Document
-4 I NATIONAL AEROI?AU!I'ICs AND SPACE ADMINISTFATION
t3
k TECHNICAL X-303 ~~ STABILITY AND CONTROL C!€RRAclrERISTICS AT SPEEDS OF AN AIRPLANE CONFIGURATION An investigation to determine the static stability and control char- acteristics associated with two types of variable-sweep wings mounted on a fuselage representative of current supersonic fighter airplanes has been made in the Langley 300-MPH 7- by 10-foot tunnel. One of the wings having an outboard pivot was an advanced design of a previously tested wing which indicated small static-stability changes with changes in wing sweep. The second wing involved had a pivot-point location inside the fuselage, which may be a more desirable location f r o m a structural standpoint.
These low-speed tests indicated that by careful design of the wing and location of the outboard pivot the longitudinal stability could be maintained at essentially the same level for wing sweepback angles of 25O and 7 5 O ; thereby the conclusions previously reached vith the simplified research model were substantiated. An analysis of the pres- ent wind-tunnel results for three supersonic designs having combinations of a wing and horizontal tail, a wing, horizontal tail and c m d sur- face, and a w i n g and canard surface has been made for the configuration The wing and canard-surface arraage- having an outboard pivot location.
This fact ment indicated reduced static margin with increasing sweep.
may have important implications regarding reductions in the transonic stability shift and reductions in trim drag at the design Mach nuuker realized from use of a canard surface.
IIWRODUCTION In recent years there has been a renewed interest in variable-wing- sweep aircraft generated both by the desire for multimission aircraft and L by the fact that the design supersonic Mach numbers now being considered are such that considerably greater penalties, associated with the super- r' sonic wing planform requirements, are now being encountered at subsonic In addition to the obvious implications at subsonic speeds speeds.
these penalties, through excess fuel consumption, can seriously limit the supersonic phase of a given mission. The National Aeronautics and Space Administration, therefore, has undertaken a research program for the purpose of developing variable-wing-sweep configurations which would fulfill the current requirements better than would those developed in the past. For example, a considerably higher sweep range is needed because of the high supersonic Mach numbers required. Also, a method of avoiding the wing translation, utilized in previous variable-sweep aircraft as a means of minimizing the stability changes associated with the wing rotation, would be desirable.
The development of a variable-wing-sweep configuration xhich appears to satisfy reasonably the current requirements is described in refer- ence 1, and detailed subsonic, transonic, and supersonic aerodynamic data for the configuration are presented in references 2 to 6. This con- figuration possessed essentially the same longitudinal stability charac- teristics at both PjO wing sweep and 750 wing sweep without wing transla- However, the wind-tunnel model used in this study was simplified tion.
in order to be more adaptable to configuration development; therefore, ?
it appeared desirable to test the variable-sweep wing developed on a model more representative of current fighter airplanes. A research pro- gram which will provide such information for a Mach number range from low subsonic to a Mach number of 2 . 0 has, therefore, been initiated.
In the present investigation, two sets of variable-sweep wings were studied. One, which is referred to as configuration I, is an advanced version of the outboard pivot design described in references 1 and 2.
Although the aerodynamic superiority of this type of variable-sweep wing over one having an inboard pivot has been fairly well established, the possible structural penalties must, of course, be considered in applying the principle to an aircraft. It therefore appeared desirable to pro- vide aerodynamic data for both types of wings on a given configuration in order to facilitate the weighing of aerodynamic and structural con- siderations. In view of this a second wing, referred to as configura- tion 11, having a more conventional planform and a pivot located within the fuselage was also tested.
The purpose of this paper is to present the results of the low- speed tests made in the Langley 300-MPH 7- by 10-foot tunnel.
I 3 I - c SYMBOLS * The forces and moments are referred t o the body axis system except the l i f t and drag which, of course, are referred t o the wind a x i s , (See It is i q n r t m t te rrc?te that all C r \ P f f i C i P I l t S m e baaed e" the fig. 1 . ) highest sweptback-wing geometry of the configuration i n question, and the moment reference point f o r both configurations is located 0.609 inch above the f'uselage center l i n e a t a body s t a t i o n 67-03 percent of the The e f f e c t s of changes i n center of gravity due t o weight body length.
associated with changes i n the sweep of the wing panels have been neg- lected. The coefficients and symbols are defined as follows: L i f t
lift coefficient, -
ss
d r a g coefficient, - Drag
qs Pitching moment pitching-moment coefficient, qsz Rolling maanent rolling-moment coefficient , q= yawing-moment coefficient, Y a w i n g moment Lateral force l a t e r a l - f orce coefficient , qs dynamic pressure, lb/sq f t angle of attack, deg angle of sideslip, deg effective-dihedral parameter,
directional-stability parameter, -, *n per deg
as
3 9
side-force parameter, -, per deg
as
............... . . e.. 0 .
..
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . . .
...... .........
I section lift coe f f i cient cZ t ' longitudinal-stability parameter
lift-curve slope, *, per deg unless otherwise noted
aa L yawing-moment effectiveness parameter due to roll control, C n6h acn
0 -, per deg -
2 a ' h ac1
roll-control effectiveness parameter, -, per deg
C 28h %h
CL
-
maximum lift-drag ratio, ( L / D ) , , CD
- J O
, in.
C mean aerodynamic chord, r b / 2 C d Y J o ' local chord, in.
C average chord, in.
Cav longitudinal distance from wing apex to O.25c, in.
xc/4 longitudinal distance from w i n g apex to 0.255, xc/4 ~ b / 2 spanwise distance measured from root chord, in.
Y
-
l a t e r a l distance from fuselage center l i n e t o Y E ,
sgb/2 CY ay
, in.
p / 2 c Q- J o S wing area including enclosed area, 2Lb’2 c a y , sq f t b wing span, in.
h taper r a t i o A aspe c t r a t i o wing leading-edge sweep angle, deg % E horizontal-tail deflection, positive with t r a i l i n g edge ‘ h down, deg canard-surface deflection, positive with t r a i l i n g edge down, deg v e r t i c a l - t a i l deflection, positive with t r a i l i n g edge left, deg w i n g nose-flap deflection, positive with t r a i l i n g edge down, deg horizontal-tail dihedral angle, negative with t i p chord
rt
down, deg angle, negative with t i p chord down, deg canard dihedral Configuration component part designations: W Kina, B bow C canard surface T horizontal t a i l ............... ' .......
c 0 . e * e e 0 . 0 ..a e . .
. . . . . . . . . . . . . . . . . .
. . . . .
...... .........
P d MODELS Y' Configuration I Geometric characteristics of configuration I are presented in fig- ure 2, and photographs of this configuration in the tunnel showing the wing-sweep angles of 25O and 75O are presented as figure 3.
The fixed portion of this configuration had a leading-edge sweep of 60°, and the leading edge intercepted the fuselage center line at a station 17.9 inches from the nose. The outer panel with the wing leading edge swept back 25O had an NACA 6 5 ~ 0 0 6 airfoil section in the streamwise direction, and the inboard or fixed portion of the wing had an NACA 65A004.4 airfoil section in the streamwise direction.
The pivot-point location for the configuration corresponded to 51.08 percent of the semispan of the wing swept back 7 5 ' and was located 0.735 inch behind the moment reference point.
This pivot location was selected primarily from stability considerations. (See ref. 1.) The wing was displaced vertically 1.775 inches above the fuselage reference line.
A 15-percent-chord leading-edge flap was tested with the wing swept back 2 ! j 0 in an effort to increase the maximum lift-drag ratio of this configuration.
The horizontal tail employed in this investigation had an aspect ratio of 2.425, based on the exposed area and span with a leading-edge sweepback of 5 1 . 7 ' . These panels could be deflected 5, 0, -5, -10, and -15 to provide pitch or roll control and could be set at dihedral angles of 00 and -2OO.
The canard surfaces were of wedge airfoil section with a fixed dihedral position of -2OO.
Incidence angles of flOo could be obtained with these controls.
The fuselage used in this investigation was representative of cur- rent high-speed, twin-engine fighter airplane having a high-fineness- ratio forebody ahead of the engine inlets and with the engines housed in the fuselage. The entrance or capture area of the inlets was 6.020 s q in. and the exit area was 7.192 sq in.
Configuration I1 Geometric characteristics of configuration I1 are presented in figure 4, and photographs of this configuration in the tunnel with the wings swept back 4 3 . 0 3 ' and 70.50° are presented as figure 5. small The fixed glove at the wing-fuselage Juncture had a leading-edge sweep a . *Le e e e e . eo e e . * e e** e .
._ e * e .
e . . . e **I-; e* e e * e* e e . e
e e e e 7 0 e** 0. e* * e * *e e a i , of 60°, with the wing-apex intercept a t the same position as for con- figuration I. The wing w i t h the leading edge swept back 43.03' had an NACA 65~005 a i r f o i l section i n the streamwise direction. The wing pivot point was located within the fuselage and corresponded t o 28.05 percent of the wing semispan f o r the 70.50° sweepback condition and was located 3.85 inches ahead of the moment reference point. T h i s pivot iomj-ion was seiecied p r u = i l 3 s t y u c t - u - d co~~~~z~~i~~s.
The wing w a s displaced vertically 1 . 7 ' 7 5 inches above the fuselage reference l i n e and was tested a t leading-edge sweepback angles of 4 30.0O0, 43.03O, 60.50°, and 70.50° with aspect-ratio variation of L 4.494, 4.000, 2.582, and 1.754, a t zero incidence and dihedral.
2 Control surfaces f o r this configuration were the same as those for 4 configuration I, and the relative positions and pertinent geometry of these controls are presented i n figure 2.
TESTS AND CORRECTIONS The t e s t s were made i n the Langley 300-MPH 7- by 10-foot tunnel a t a tunnel aynaslic pressure of 80 l b / s q f t far configuration I and 63 lb/sq f t f o r configuration 11.
The Reynolds number for configuration I, based on the mean aero-
dynamic chord of the w i n g swept back 75O, w a s 2.382 x 10 , and the
Reynblds number for configuration 11, based on the wing swept back 70.50°,
was 2.219 x 10 . The models were sting-mounted (figs. 3 and 5) and a l l
forces and moments were measured w i t h a six-component strain-gage balance.
Jet-boundary corrections as determined from reference 7 have been applied t o pitching moment, drag, and angle of attack. Blockage cor- rections as determined by the methods of reference 8 have been applied t o the dynamic pressures and drag. The base pressure was measured and the drag was adjusted t o a condition of free-stream s t a t i c pressure a t The internal duct drag was also measured and subtracted f r o m the base.
the t o t a l drag. The angle of attack and angle of sideslip have been corrected for sting bending and balance deflections under load.
Transition w a s fixed on a l l surfaces including the fuselage. Num- ber 100 carborundum grains were used a t the 10-percent streamwise chord l i n e s f o r the w i n g s of configurations I and I1 swept back 25O and 30.00', respectively, and on the horizontal and vertical t a i l s e a t similar posi- tions. The fuselage transition s t r i p was placed a t a position 10 per- cent of the fuselage length aft of the nose.
............... .......
. .
. . . . . . . . . . . . . . .
..... .....
. . . . *: ......
....... ..a ..
J 8 PRESENTATION OF NSULTS d A f a i r l y extensive investigation of the s t a b i l i t y and control char- a c t e r i s t i c s of configurations I and I1 has been made a t low subsonic speeds, and i n order t o aid i n locating a particular s e t of data, tables showing the locations by figure number of the basic data for both con- figurations a r e given i n t h i s section.
Configuration I.- Figure Longitudinal aerodynamic characteristics: Wing-sweep e f f e c t s with canard surface off; rt = 0';
6 h = o o . . . . . . . . . . . . . . . . . . . . . . . . . 6
Effect of horizontal-tail incidence with canard surface
off; rt = oo -
A u = 2 5 O . . . . . . . . . . . . . . . . . . . . . . . .
A u = 7 5 ' . . . . . . . . . . . . . . . . . . . . . . . . 8
Effect of horizontal t a i l with canard surface on; 6, = 0 ' -
A~ = 2 5 O ; rt = oo . . . . . . . . . . . . . . . . . . . .
A m = 7 5 O ; rt = Oo or -20' . . . . . . . . . . . . . . . . 10
Effect of horizontal-tail incidence with canard surface off;
rt = -200 -
A u = 2 5 ' . . . . . . . . . . . . . . . . . . . . . . . . 1 1
A u = 7 5 ' . . . . . . . . . . . . . . . . . . . . . . . .
Effect of horizontal t a i l w i t h and without canard surfaces -
A m = 2 5 O . . . . . . . . . . . . . . . . . . . . . . . .
A u = 7 5 ' . . . . . . . . . . . . . . . . . . . . . . . .
Effect of wing leading-edge f l a p with canard surface off; A u = 2 5 ' . . . . . . . . . . . . . . . . . . . . . . . .
Effect of inboard wing leading-edge extension with canard surface off; A m = 7 5 O . . . . . . . . . . . . . . . . . .
Lateral aerodynamic characteristics:
Effect of various component parts -
& = 2 5 0 . . . . . . . . . . . . . . . . . . . . . . . .
Am-i7'jo . . . . . . . . . . . . . . . . . . . . . . . .
Wing sweep effects, canard surface off, horizontal t a i l o n . . . . . . . . . . . . . . . . . . . . . . . . .
Vertical-tai& control w i t h canard surface off; X) A u = 7 5 O . . . . . . . . . . . . . . . . . . . . . . . .
2 s I
Configuration 11.- Figure Longitudinal aerodynami c characteristics :
Wing sweepback effects w i t h canard surface off -
Horizontal t a i l on, = 0 ' . . . . . . . . . . . . . . .
Horizontal t a i l off 22 . . . . . . . . . . . . . . . . . . .
Effect of horizontal-tail deflection w i t h canard surface
off; rt = oo -
+ ~ = 3 0 . 0 0 ~ . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . . . .
= 43.03' 24 &-s=60.50° . . . . . . . . . . . . . . . . . . . . . .
b = 7 0 - 5 0 °
. . . . . . . . . . . . . . . . . . . . . . 26
Effect of horizontal-tail deflection with canard surface
off; rt = -2oO -
A m = 60.50° . . . . . . . . . . . . . . . . . . . . . .
= 70.w'
. . . . . . . . . . . . . . . . . . . . . . 28
Effect of horizontal-tail deflection w i t h canard surface
on; rt = OO -
A ~ ; E = ~ O . O O O
. . . . . . . . . . . . . . . . . . . . . . 29
A m = 43.03' . . . . . . . . . . . . . . . . . . . . . .
A m = 6O.5Oo
. . . . . . . . . . . . . . . . . . . . . . 31
Am=70.50°
. . . . . . . . . . . . . . . . . . . . . . 32
W i n g sweepback effects, canard surface on, horizontal t a i l on . . . . . . . . . . . . . . . . . . . . . . . . .
Effect of canard-surface control f o r w i n g sweepback of 70.50~; rt = oo . . . . . . . . . . . . . . . . . . .
Effect of various component parts -
Am=30.00° . . . . . . . . . . . . . . . . . . . . . .
%=43.03' . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . . . . . . .
A m = 6 0 . W o 37 4% b = 7 0 . ' j O o . . . . . . . . . . . . . . . . . . . . . .
Effect of adding rfing glove t o wing Kith canard surface off; = 43.03' . . . . . . . . . . . . . . . . . . . .
Lateral aerodynamic characteristics: * . . . . . . .
Effect of various component parts; = 70.50' 40 Horizontal-tail r o l l cohtrol with canard surface off; . . .
41 % =70.wo : . . . . . . . . . . . . . . . . . .
J I Summary plots are presented in figures 42 to 44 and for the most ?
part the discussion will be limited to these figures in order to expe- dite publication.
DISCUSSION Longitudinal Stability and Control Figure 42(a) presents the variation of pitching moment with lift L coefficient for configurations I and I1 with the horizontal tail off for the range of wing leading-edge sweep angles tested. Both configu- 0 rations indicate increasing instability with increasing lift coefficient. 2 In addition, the results for configuration I1 indicate a rather large 4 variation in stability level with sweep variations.
The results with the horizontal tail on are presented in fig- ure 42(b), and it will be noted that for both configurations the addi- tion of the tail tended to linearize the variation of pitching moment with lift coefficient. This favorable effect is associated with the location of the horizontal tail below the chord plane. However, a wide variation of stability with sweep angle is still encountered for con- figuration I1 with appreciable nonlinearities in pitching moment indi- cated for all sweepback angles except 70.50°. For configuration I, the change in stability level realized from increasing sweep is very small compared with the large change in stability encountered for configuration 11.
A comparison of the variation of longitudinal low-lift stability, untrimmed maximum lift-drag ratios, and lift-curve slope with changes in wing leading-edge sweep angle is presented in figure 43 for the two configurations tested. A change in static margin of approximately - 11.5 percent c is noted for the wing-body combination and a change
-
of approximately 1 4 percent c in static margin for the wing-body-tail combination of configuration I1 is encountered when sweeping the wing through a range of 4 0 . 5 0 ° . For the wing-body-tail combination of con-
-
flguretlon I, however, an increase in stability of only 2 percent c is noted for an increase of 50' in sweep. It must be kept in mind that at some intermediate sweep a somewhat larger rearward shift will be encountered (see ref. 2 ) . However, since this higher static margin will be encountered only during transition between the design sweep angles, it appears to pose no problem. An interesting point shown in figure 43 for configuration I is that for the tail-off configuration the aerodynamic-center location for the 7 5 O sweep condition is ahead of the aerodynamic center location for the wing leading-edge sweep of 2'3'.
This is a rather graphic illustration of the effectiveness of this type of variable-sweep wing in controlling the stability. The o m m a m a m a m a m o m m m a m o m om m a m m m m a m a increase i n s t a b i l i t y of the low-sweep condition over that of the high- sweep conditioo i s due t o the f a c t that the i n s t a b i l i t y associated with the fixed portion of the wing i s reduced because of the increase i n overall lift-curve slope which accompanies the reduction i n wing sweep.
For the case presented herein t h i s more than compensates f o r the for- ward movement of the outer panel aerodynamic center.
W i t h regard t o the variation of the lift-curve slope with sweep angle, the results (fig. 43) for both configurations indicate a decrease of approximately 50 percent a t low angles of attack as the sweep i s L increased f r o m the minimum t o the maximum angle. This decrease, of 1 course, would provide appreciable reduction i n the gust loads encoun- tered i n a low-level mission with the wing i n the high-sweep position.
~O For landing, take-off, and low-speed l o i t e r , advantage could be taken of the higher l i f t and corresponding lower drag due t o lift (as indi- cated by the l i f t - d r a g r a t i o s of f i g . 43) associated with the low-sweep high-aspect-ratio w i n g setting. Large decreases i n (L/D),, with increases i n sweep are noted f o r both configurations as would be expected. A comparison of values of untrimmed (L/D),, with and without the horizontal t a i l is a l s o presented i n figure 43 f o r both configurations. Addition of the horizontal t a i l reduces the values of (L/D),, f o r both configurations with t h e wing i n the least sweptback position A m = 25' f o r configuration I and ALE = 30.00' f o r con-
(
figuration 11). This e f f e c t i s seen t o diminish with increasing sweep, apparently due t o changes i n wing-induced-flow characteristics on the With regard t o t h e l i f t - d r a g r a t i o s , it m u s t be horizontal tails.
realized that considerably higher values would be expected at f l i g h t Reynolds numbers.
the basic-data p l o t s indicate suf- Regarding longitudinal control, f i c i e n t horizontal-tail effectiveness throughout the e n t i r e l i f t range f o r a l l configurations a t subsonic speeds.
Since configuration I appears more desirable frm a longitudinal- s t a b i l i t y standpoint (through a minimum of aerodynamic-center shift with w i n g sweep) while maintaining essentially the same variation i n an analysis was made of three l i f t - c u r v e slope and l i f t - d r a g r a t i o , air- arrangements of configuration I considered as possible supersonic The arrangements planes, and the r e s u l t s a r e presented i n figure 44.
presented include combinations of a wing and horizontal tail, a wing, and canard surface, and a w i n g and canard surface.
horizontal tail, For comparison purposes, the moment reference location f o r the three configurations has been adjusted so that f o r the 25O wing-sweep con- d i t i o n each configuration has the same s t a t i c margin of about
1 -
5- percent c .
4 d Sweeping the wing fr0m.25~ to 75O for the combination of the wing and horizontal tail results in an increase in stability of about 2 per- cent C; whereas the combination of wing, horizontal tail, and canard surface indicates no increase in stability, and the combination of the wing and canard surface provides a decrease in stability for increasing For the moment reference chosen, the combination of the wing sweep.
and canard surface shows essentially neutral stability for the 7 5 ' sweep An arrangement of this type may prove useful as viewed from condition.
the effects of minimizing transonic aerodynamic-center shift. If the
-
c
configuration were considered to have a static margin of 9 percent
2 L with a wing sweep of 2 5 ' at subsonic speeds, and were designed for supersonic speeds with the wing swept back 75O, then, the total increase in static margin realized from sweeping to 750 and increasing Mach nun-
-
ber to supersonic would only amount to approximately 1 0 percent c, assuming a typical transonic aerodynamic-center shift of approximately - 16 percent This fact, plus the reduction in trim d r ' a g realized c.
from comparison of canard arrangements with conventional tail-rearward configurations, appear to make this type arrangement promising from performance and stability viewpoints. It must be kept in mind, however, that some type of stability "fix" would probably be required to reduce the nonlinearities encountered at the high lift coefficients.
Figure 45 presents the span-load distribution for configuration I with the wing swept back 2 5 ' and 75O as calculated by the methods of reference 9 . Wing alone aerodynamic-center locations and lift-curve The cal- slope are also presented in the table included in figure 45.
culated loadings are based on the areas and spans of the respective wing in question, as indicated by the figure.
Lateral-Directional Stability and Control None of the configurations tested indicated any unusual lateral or directional stability characteristics. (See figs. 17 to 20, 40, Adequate directional stability was maintained to angles of and 41.)
attack of 2 0 ' or better which should cover the range of acceptable landing attitudes. Positive effective dihedral was obtained over same range of angles of attack.
approximately the
. Figure 41 presents the effectiveness of the roll-control tail
(differential deflecting of the horizontal tail) for configuration I 1 Wing-off results are with the wing in the 7 0 . 5 0 ' sweepback condition.
also presented and illustrate the effect of the wing in delaying the reduction in roll effectiveness to higher angles of attack. This is apparently due to wing downwash allowing the tail to operate in the linear portion of its lift curve.
.
t * !he yawing moments induced by the r o l l control t a i l are a l s o pre- Over a good portion of the angle-of-attack range, * sented i n figure 41.
a value considerably lower than the yaw-to-roll r a t i o is about 0.3, with the configuration of reference 2. This favorable t h a t obtained condition i s probably associated with the shorter wing-tail coupling At the higher angles the r a t i o is reduced of the present configuration.
even I”-~~=tiiei= due, in pari ai least, t o the horizontal-tail d i r f e r e n t i a l drag- CONCLUDING REMARKS These law-speed tests of two types of variable-sweep wings on a representative f i g h t e r airplane indicated t h a t by careful wing design the subsonic longitudinal s t a b i l i t y can be maintained at essentially
the same l e v e l f o r wing sweep angles of 2 5 ’ and no f o r configura-
t i o n I; thereby the conclusions previously reached with a simplified An analysis of the present wind- research model w e r e substantiated.
tunnel results f o r three supersonic designs having combinations of a wing and horizontal tail, a wing, horizontal t a i l and canard surface, and a wing and canard surface has been made f o r the configuration having an outboard pivot location. The combination of wing and canard surface indicated reduced s t a t i c margin with Increasing sweep. “his f a c t may have important implications regarding reductions i n the tran- sonic s t a b i l i t y s h i f t and reductions i n trim drag at the design Mach number which were realized from use of a canard surface.
Langley Research Center, National Aeronautics and Space Administration, Langley Field, V a . , May 13, 1960.
a m . a a REFERENCES 1. Alford, William J., Jr., and Henderson, William P.: A n Exploratory Investigation of the Low-Speed Aerodynamic Characteristics of Variable-Wing-Sweep Airplane . . Configurations. NASA TM X-142, 1939.
2. Alford, William J., Jr., Luoma, Arvo A., and Henderson, William P.: Wind-Tunnel Studies at Subsonic and Transonic Speeds of a Multiple- Mission Variable-Wing-Sweep Airplane Configuration. NASA TM x-206, 1959 * 3 . Spearman, M. Leroy, and Foster, Gerald V. : Stability and Control Characteristics at a Mach Number of 2.01 of a Variable-Wing-Sweep Configuration With Outboard Wing Panels Swept Back 75'. NASA i rm x-32, isg.
4 . Spearman, M. Leroy, and Foster, Gerald V.: Effects of Various Modi- fications on the Supersonic Stability Characteristics of a Variable- NASA TM x-260, Wing-Sweep Configuration at a Mach Number of 2.01.
1960.
5 . Foster, Gerald V.: Stability and Control Characteristics at Mach Numbers of 2.50, 3.00, and 3.71 of a Variable-Wing-Sweep Con- figuration With Outboard Wing Panels Swept Back 75O.
NASA m x-267, 1960.
6. Foster, Gerald V. : Effects of Spoiler-Slot-DeflectorControl on the Aerodynamic Characteristics at a Mach Number of 2.01 of a Variable-Wing-Sweep Configuration With the Outer Wing Panels Swept Back no. NASA TM X-273, 1960.
7. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, With Consideration of the Effect of Compressibility. NACA Rep. 993, 1930. (Supersedes NACA RM ~ 7 ~ 2 8 . ) 8. Sanders, J., and Pounder, J. R. : Wall Interference in Wind Tunnels of Closed Rectangular Section. Aero. Rep. AR-7, Nat. Res. Council of Canada (Ottawa), 1949.
9. Campbell, George S.: A Finite-Step Method of Calculations of Span Loadings of Unusual Plan Forms. NACA RM L50Ll3, 1951.
c 6
*
Figure 1.- System of axes used showing the positive direction of forces, moments, and angles.
ma earn a moa a m a am a a a o a a am o m a m m a a a a a a a a o a a a m a a a a a a a a m a rn 0 m a m a a 0 m a a a a a a ma a a a a a ma o a a ma , S c 1 * (a) W i n g swept back, A m = 2 5 ' .
L-59- 8261 (b) Wing swept back, ALE = 7 5 O .
L- 59- 8263 Figure 3.- Photographs of configuration I showing w i n g i n 750 and 2 5 O sweepback conditions.
1 1 I Wing in intermediule sweep conditions Wing swept 39x)w' und 7050.
Figure 4.- Two-view drawing of configuration I1 showing pertinent dimen- sions.
A l l dimensions i n inches, except as otherwise noted.
(a) W i n g swept back, A m = 43.03O.
L-59-8- I (b) Wing swept back, A m = 70.500.
L- 59- 8262 Figure 5.- Photographs of configuration 11, showing w i n g i n 70.50° and 43.03' sweepback conditions.
A .
k 0 .
m o V .
: Q k G k + o El \ 9, .
h C P I .
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the various parts of configuration I .
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4 fi 8, 0 W B T 0 0 O f f n WBCT 0 0 0 Figure 18.- Variation of s i d e s l i p derivatives with angle of attack f o r the various component parts of configuration I. A m = 750.
Angle o f attack,o,deg Figure 19.- Variation of s i d e s l i p derivatives w i t h angle of attack f o r configuration I w i t h canard surface off and horizontal t a i l on.
A m = 25' and 75'.
8, 4 8, 0 0 0 Off 0 0 0 0 Angle o f sides/ip,p, deg Figure 20.- Effect of v e r t i c a l t a i l on t h e lateral aerodynamic charac- t e r i s t i c s of configuration I with canard surface off; ALE = no; a 4.0’.
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.
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Figure 41.- Lateral control characteristics of configuration I1 w i t h canard surface o f f . A m = 70.50°; rt = Oo.
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12s
I - I
I -
c L i f t c o e f f i c i e n t ,c, Figure 44.- A comparison of the longitudinal stability characteristics for three arrangements of configuration I having wing leading-edge For comparison purposes, the moment sweepback angles of 25' and 75'.
reference location has been adjusted so that all three arrangements c All have the same static margin in the 2 5 O wing-sweep condition.
00 deflection.
control surfaces are at * e o o m D O : - : . m a ma: m a am m a m a o o ma o m - D' omm m m m o m m a mo *ma mo mo amo ma mom ma .
t
r; C R , .
Figure 45.- Span-load distributions for configuration I without hori- zontal tail and canard surface.
Am = 25' and 75'.
NASA - Langley Field, Va.
L-1024