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WIND-TUNNEL INVESTIGATION AT SUBSONIC AND SUPERSONIC SPEEDS OF A MODEL OF A TAILLESS FIGHTER AIRPLANE EMPLOYING A LOW-ASPECT-RATIO SWEPT-BACK WING - STABILITY AND CONTROL By Willard G. Smit h Am es Ae ro nau tical Labor a tor y MoffeU Fi e ld, Calif.
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CLASSIFIED DOCUMENT This material contains informat1on affecting the Nat10nal Defense of the United States within the meaning of the espionage laws, Title 18, U.S.C., Sees. 793 and 794, the transmission or revelation of which 1n any manner to an unauthorized person 1s prohibited by law.
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
WASHINGTON Ja nu a r y 12, 1953 .........................
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NATIONAL ADVISORY COMMITTEE F O R AERONAUTICS RESEARCH MEMORANDUM WIND-TUNNEL INVESTIGATION AT SUBSONIC AND SUPERSONIC SPEEDS O F A MODEL OF A TAILLESS FIGEEEB AIRPLANE EMPLOYING
A LOW-ASPECT-RATIO SWEPT-BACK WING -
STABILITY AND CONTROL By Willard G. Smith This report presents the r e s u l t s of a wind-tunnel investigation of of a model of a f i g h t e r the s t a t i c s t a b i l i t y and control characteristics airplane employing a low-aspect-rati o swept-back wing with trailing- The edge elevons, a swept-back v e r t i c a l t a i l , but no horizontal tail.
investigation was conducted over a Mach number range of 0.60 t o 0.90 and 1.20 t o 1.70, a t constant Reynolds numbers of 2.0 million f o r the s t a b i l i t y t e s t s and 3.2 million f o r the control effectiveness t e s t s .
All r e s u l t s are presented i n tabular form and typical d a t a are pre- sented i n graphic form as w e l l .
The r e s u l t s indicate t h a t , for the t e s t conditions at which the investigation was conducted, the model, with elevons undeflected, was Sufficient control effective- longitudinally and directionally stable.
ness was provided by the trailing-edge elevons t o permit longitudinal balance of the model t o a l i f t coefficient of 0.44 at a Mach number of 0.90, and t o lift coefficients of 0.25 and 0.11 at Mach numbers of 1.20 With the rudder deflected 80 and the model a t and 1.70, respectively.
an angle of attack of -O.>O, the results indicate t h a t the model w i l l have s u f f i c i e n t directional control t o maintain s i d e s l i p angles of 3 . 6 O a t 0.90 Mach number and 2.3' a t 1.40 Mach number.
INTROlXTCTION The stab i lit y and c ontr ol e f f e c t ivene s s character i s t i c s of a i r c r a f t impor- flying a t high subsonic and supersonic speeds are of paramount A wind-tunnel tance i n the design of present-day fighter a i r c r a f t .
investigation has recently been conducted i n the Ames 6- by &foot supersonic wind tunnel t o study the s t a b i l i t y and control characteristics of a p a r t i c u l a r high-speed fighter model.
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$iCA RM A52530 The model had a lar-aspect-ratio swept-back wing and. a swept-back vertical t a i l . Two wing plan forms (the basic wing with rounded t i p s and a modified wing with triangular t i p s ) were tested i n the s t a t i c longitudinal s t a b i l i t y investigation. The model had no horizontal tail, The longitudinal control bein& obtained with trailing-edge elevons.
control effectiveness for full-span constant-chord elevons on the basic- wing model was investigated through a Mach number range of 0.60 t o 1.70.
A limited study w a s a l s o made of the effectiveness of elevons extending over approximately the outboard half of the wing panels. Rudder effec- tiveness was determined for the basic model a t 0.90 and 1.40 Mach numbers.
NOTATION Force coefficients are referred t o the wind axes. Moment coef- t h e origin on the fiicients a r e referred t o the s t a b i l i t y axes, with fuselage longitudinal a x i s a t the l a t e r a l projection of the quarter- chord point of the mean aerodynamic chord. In those t e s t a where yawing-moment coefficients were not measured, rolling-moment coef- axis.
f i c i e n t s are referred t o the fuselage longitudinal wing span, f e e t l o c a l wing chord measured p a r a l l e l t o wing plane of symmetry, f e e t f e e t free-6tream dynamic pressure, pounds per square foot drag coefficient l i f t coefficient coefficient cros s-wind-f'txce h i n g e 4 m e n t coefficient (hing; m n e n t e. e.. e.. .e. e... e.. e... e..
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NACA RM A52J36 3
rolling+noment coefficient ( r o l l i n tS oment
cz
p i tching-moment coefficient tc,ngFmoment Cm ( p i
yawing-moment coefficient ( yaw intSoment
Cn r a t e of change of yawing-moment coefficient with ~ g l e of CnB sideslip, per degree r a t e of change of rolling-moment coefficient with angle of c z P s i d e s l i p , per degree r a t e of change of l i f t coefficient with elevon deflection, C L6e measured a t zero elevon deflection, per degree C r a t e of change of rolling-moment coefficient with elevon ‘8, deflection, measured a t zero elevon deflection, per degree r a t e of change of pitching-moment coefficient with elevon Cm Fe deflection, measured at zero elevon deflection, per degree r a t e of change of cross-wind-force coefficient with rudder cch deflection, measured a t zero rudder deflection, per degree C r a t e of change of yawing-moment coefficient with rudder deflec- n8r t i o n , measured a t zero rudder deflection, per degree
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slope of the l i f t curve measured a t zero l i f t , per degree da
- ‘ % I
slope of the pitching-moment curve measured at zero l i f t dCL L - l i f t - d r a g r a t i o D m a x i m l i f t - d r a g r a t i o
( 9 -
M free-stream Mach number f i r s t moment of area of control surface aft of hinge l i n e , Ma f e e t cubed e.. e... ..e e... e.. e.. .e. e. e e. . e e e .
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4 .e. e... ..e * e * : * e * * $7A8A*RMeA52J30 R Reynolds number based on wing mean aerodynamic chord S t o t a l projected wing area, including area formed by extending leading and t r a i l i n g edges t o model plane of symmetry, square f e e t Y spanwise distance from plane of symmetry, f e e t
a angle of attack of fuselage longitudinal axis , degrees
P angle of s i d e s l i p of fuselage longitudinal axis, degrees 6 angle of deflection of control surface (angle between wing chord or vertical-tail chord and control chord), measured i n a plane perpendicular t o the controldurface hinge l i n e , degrees Sub s c r i p t s e combined inboard and outboard elevons e i inboard elevon e0 outboard elevon r rudder I a t o t a l d i f f e r e n t i a l elevon deflection, degrees APPARATUS I Wind Tunnel and Equipment This investigation w a s conducted i n the Ames 6- by &foot super- sonic wind tunnel. This wind timnel i s a closed-throat, variable- pressure wind tunnel i n which the stagnation pressure and the Mach num- ber can be continuously varied. The stagnation pressure can be varied from 2 t o 17 pounds per square inch absolute and the Mach number can be Further information varied from 0.60 t o 0.90 and from 1.15 t o 2.00.
this wind tunnel i s presented i n reference 1.
regarding The model w a s mounted on a s t i n g having a diameter which was 64 percent o f the diameter of the base of the model. The s t i n g support system allowed the model angle of attack t o be varied continuously from - 1 2 . 3 ' t o 22.5O.
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The aerodynamic forces and moments were measured by a four- component e l e c t r i c a l strain-gage balance mounted i n the body of the model. The balance i s similar t o t h a t used i n reference 2. The forces and moments were registered by recording-type galvanometers the balance.
calibrated by applying known loads t o Model A moriei of a higIi-speeci fighter airpiane (fig. i j having a low- aspect-ratio, swept-back wing, sweptaack v e r t i c a l tail, and no hori- zontal tsil w a s used i n this investigation. P r ~ T r l s i m s v e r e mde Pgr a l t e r i n g the plan form of the basic wing of the model by the addition These extended t i p s had a constant section of triangular wing t i p s .
thickness of 4.5 percent. A three-view drawing of the basic-wing model and the model with the modified wing i s shown i n figure 2.
The basic wing had a modified trapezoidal plan form with a 52.5' The modification leading-edge sweep angle and a taper r a t i o of 0.332.
consisted of rounding the wing t i p s t o f a i r i n t o the leading and trail- i n g edges (see f i g . 3). The wing was composed of symmetrical sections having a thickness of 7.0 percent of the chord (streamwise) at the wing root and tapering t o 4.5 percent of the chord ( s t r e m i s e ) a t the theo- r e t i c a l t i p . (See table I for wing-section coordinates.) These sec- tions were modified somewhat t o f a i r into the trailing-edge elevons which were flat sided.
The movable control surfaces on the m d e l consisted of c o n s t a n b chord trailing-edge elevons, each divided i n t o two spanwise segments, and a constant-percenhhord rudder (figs. 3 and 4). The control sur- faces on one wing panel and the rudder were restrained by beams f i t t e d s t r d n gages for measuring the control hinge moments.
with e l e c t r i c a l w a s f i t t e d with i n l e t s housed i n wing-body fairings with The mcdel and exhaust a t the rear i n t e r n a l ducts allowing the a i r t o flow through of the fuselage. In t h i s investigation, the mass flow of a i r through the ducts was not adjustable; however, the ducts were constructed s o that at supersonic speed the e x i t was choked, limiting the i n l e t Mach number t o 0.4.
I n order t o accommodate the annular duct e x i t and the mounting s t i n g , the boattailing on the model was somewhat l e s s than would be expected on a f'ullecale airplane.
A conventional canopy was used on the model with a dorsal f i n extending from the canopy t o the leading edge of the v e r t i c a l tail.
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e.. e... e.. e e.. 0 * e . 8 T k A RM A52J30 Provisions were m a d e f o r t e s t i n g the model without the v e r t i c a l t a i l but with the dorsal f i n f a i r e d i n t o the body. Table I1 presents the coordinates f o r the v e r t i c a l - t a i l sections.
TESTS AND PROCEDURE The aerodynamic c h a r a c t e r i s t i c s of both the basic-wing and modified- wing models were determined with control surfaces undeflected. L i f t , drag, pitchingdoment, and rollingdoment d a t a were obtained through an angle-of-attack range of approximately -3' t o +Eo at Mach numbers of 0.60, 0.80, 0.90, 1.20, 1.35, and 1.70. Tests of both models were con- ducted a t a constant Reynolds number of 2.0 million based on the m e a n aerodynamic chord of the basic wing (1.8 million based on the mean aero- I n the longitudinal s t a b i l i t y dynamic chord of the modified wing).'
phase of the investigation, the m d e l w a s mounted w i t h the wings verti- c a l i n the wind tunnel t o u t i l i z e the most favorable stream conditions (reference 1 ) .
The longitudinal control effectiveness of the elevons was investi- gated f o r the basic-wing configuration only. Tests of the model were conducted with the elevons on the r i g h t wing panel deflected. Incre- ments of l i f t , drag, and pitching moment due t o control deflection on the one wing panel were doubled and added t o the corresponding values I n this manner pitching- f o r the model with undeflected controls.
moment and rolling-moment data were obtained simultaneously, thus reduc- i n g the number of tests required. The v a l i d i t y of this procedure was checked by t e s t i n g the model through the speed range of the investiga- t i o n with the elevons on both wing panels deflected. Results of these two methods were i n excellent agreement. With the combined inboard and outboard elevons deflected through a range of 30 t o -200, l i f t , drag,
pitching-moment , rolling-moment, and hingeaoment data were obtained f o r
an angle-of-attack range of approximately -30 t o 120 at Mach numbers of 0.60, 0.80, 0.90, 1.20, 1.35, and 1.70 and a constant Reynolds num- ber of 3.2 million. Similar data were obtained at Mach numbers of 0.90 and 1.20 with the outboard control surface alone deflected through a range of OO t o 1 5 O .
The r e s u l t s of preliminary t e s t s of the basic-wing model a t Reynolds numbers of 1.0 t o 4.0 million a t supersonic speeds and 2.0 and 3.2 million a t subsonic speeds indicate that, within this range, Reynolds number variation had no s i g n i f i c a n t e f f e c t on the aerodynamic charac- t e r i s t i c s of the model with controls undeflected. The e f f e c t s of however, Reynolds number variation on elevon and rudder effectiveness, were not investigated.
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NACA RM A52530 7 The l a t e r a l s t a b i l i t y characteristics and rudder effectiveness of the basic-wing model were investigated with the elevons undeflected.
w a s mounted with the wings horizontal i n the tunnel, and the The model angle of s i d e s l i p was varied a t preset angles of attack. With the rud- der deflected through a range of 0 ' t o 8O, cross-wind-force, yawing- moment, rolling-moment, and rudder hinge-moment d a t a were obtained through an angle--of-sideslip range of 5 O t o -5O a t - 0 . 5 O , ?.lo, and 10.5O angles * of attack. Corresponding data were obtained under similar test condi- tions f o r the model with the v e r t i c a l t a i l removed. The l a t e r a l stabil- . - . . a a .. -E@- a&-. ...... -L- ..
< c , , - . . . a LLtJ C L L l U I U U U G I G A A G b L t I V G L I C : 3 3 YUU3G of t h Investig&tifin v s cc&.Uct& at Mach numbers of 0.90 and 1.40 and a t a constant Reynolds number of 3.2 million.
A tabulation of the t e s t conditions is presented i n table 111.
Reduction of Data The t e s t data have been reduced t o the standard NACA coefficient form based on the t o t a l projected wing area of the appropriate model configuration, including the area i n the region formed by extending the leading and t r a i l i n g edges t o the plane of symmetry. Factors which could a f f e c t the accuracy of these results and the corrections applied are discussed i n the following paragraphs.
Angles of attack and sideslip.- The determination of the actual angles of attack or s i d e s l i p of the m d e l under load required t h a t several corrections (determined from s t a t i c calibrations) be applied t o the nominal angle. Corrections of from 5 t o 10 percent of the nominal angle were applied f o r the angular deflection of the s t i n g and balance under aerodynamic load and for the angular movement due t o s t r u c t u r a l clearances i n the model support and balance.
Control-surface deflections.- A correction w a s applied t o the nominal c o n t r o l e u r f a c e deflection angle for the deflection under load as determined from the s t a t i c calibrations. The maximum correction amounted t o about 3 percent of the nominal deflection angle. The r e s u l t s presented herein are for the corrected control deflection angles except i n the figure showing variation of l a t e r a l s t a b i l i t y characteris- t i c s with s i d e s l i p angle a t various nominal rudder deflection angles.
Tunnel-wall interference.- Corrections t o the data f o r the e f f e c t s of the tunnel w a l l s at subsonic speeds were made by the method of refer- ence 3. The reflected bow wave did not intersect the model and s o no corrections were made at supersonic Mach numbers. T h e s e corrections, which were added t o the data, were as follows: ......................... . 0 . . 0 . .
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& NACA RM A52J30 ED = 0.0066 C L ~ A t subsonic speeds the e f f e c t s of constriction of the flow due t o the presence of the model were taken i n t o account by the method of reference 4. This correction was calculated for conditions a t zero A t angle of attack and w a s applied through the angle--of-attack range.
a Mach number of 0.90, this correction amounted t o a 1-percent increase i n Mach number and dynamic pressure over that determined from a calibra- t i o n of the wind tunnel without a n o d e 1 i n place.
Support interference.- The e f f e c t s of support interference were believed t o consist primarily of a change of pressure at the base of the model. A base-pressure correction was applied t o adjust the pres- sure a t the base of the m d e l t o free-stream s t a t i c pressure. The base area used i n t h i s correction was the e n t i r e base area l e s s the duct exit area. Drag values are, therefore, forebody drag coefficients. It was assumed, on the basis of information contained i n reference 5 , that the e f f e c t of sting-body interference on the forebody drag was negligible.
at subsonic and Stre\m variations.- Tests of the model were made supersonic speeds, i n upright and inverted attitudes. Results of these tests showed no measurable e f f e c t s of stream angle or stream curvature i n the horizontal plane of the wind tunnel.
Stream surveys conducted i n the Ames 6- by &foot supersonic wind tunnel (reference 1) show some curvature i n the v e r t i c a l plane of the wind tunnel, but the r e s u l t s of a previous investigation (reference 6 ) indicate t h a t this curvature had l i t t l e e f f e c t on the longitudinal s t a b i l i t y characteristics of the model when pitched i n the horizontal plane. For the lateral s t a b i l i t y t e s t s , the model was mounted with i t s wings horizontal s o that it yawed i n the plane of l e a s t stream curvature. No attempt was made t o determine the effects of the stream-angle variation i n the v e r t i c a l plane of the wind tunnel on the l a t e r a l directional data. The data obtained showed a s m a l l effect of stream angle on the r o l l i n g moment due t o s i d e s l i p and no effect on the yawing moment due t o sideslip.
Internal duct drag.- The model was equipped with twin ducts through which air could flow. However, provisions were not made t o vary the mass flow, s o a study of the duct drag characteristics was not feasible i n t h i s investigation. The drag data presented herein are f o r the com- plete model; t h a t i s , the drag due t o flow through the ducts has not been subtracted from the f i n a l coefficients.
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Precision of Data The accuracy of the t e s t r e s u l t s , excluding stream e f f e c t s , i s shown by the repeatability of the d a t a in those cases where t e s t condi- tions were duplicated i n several t e s t s . A n interim of three months elapsed between t e s t s during which the model ard balance were disas- sembled. The effects of changes i n clearance or alinement i n the model and balance determine t o a large extent the precision of these data.
Examination of the r e s u l t s showed the data t o be repeatable within the accuracy shown i n the following table: Accuracy Quantity CL = 0 CL = 0.4 f0. 001 f 0.002 &. 003 f .005
f . 001 f ,001
f ,0007 f .0017
f . 001 f . 001
k.003 5.005 f .008 5.013 f .03 f .03
f . 0 3 x io6 5.03 x 106
f .10 f .15 k .25 f 035 RESULTS AND DISCUSSION All the r e s u l t s of the investigation are contained i n t a b l e I V .
Brief discussions are presented of the longitudinal s t a b i l i t y charac- t e r i s t i c $ , the longitudinal control effectiveness, and the l a t e r a l s t a b i l i t y characteristics and rudder effectiveness i n the following paragraphs. Typical data, pertinent t o the discussion, are presented i n the figures.
Longitudinal s t a b i l i t y characteristics .- L i f t coefficient as a
function o f angle of attack, and the variation of drag and pitching- moment coefficients with l i f t coefficient are presented i n figure 5 for the basic-wing and modified-wing configurations with elevons unde- flected a t Mach numbers of 0.90, 1.20, and 1.70. Both configurations were longitudinally stable up t o a l i f t coefficient of 0.5 throughout ......................... . 0 . . 0 . .
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10 NACA RM A52J30 the Mach number range of the investigation.
The variation of pitching- moment coefficient with l i f t coefficient f o r the basic-wing model ( f i g , ?), although linear at 1.70 Mach number, exhibited a s l i g h t non- l i n e a r i t y at 1.20 Mach number, and was markedly nonlinear at a Mach num- ber of 0.90. The s t a b i l i t y of the basic-wing model (dCm/dCL) increased from 0.04 at zero l i f t coefficient t o 0.16 at a l i f t coefficient of 0.30 at a Mach number of 0.90. With the addition of triangular wing t i p s (modified wing), the s t a b i l i t y remained nearly constant with increasing l i f t coefficient up t o a l i f t coefficient of 0.30 at a Mach number of 0.90. Thus this increase i n s t a b i l i t y with increasing l i f t coefficient f o r the basic-wing model appears t o be a plan form e f f e c t . T h i s obser- vation i s substantiated by comparison of the r e s u l t s of an investiga- tion of the pitching-moment characteristics of a plane triangular wing of aspect r a t i o 4 (reference 7) with the r e s u l t s of a l a t e r investiga- tion (as yet unpublished) of the same wing with the t i p s cut off.
A summary of the aerodynamic characteristics of the two configur- The differ- tions, as a function of Mach number, i s shown i n figure 6.
ence i n s t a t i c margin at zero l i f t shown by the two plan forms of this investigation (fig. 6 ) decreased w i t h increasing supersonic Mach numbers.
It i s evident from examination of figures 5 and 6 t h a t the basic-wing model exhibited a greater change of s t a b i l i t y with increasing lift coefficient at subsonic speeds and a greater change of s t a b i l i t y ( a t zero l i f t ) w i t h Mach number than did the modified-wing model.
Longitudinal control effectiveness .- The longitudinal control
effectiveness investigation was conducted f o r the basic-wing configura- tion with the control surfaces shown i n figure 3. As noted previously, the control surfaces on only one wing panel were deflected and the increments of l i f t , drag, and pitching moment due t o the control deflec- tion were doubled.
The relationships of lift coefficients t o angle of attack, control- surface deflection, and drag coefficient for the airplane balanced with the combined control surfaces and with the outboard elevons alone are shown i n figure 7. These data indicate that, f o r the elevon deflection range of this investigation,the combined elevons would be capable of balancing the airplane (center of gravity a t 0.25 7 ) t o a l i f t coeffi- cient of 0.44 at a Mach number of 0.90, and t o l i f t coefficients of 0.25 and 0.11 at Mach numbers of 1.20 and 1.70, respectively.
A limited study of the control characteristics with only the out- board elevons deflected shows t h a t these elevons w i l l balance the model t o l i f t coefficients of 0.31 and 0.14 a t Mach numbers of 0.90 and 1.20, respectively, but at the cost of considerably greater control deflec- tions and consequently higher drag than with the combined control surfaces.
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NACA R M ~ 5 2 ~ ~ 3 0 ..e .
Examination of figure 7 reveals a decrease in the r a t e of change of balance l i f t coefficient with control deflection at 0.90 Mach number for both the combined elevons and the outboard elevons beginning at a l i f t coefficient of about 0.10. This apparent decrease i n effective- ness coincides with the increase i n s t a b i l i t y w i t h increasing lift coef- f i c i e n t discussed previously, and s o appears t o be the r e s u l t of the inherent s t a b i l i t y characteristics of the wing. Similar gradual decreases i n control effectiveness a t 1 . 2 0 and 1.70 Mach numbers are also presumed t o be due t o the increases i n s t a b i l i t y with l i f t coeffi- cient. The variations with Mach number of elevon l i f t , pitchingaoment, m& ruiiing+noment effectiveness f o r the combined elevons deflected are presented i n figure 8. It should be noted t h a t the values of rolling- moment effectiveness shmn are those f o r the elevm deflected 011 m e wing only, while the l i f t and pitching+noment effectiveness values are f o r deflection of the elevon on both wings.
at 0.90 and 1.20 Mach The stick-free s t a b i l i t y of the airplane numbers i s i l l u s t r a t e d i n figure 9 f o r the combined elevons f r e e and f o r only the outboard elevons free. The stick-fixed s t a b i l i t y curves, f o r the model with elevons fixed a t zero deflection, are a l s o shown f o r com- parison. It i s of i n t e r e s t t o note that f o r a Mach number of 0.90, the model exhibited a greater s t a b i l i t y s t i c k f r e e than s t i c k fixed, below a l i f t coefficient of 0.10. A n explanation f o r this greater s t a b i l i t y at low l i f t coefficients with the elevons f r e e can be found i n the tabu- lated hinge-moment data (table I V ) which show that the elevons f l o a t downward w i t h increasing angle of attack f o r angles of attack up t o 8 O .
The stick-free neutral points for the model with the combined elevons free are located at 32 and 41 percent of the mean aerodynamic chord at I Mach numbers of 0.90 and 1.20, respectively. With the inboard elevons fixed and outboard elevons free, the neutral points a r e a t 33 and 42 per- cent of the mean aerodynamic chord a t Mach numbers of 0.90 and 1.20, r e spec t i vely .
Lateral s t a b i l i t y characteristics and rudder effectiveness.- The variations of rolling-moment, yawingaooment, and cross-wind-force coef- f i c i e n t s with s i d e s l i p angle f o r the basic-wing model with zero elevon deflection a t 0.90 and 1.40 Mach number are shown i n figure 10 f o r angles of attack of 4 . 5 O and 5.1'. Also shown i n figure 10 are data f o r an angle of attack of l0.5O, obtained a t Mach numbers of 0.80 Since the data i n figure 10 revealed nonlinearities i n the and 1.40.
variations of yawing-moment and rolling-moment coefficients with side- s l i p angle, the variations of l a t e r a l s t a b i l i t y characteristics with angle of attack (fig. 11) are presented for both zero s i d e s l i p and a s i d e s l i p angle of 2O. Examination of figures 10 and 1 1 indicates that the model was directionally stable through the angle-of-attack and of the investigation and exhibited a positive angle-of-sideslip ranges dihedral e f f e c t a t the positive angles of attack.
......................... 0 . . 0 . .
0 . 0 . 0 . . 0 . . ........
. . . . . . . . . . . . . . . . . ......
0 . 0 . . . . .
. ....
0 . 0 . .
.......... .......
e 0..
YACA RM A52J30 12 The effectiveness of t h e rudder i n directionally controlling the model w a s investigated for the same range of t e s t conditions as were the l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s of the model w i t h controls unde- flected. Cross-wind-force, yawing-moment, rolling-moment, and rudder- hinge-moment data were obtained a t rudder deflections of 0 ' t o 8' and with the v e r t i c a l t a i l removed. Results of these tests, w i t h the excep- t i o n of rudder-Mnge-moment data, are shown i n figure 10 only f o r the model with 0 ' and 8 ' of rudder deflection since the variations of lat- eral s t a b i l i t y c h a r a c t e r i s t i c s with rudder deflection angle were found t o be l i n e a r f o r t h e range of rudder deflections tested. The model was capable of maintaining s i d e s l i p angles of 3 . 6 O and 2 . 3 O a t 0.90 and 1.40 Mach numbers, respectively, with the rudder deflected 8 ' at an angle of attack of -0.5'.
The variation of rudder effectiveness with angle of attack i s shown i n figure 12.
The variation of elevon-rolling-moment effectiveness w i t h s i d e s l i p angle was not investigated. However, a comparison of the m a x i m u m recorded r o l l i n g moment due t o combined angles of a t t a c k and s i d e s l i p with t h e elevon-rolling-moment effectiveness obtained at zero s i d e s l i p provides some indication of t h e a b i l i t y of the elevons t o balance the model i n r o l l at angles of s i d e s l i p . It w i l l be noted, from the data presented i n figure 10, t h a t the m a x i m u m r o l l i n g moments obtained f o r the model w i t h control surfaces undeflected occurred at an angle of s i d e s l i p of 5 O and a nominal angle of a t t a c k of 5 ' f o r both 0.90 and 1.40 Mach numbers. By comparison of these values of rollingaoment coeffi- cient with the data presented i n t a b l e I V , f o r the elevon-rolling-moment effectiveness at zero s i d e s l i p angle, it i s apparent that these rolling- moment coefficients a r e of approximately the same magnitude as those produced by a go t o t a l d i f f e r e n t i a l deflection of the combined elevons a t 5 ' angle of attack a t a Mach number of 0.90, and a 1 4 ' t o t a l differ- e n t i a l elevon deflection a t angle of a t t a c k at a Mach number of 1.40.
CONCLUSIONS A b r i e f analysis of the r e s u l t s of this investigation indicated that t h e following observations a r e worthy of note: 1. Both the basic-wing (rounded wing t i p s ) and the modifiedlring (triangular wing t i p s ) models w i t h elevons undeflected were longitudi- nally s t a b l e , through the Mach number range f o r which data were obtained, t o l i f t coefficients beyond those t o which the elevons were capable of balancing the basic-wing model a t t h e maximum elevon deflections c ons ider ed .
2. The modified-wing model (triangular wing t i p s ) exhibited a smaller change of s t a b i l i t y with increasing l i f t coefficient and with increasing Mach number than did the basic-wing m o d e l .
A t the m a x i m elevon deflection angles for which data were 3.
obtained, the conibined elevons provided sufficient longitudinal control t o balance the airplane t o a lift coefficient of 0.44 at a Mach number of 0.90, and t o lift coefficients of 0.25 and 0.11 at Mach numbers of 1.20 and 1.70, respectively.
With only the outboard elevons deflected,, the longitudinal control was snmevh~t less, but v w l d 2 s ---pp2 U U I I L L l e u L J --* * b V - balance the model t o lift coefficients of 0.31 and 0.14 at Mach numbers of 0.90 and 1.20, respectively.
4. The basic-wing model was l a t e r a l l y and directionally s t a b l e
through a nominal angle-of-attack range of Oo t o loo at Mach numbers
of 0.90 and 1.40.
The model w a s capable of maintaining s i d e s l i p angles of 3.6' 5.
a,nd 2.3O at Mach numbers of 0.90 and 1.40, respectively, with the rudder deflected 8O and at a -0.5' angle of attack.
A m e s Aeronautical Laboratory National Advisory Committee for Aeronautics Moffett Field, Calif.
R i c k , Charles W., and Olson, Robert N . : 1 . Flaw Studies i n the Asymmetric Adjustable Nozzle of the Ames 6- by &Foot Supeisonic Wind Tunnel. NACA RM A9E24, 1949.
2.
Olson, Robert N . , and Mead, Merrill H.: Aerodynamic Studyoof a
W i n g h e l a g e Combination Employing a Wing Swept Back 63 . -
Effectiveness of a . n Elevon as a Longitudinal Control and the Effects of Camber and Twist on the Maxim Lif't-Drag Ratio a t Supersonic Speeds.
NACA RM A50A31a, 1950.
3. Silverstein, Abe, and White, James A.: Wind-Tunnel Interference of the Wing with P a r t i c u l a r Reference t o Off-Center Positions and t o the Downwash a t the T a i l .
NACA Rep. .547, 1935.
4. Herriot, John G. : Blockage Corrections for Threedimensional-Flow Closed-Throat Wind Tunnels, with Consideration of the Effect of Compressibility. NACA Rep. 995, 1950. (Formerly NACA RM ~71328) ......................... .
0 . 0 . 0 . . 0 . .
.
................
0 . 0 . . . . ...
.......... . 0 0 . 0 -
Wind-Tunnel Invest i gat i on 5 . Phelps, E. R a y , and Lazzeroni, Frank A. : of-the Aerodynamic Characteristics of a l/l-cale b d e l of the NACA RM A51E28, 1951.
Northrop M X - m A Missile.
Aerodynamic Study of a 6. H a l l , Charles F., and Heitmeyer, John C . :
W i n g a s e l a g e Combination Employing a Wing Swept Back 630. -
Characteristics a t Supersonic Speeds of a Model with the Wing Twisted and Cambered f o r Uniform Load. NACA RM AgJ24, 1950.
7. Heitmeyer, John C.: L i f t , Drag, and Pitching Moment of Low-Aspect- Ratio Wings at Subsonic and Supersonic Speeds -Plane Triangular Wing of Aspect Ratio 4 with 3-Percent-Thick, Biconvex Section.
NACA RM A5lD30, 1951.
TABU I . - WING SECTION COORDINATES [Ccordinates given i n percent of local chord, measured p a r a l l e l of symmetry] t o plane Wing-root section Wing-tip seckion NACA 0007-63/30-9.5~ mod. NACA 0004.5-63/30-6.6° mod.
Statior! O r dinat e Station Or dinat e Station S t z t i s r ~. ~ ~- 0 4 2 . 5 0 0 42.5 0 2.250 3 452 .1 3 O E 43. . i .PO9 3.421 3 L / 45 .2 .2 .294 .458 4 7 . 5 3.378 47.5 .4 . 4 .413 .643 50 3.324 . 6 . 6 .504 .784 52.5 3.258 52.3 .8 .%
. go1 55 3.178 579
1.0 1 1.003 3.084 .645 57-5 57.5 1.2 1.2 .704 60
1 095 2 978 1
1 . 6 1.6 6 2 . 5 .807 62.5 1.255 2.857 2.234 .896 1.394 65 2 . 7 2 3 2.189 2.5 67.5 2 . 5 1.547 2 0 576 47.5 2.122 3 3 1.681 70 1.081 2 . 4 1 7 70 2.034 1.914 2.247 7 2 . 5 1.230 7 2 . 5 1.930 2.110 2 . 0 6 7 5 7 5 1.811 1.356 75 * 7 . 5 2 . 4 9 4 7.5 77;5 1.604 1.877 77.5 1.679 8 0 1.681 80 2 9 7 7 9 1.786 1.536 12.5 12.5 82.5 8 2 . 5 2.994 1 . 4 7 8 1.925 1 . 3 8 3 85 1 . 2 7 2 3 158 2.030 85 1.220 1 7 . 5 17.5 3.281 87.5 1.065 2.109 87.5 1.048 20.
20 -
.858 3 . 3 7 1 90 2.167 90 .869 22.5 22.5 3 433 9 2 . 5 .650 2.207 92.5 . 6 8 3 3 . 4 7 2 95 . 4 4 3 2.232 95 .49?
2 7 . 5 2 7 . 5 3.494 9 7 . 5 .236 2.246 97.5 2 9 2 100 30 * LOO 3.500 0 3 2 . 5 32.5 3 499 3 5 3 . 4 9 6 3 7 . 5 37.5 3.489 4 0 3 475 L.E. radius: 0 . 5 3 9 percent chord L.E. radius: 0 . 2 2 3 percent chord T.E. radius: 0.032 percent chord T . E . radius: 0 . 0 9 5 percent chord 1 6 TABLE 11.- VERTICAL TAIL SECTION COORDINATES [Coordinates given i n percent of l o c a l chord, measured p a r a l l e l t o t h e fuselage longitudinal axis] Tip section Root section NACA 0008-6 3130-9' NACA 0006-63/30-6°45' S t a t ion Ordinat e Station Ordinate 0 -1 0 -1 0,279 Q * 371 .2 .2 9 523 .4 .4 .551 .6 .6 .672 .895 .8 .8 1.029 1.0 1.0 1.146 .860 1.593 1.195 3 1.441 1.922 4 1.641 2.187 2.411 5 1.808 2.382 3 0 176 15 3.609 2.707 2.889 3.852 25 3 9 969 2 976 30 4.000 3.000 -r I 35 3.981 2 992 4-0 2.960 3.916 50 3.800 2.893 55 2.784 3.627 2.630 3 399 65 3.118 2.431 70 65 2.192 2 790 H.L. 75 2.426 1.921 H.L. 75 99 923 1.631 2.039 100 99.833 .167 .077 0 0 ~~~ L.E. radius: 0.396 percent L.E. radius: 0.704 percent chord chord; rudder has f l a t sides T.E. radius: 0.167 percent P.E. radius: 0.077 percent chord chord TABLE 111.- TEST C O N D I T I O N S [B, basic model; A, triangular wing t i p ; e i , inboard elevons; eo, outboard elevon; V, vertical tail; r, rudder] Reynolds No.
2onf iguration Test N o . Mach No.
(million ) o f model 1 0.6 2.0 B 2 .8 3 -9 4 1.2
1 1
1.35 6 J. J.
1.7 .6 i .8 B+A 8 .8 9 - 9 1 0 1.2 1 1 1-35 1.7
3 1 9 2 B 1
.6 14 .8 15 -9 16 1.2 1-35 1.7 . 6 19 1 20 .8 1.2 23 1.35
1.7 1
.6 .8 27 -9 1.2 29 1-35 30 1 . 7 31 .6 -8 -8 .8 -9 34 1.2 35 1.35 3 6 1.7 . 6 -3 -3 .8 39 -9 40 1.2 1-35
1.7 1
V .6
- -
.e. a*.. . E . .*.a e.. .a. . 0 . , a a. a . a a . a . a . a a . * a . . . . a a c a * . a .an a 4. a.
. . . . a a . * a . . a a . a 1P L a . *.D. .a. a. * * a L V TABLE 111.- ONCLUDED
- -
Reynolds No. Configuration Fach No.
Test 20.
j e (million ) of model
-
-
0.8 3 . 2 B 43 -9 46 1.2 1-35
4 4 8 1 I O
1 . 7 .6 . 8 51 09 1.2 53 1-33
-15 1
1 . 7 0 1
-9
; z 1.2
-15 57 -8 -9 1.2 36 -8 59 -9 -3 60 1.2 -3 *9 62 1 . 4 -9 64 1 . 4 65 -9 66 1 . 4 -9 65 1.4 -9 1.4 I V 71 09 1 . 4 B-V B 73 -9 74 1 . 4 75 -9 1 . 4 77 -9 1.4 79 B-V -9 80 1.4 B-V 8 1 .8 B 82 1 . 4 .8 84 1.4 . 8 1.4 .8 B -V 8 6 1.4 B-V V ir . . . .
I r- co
m 1
I l l m m n j w I I ~cucurlmwn~cu I I ~ r l m ~ c u c o r l m m : I I??????"? : 1"9"7c;"'-;??r: I I I L 1 ? " " u N ? 9 I I i o d c u f n f t m I I I r l c u f t j d m I I I r l c u m f f m m rl c u mcufwww mt-o n m o m m n w m o OW m w m r l cuw r l f ~ N W 0 0 ~ 0 0 0 0 0 0 ~ ~ O O r l r l d r l r l r l d O O O O O r l d r l N N c u m m
?S???SSSSSS 8 O o o o 0 0 0 0 0 0 0 0 0 0 0 0 0
. ? ? ? ? 8 8 ? ? ? ? ? ? ? ? ? ? ? ? ? ? ?
0 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 I I I I I I I I I I I f
W I
rl c u m e.. e... e..
. * e * e.. 0.4 * e * e. e .e e . * e e .
e .
e . e
e . . * e * . *
.
* e . e e.. e e.
Y e e e . . . . e e e .
e . . e * e e... e e . e . e .e. **.e e..
e e.. e. . * e e e NACA RM A52J30 .........................
. 0 . . 0 . .
........ . 0 . a . . 0 . 0 .
. I.. 0 . . . . . . . . . . . . . . . . .
G V .rl c V - w n W & Z f cu c u c u a" V w t - c u d f f m o d e-cn P -
......................... . e . . 0 . e
0 . 0 . . . e 0 . . ........
e e . e . 0 . 0. . ......
0 . 0 . . . . . .... . 0 . 0 .
.......... .....
0 0 . . * .
22 %ACA RM A52J30 I I .........................
Y e r n 0 . . . . . . .
........
...... ................
. . . . . . .... . . . . . . e r n
........ .... ..........
NACA RM A'j2J3d ......................... . . . . . . .
0 . 0 . 0 . . 0 . . ........
. . . . . . . . . . . . . . . . * ......
0 .
. . . .
0 . .... . . . . . .
.......... e 0 . . .........
24 NACA RM A52J30 m t Ln rl
8 Ln x 0
cn 0 In t - f W f 9 f
I
t - n ......................... . 0 . . a .
a . 0 . 0 . . 0 . . ........
. . . . . . . . . . . . . . . . . ......
0 . 0 . . . e . . e . . . .. 0 . .
.......... . 0.. 0 . e . 0 .
26 NACA RM A52530 m L n W c- W W W w .........................
. 0 . . 0 . .
........ . 0 . . a . 0 . 0 .
...... . . 0 . 0 . a e.. 0 . 0
. 0 . 0 . . .... . . e . 0 . 0 .
a 0 e.. 0 ..........
NACA RM ~52530 * * *.* ......................... . 0 . . 0 . .
0 . 0 . 0 . . 0 . . ........
. . . . . . . . . . . . . . . . . . . e . * .
0 . 0 . . . . 0 . 0 . .
.......... ...... ......
28 NACA R M A52530 r-r-mcomft-m f t c o w o m r l c o m m w - t m f w m ~WI~OO'M'U t-t-momr-t--4) c u c v c u m m w w w F ; A W S A A W W NCUNCUNCUCUN m n n n X k 2 X K ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ? ?
m m e m m eom o m m m e o m mea m m o m 0 0 .
o e a m m e o o m m a m e 0 m a m a a 0 a m om e m m m m m m 0 m m e o m a 0 m a m a e m m a m a o m m m e o m m e a NACA RM A52J3bcm " 29 t- Q) ...
...
...
• • •• ••• •• •• ••• •• •• • • • • • •
.. • • •
• • •
· • • • • •
• • •• • • •• • • •
•• · •• •
• • • • • • • • • • • • • • •••• • • • • • ....
• • • ... ...
... •
•••• • ••• • •• • NACA RIv1 A52J30 Figure 1.- The model mounted in the Ames 6- by 6 -foot wind tunnel.
. . . . . . . .........................
........ . . . . . . . . . .
...... . . . . . . . . . . . . . . . . .
. . . . . . .... . . . . . . . .
....................... ..........
NACA RM A52530 Figure 1 . - The model mounted i n t h e Amev 6- by 6-foot wind tunnel.
.** * * a * * * a *.*a a * * * * a * a * a * a * * a a . a * * * * * a . * * * a * * a * a * * * *
* a . * * . . a
a * * * * * a a * * a . * a a * * * a * * * a .** a * * * * a * * a * a NACA RM A52J30
T
B lo
i
(d b P a. a . ..
NACA RM ~ 3 2 ~ 3 0 33 ......................... . 0 . . 0 . .
0 . 0 . 0 . . 0 . . ........
. . . . . . . . . . . . . . . . . ......
0 . 0 . 0 . . 0 . 0 . .
.......... ....... .......
34 NACA RM A52J30 I I ' '.
\
i
I I a: \ \ \ \ \ \ \ I ......................... . . . . . . .
0 . 0 . 0 . . . . . ........
. . . . . . . . . . . . . . . . . ......
0 . 0 . 0 . . . .... . . . . . .
.......... ...... .......
36 NACA RM A52J30 . 0 . . 0 . . .........................
........ 0 0 0 0 0 . 0 0 0 0
...... . ................ - - - _ _
. 0 0 . i 0 i... - . 0 . . 0 . 0 .
....... 0.0 . ..........
RACA RM A52J3d0 37 .08
8 .06
!ii Q . .04 Basic wing solid line Modified wing dashed fine 4 .6 .8 L O 1 2 I4 L 6 L 8 Mach number Mach number Figure 6. - Summary of aerodynamic chorocieristics of the basic- wing and modified-wing models as functions of Mach number.
Reynolds number, 2.0 milkon (nomhal).
e a .a. a a e a o a a o a a a a a a a a a e a * a c a . a a.
a a . a . a . a a . a a * a * * * a a 0 a. .a a * a * a a * . 0 a a a * * * . a a * * *..a a . a * a a a . a .
* * a . a * * * a * a * * a NACA RM A52J3da* * * * Moth number, M Figure 6. - Continued.
.20 ./8 ./6 . / 4 . /2
. /o
.08 .06 .04 .02 .4 . 6 .8 1 . 2 1 . 6 /.8 Mach number, M / e ) C, vs M Figure 6.- Concluded.
......................... . t . . 0 . . a .
. e . e ........ 0 . 0 .
. . . . . . . . . . . . ......
0 . 0 . 0 . .
0 .
.......... om. o m o m s NACA F N A52530
tl
I co4
CL I ~~ ~ . . . . . . . .........................
........ e o . . . . . .
. . . . . . . . . . . . . . . . .
. . . . . .... . . . . . . .
......
. ..me at. m e . *.e 8 ...........
NACA RM A52530 43 \, ......................... . 0 . 0 0 . .
.
0 . 0 . 0 . . e 0 ........
.
. . . . . . . . . . . . . . . . ......
. . .
0 . 0 . m . 0 .
............... , e . * . a .
NACA RM A52530 44 c .4 . 6 .8 /.O /.2 1 4 1 6 /.8 .4 . 6 .8 1.0 12 1 4 /.6 /.8 Mach number, M -.4 . 6 . 8 LO L 2 / . 4 /.6 18 Mach number, M figure 8.- Summary Of elevon effectiveness charac?erisfics ff f zero lift coefficient as functions of Moch number. Reynolds number, 3.2 million.
e.. e e.. . * e e... e.. e... .*e e.
e . e . e e . e .e m .
. e .
e e . * . * * e
e. .e e ... . .ea e e
e . e . . . .
e . e m . e * e e . e .
.e. m e . . .e.
.e. e NACA RM A52J30 * e - * e Combined elevons fixed
- - Combhed elevons free
-- - - Outboord elevons free
. 6
-I .4
c ?
I
- 8 2
.04 0 -04 for M = 0 . 9 Pitching -moment coe f ficien f, Cm ~
Hgure 9. - The variution of 'pitchhg-moment coefficient with
lift coefficient for the mode/ with controls free und controls fixed at zero de flec fion . Reyno/ds number, 3.2 million.
b 0 0 w
p 9
P ......................... 1 e. . e . .
.
0 . * . 0 . 0 . . ........
.
e .
earn 0 0 . 0 e m 0. ...... e .
e . 0 . . . * e . e . 0 . e
.......... ............ .._e
.......
NACA RM A52J30 ?
co 9 .
I M =Q90
4 -- M = 140
-2 0 2 4 6 8 IO /2
Angle of offack, a, deg
-. 003
Y O 0 2 7 0 0 1 -2 0 2 4 6 8 IO /2
vj
Angle of attack, a, deg
Figure / A - The variation of the luferol stabikw cbaracteriktics
with of attack for the basic- wing mode/ with rudder and elevons UndefleCted ff eynolds number, 32 mil/ion.
.00/6 .00/2 F .0008 .0004 -2 0 2 4 6 8 IO 1 2 Angle of attack, a, deg ..002 -.001 -2 0 2 4 6 8 IO 1 2 Angle of attack, a, deg -$P= 1 6 ) p = 20 Figure I L - Conc/uded , OO/P .0008
L.F nnnd
. V Y " T 0 2 4 6 8 /O 12
Angle of oituck, a, deq
Ang/e of ottock, a, deg
Figure 12.- Variation of the rudder effectiveness churocteristics with ungle of attack for the basic-wing model with elevons undeflected. Reynolds number , 3.2 milfion .
NACA-Langley - 1-12-53 - 325 . ... ...
...
... ....
•• • •• • •• •• • • • • • • • • • • • • • •• • • • • • •• • • •• • • • •• • •• • • • •• • • • • • • •• • • • • • •••• • •
• . .. • ••• •••
• • • . ..
• • • •• • •••• • ••• ••• •• •
••• • ••• SJE: c: 1....U::~ I. T. Y. •• I f'.J FOR M AT ION
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