NACA RM A57K2 TABLE I.. GEOMFX!RIC CHARACTERISTICS OF MODEL. Concluded T'Horizontal stabilizer Area (movable portion only). sq in span. total (movable portion only). in Area moment. cu in Trailing-edge angle. deg Semisubmerged tank Plan-form area. sq in Length. in Maxirmun diameter. In Leading-edge location (fuselage station). in Small center-line store and saddle tank Saddle tank length Saddle tank leading-edge location (fuselage station). in Store length. in Store maximum diameter. in Store leaang-edge location (fuse-e station).in

10 ..... I / Egure 1. - System of axes and positive direction of force^, moments, and angles,

11 , (a) Threequarter top view of basic configuration with spoilers deflected 7. A * 3 Figure 2. - Model photographs

12 I I I...

13 - A (c) Close-up of spoiler detail. Figure 2. - Continued.

14 - (a) Close-up of deflector details. Figure 2. Concluded.

15 14 NACA RM A57KX) J I in inches (a) Three views of basic configuration with spoiler-slot-deflector system. Figure 3.- Configuration details. L

16 ra I L-cA I I Section A-A

17 16 NACA RM ~57~~2 line Spoiler -slot deflector Uensions Chord 4 c mean, 3n.l Spoiler No Spoiler No Spoiler No Spoiler NO Slot Deflector NO Deflector No Deflector No Deflector NO &an chord - measured perpendicular to the hinge line from the hinge line to the apposite edge 2 PLen form - length A times the mea chord (c) Spoiler-slot-deflector system, dimensions, and direction of deflections. Figure 3.- Concluded.

18 cl.4.e o.&.os.e o

19 (b) M = 1.8 Figure 4. - Continued. I... "

20 L 1 1 Figure 4.- Continued.....

21 ~ I I.....

22 (e) M F'irmre 4.- Concluded.

23 , Q, deg.a -.a -.a !a c, (a) M = 1.6. Flgure 3.- Effects of store configuratians on longitudinal charecteri6t;ics I

24 * I 1

25 cl I.2 i I! I.o4.8.E.16.2 ID -8-4 O 4 8 ~ 216 a, aeg.4 O O8 cm (c) M = E O -.e Figure 5.- CondLUded

26 4T NACA RM ~57~2 - Basic configuration ----Saddle tmk saddle tank and store - - Tail off a.c., percent E a.c., - percent c o M Figure 6.- Variation of pitching moment at zero fift and aeroaynamic center with Mach nuniber for configurations tested.

27 cz -.1 B, aef3 (a) M = 1.6 Figure 7.- Angle-of-attack effects on lateral characteristics of the basic configuration.

28 NACA RM A57K c2 m.1 BJ (b) M = 1.8 Figure 7. - Continued... mmmmmmm

29 2 - MACA FN ~57~2.16.E Oh c Bt (c) M = 2. Figure 7. - Continued.

30 . NACA RM ~57~2 -.x) E.8, Ob % Figure 7.- Continued.

31 - NACA RM A57K.2 - " h.".1 cz -.1

32

33 "..." Bt deg (8) M = 1.6 Figure 9.- Angle-of-attack effects on lateral characteristics of the vertical tail off configuration.

34 3T NACA EM ~ 57~2 "a 33 b Figure 9.- Continued.

35 NACA RM A57K L.8.1 Ct J.1 -.Ol I! Y B, -(c) M = 2.2 Figure 9. Concluded.

36 NACA RM Ag 7Kx> I._ 35.. Figure 1. - Variation of rolling-moment yawing-moment derivatives with angle of attack for the vertical tail off configuration.

37 . - B, d=g (a) M = 1.6 Figure I".- Effect of vertical tail on lateral characteristics of basic configuration; a =.

38 cz -.Ol 1.2 Figure Continued.

39 Figure ".. ~ U. - Concluded

40 NACA RM A57Kx) Figure Effect of vertical tail on lateral characteristics of the saddle tank and store configuration; a = W.

41 .1 cz cn.1 m L B, deg r v (b) Vertical tail off. Figure Concluded.

42 .2 41.r6.E.8.4 c Ob m s k.4.1 cn Figure Effect of mass-flow ratio on lateral characteristics of basic configuration; M = 2.2.

43 42 I^u NACA RM ~57~2.16 w e matched.12 Supercritical.8.4 cy -.Ob I ma/mm.4.2 c '

44 .16.E Oh B, aeg (c) u = 14 Figure 13.- Concluded.

45 44 " NACA. R M A -16 " t.e Ol 3..o.a "" Figure 14.- Effect of mass-flow ratio on lateral characteristics of vertical tail off configuration; M = 2.2.

46 45.1 Ct c cp -.1

47 7 NACA RM A57K Ol oe - Figure 14. Concluded.

48 m.2 -a a> dee (a) Equd spoiler and deflector angles. Figure l5.-,rolling=ment effects of spoiler deflection for basic configuration.

49 ' m x) 24 a, - deg (b) Various deflector angles. Figure 15.- Continued.

50 7T MACA RM ~57~2.2 " 49 a.1 C2 U A loo oo loo " 4" Oo. 4" " 7" oo 7" oo Oo 4" " 4" d cz a, de@; (c) Inboard spoiler and deflector only. Figure 15. Concluded.

51 Ol -.2. c1 -.Ol -.2 Figure 16.- Yawing-moment effects of spoiler deflection for basic

52 .2 a ' ar w3 (b ) Various deflector angles. Figure 16.- Continued.

53 .2 1" " 1 oo. 1 4" oo 4" oo 7" oo 7" " A " 4" " 4" 1.1 I I I I I I I I I 1 I I I Figure Concluded. c LI

54 * '' I I'; -8-4 O 4 8 E cr, &!3 a 4 -.Ob E PO c p cm (a) M = 1.6 Figure 17.- Effect of spoiler deflection on longitudinal characteristics of the basic configuration.

55

56 1 I I 1 I Flgure 17.- Concluded

57 f t # I I