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RESEARCH MEMORANDUM
EFFECTS ON STATIC! STAB1L;ITY CHARACTERJSTICS O F
VAR;IOUS MODIFICATIONS TO A SWEPT-WING
NATIONAL ADVISORY COMMITTEE
F O R AERONAUTICS
WASHINGTON AprLi 3G, 1957 E NATIONAL ADVISORY C O M T F E X FOR AERONAUTICS RESEARCH MEMDRANDUM .II
s a
WIND-TUNNEL INVESTIGATION AT HIGH SUBSONIC SPEEDS OF THE
4B P
EFFECTS ON S " I C STABILITY CHARACTERISTICS OF VARIOUS MJDIFICATIONS TO A SWEPT-WING FIGHTER-TYPE AIRPLANE KIDEL By Kenneth W. Goodson An investigation w a s made at high subsonic speeds of a model of a twin-engine swept -wing f ighter-type airplane. The model w a s t e s t e d with several different tail configurations and with several wing and engine i n l e t modifications. The investigation w a s concerned primarily with at a s t a b i l i z e r incidence of Oo; however, a few longitudinal s t a b i l i t y The model w a s s t a b i l i z e r and lateral-derivative t e s t s a l s o were made.
tested i n the Langley high-speed 7- by 10-foot tunnel a t Mach numbers from 0.60 t o 0.92.
The r e s u l t s showed t h a t a horizontal t a i l raised t o t h e top of t h e v e r t i c a l t a i l and moved forward fromthe original position gave substan- t i a l improvements i n s t a b i l i t y and delayed the angle of attack and l i f t coefficient a t which pitch-up i n s t a b i l i t y occurred. U s e of a fixed hori- zontal surface, attached below t h e engine t a i l pipes, i n combination with e i t h e r the original or new horizontal t a i l (biplane arrangement) a l s o provided substantial improvements i n s t a b i l i t y ; however, the com- bination that included the new horizontal t a i l had b e t t e r characteris- t i c s . Scme additional but smaller improvements were obtained when wing leading-edge chord-extensions, modified wing trailing-edge f i l l e t s , and modified engine i n l e t s were used.
INTRODUCTION Many swept-wing high-speed airplanes experience abrupt changes i n longitudinal s t a b i l i t y (pitch-up) at moderate and high l i f t coefficients.
The present investigation w a s undertaken t o investigate various possi- b i l i t i e s of alleviating o r possibly eliminating the pitch-up problem on 1 ' ..........
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NACA RM L57Ajl A a twin-engine swept-wing fighter-type configuration. Results of rocket- model investigations of this configuration are published in reference 1.
t ' The present model was tested in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92 over an Sev- angle-of-attack range of 0 ' to 2 7 ' (for the lower Mach numbers.)
eral lateral derivative tests were made through the angle-of-attack Analysis of these data has been l i m - range at sideslip angles of + 4 O .
ited in order to expedite publication.
SYMBOLS The data are presented about the system of axes shown in figure 1.
The pitching-moment coefficients are referred to a center of gravity at the 28.57-percent mean aerodynamic chord of the theoretical wing.
Lift lift coefficient, - CL qs Drag drag coefficient, - CD qs Pitching moment Cm pitching-moment coefficient, qsc' Rolling moment rolling-moment coefficient, Cl qsb Yawing moment yawing-moment coefficient, Cn qsb Side force side-force coefficient, CY
9s
dynamic pressure, - pv2, lb/sq ft mass density of air, slugs/cu ft P
v free-stream ve lo c ity , ft / s e c
M Mach number S wing area (theoretical area, neglects inboard wing fillets, see fig. 2), sq ft C l o c a l chord p a r a l l e l t o plane of symmetry, ft
-
C
w i n g mean aerodynamic chord, 6 s,"" c+y, f t
ch . horizontal-tail mean aerodynamic chord, f t
-
CV v e r t i c a l - t a i l m e a n aerodynamic chord, ft b wing span, f t Y spanwise distance from plane of syrmoetry, f t a angle of attack of wing, deg P angle of sideslip, deg s t a b i l i z e r incidence, positive with t r a i l i n g edge down, deg it
r dihedral angle, positive with t i p s up, deg
surface leading-edge sweep, deg %e Subscripts: denotes p a r t i a l derivative of coefficient with respect t o P
- %
sideslip angle; f o r example
czp - qr-
The various components of t h e configurations presented herein a r e designated as follows:
w wing
F fuselage V O original v e r t i c a l t a i l new v e r t i c a l t a i l v1 HO original horizontal t a i l H1 new horizontal t a i l e 0 . 0.. . 0.. . 0 . 0 . . . . 0.. 0 .
0 . 0 . 0 . . 0 . . e . . 0 . .
0 . 0 . . 0 . . . . 0 . . . . . . e
\ 0 . 0 . 0 . .
0..
. ...
0 . ... . . . &c&f$dW: 0 . ... 0 . NACA RM L57Ajl '
fixed horizontal surface attached below engine tail pipes
H2
original horizontal-tail plan form in low position J "3 low-aspect-ratio highly swept horizontal surface H4 MODEL AND APPARATUS A two-view drawing of the basic complete model (WFVoG) is shown in figure 2. The model wing had an aspect ratio (neglecting fillets) of 4.270, taper ratio of 0.284, 1.67-percent positive leading-edge camber, and 4 1 . 1 ' leading-edge sweep. This nominal plan form had been modified to accommodate the engines by extending the inboard portion of the trailing edge. The original (basic) vertical- and horizontal-tail configuration also is shown in figure 2. The indicated difference between model and airplane (fig. 2) is a result of the model being adapted to the sting support.
A new vertical- and horizontal-tail combination ( V l H 1 ) was designed so that the mean aerodynamic chord of the horizontal tail was (fig. 3) In order to accomplish this, the moved forward to a higher position.
leading-edge sweep of the vertical tail was reduced from 52.0° to 35.0' by shearing the original surface about the root chord. Also, the hori- zontal tail was redesigned to move its mean aerodynamic chord forward and upward by changing the plan form and incorporating 20° of positive dihedral. This horizontal tail was mounted on the reduced-sweep verti- cal tail in a T-tail arrangement with the apex of the horizontal tail coincident with the leading edge of the tip chord of the vertical tail.
These modifications reduced the horizontal-tail length from 2 . 8 5 ~ to 2.47c' and increased the horizontal-tail height from 0.97E to 1.22C when referenced to the assumed center of gravity.
Drawings of several additional horizontal surfaces that were inves- tigated are shown in figure 4. Horizontal tail H2 was located below the jet-exit ducts to give 8 tail position below the wing-chord plane.
A horizontal tail H3 having the same plan form as the original hori- was mounted in a position directly beneath the original zontal tail H, horizontal tail and above the wing-chord plane. This tail was given - 1 5 . 0 ' dihedral in order to lower the surface further. A low-aspect- ratio, highly swept, fixed horizontal surface or strake H 4 was also tested in combination with the new T-tail arrangement. This tail also had -15' dihedral.
c Two-engine inlet modifications are shown in comparison with the original inlet i n figure 5. The modified inlet areas were made equal t o that of the basic i n l e t and the ducts were kept open f o r a l l t h e present t e s t s . When t h e i n l e t w a s f i r s t modified (modified i n l e t 1) , t h e accessory housing bodies located inside t h e ducts were r e m n v d t r . ~ These bodies were l e f t f a c i l i t a t e model changes i n t h e i n l e t region.
inasmuch as t h e i r e f f e c t was out of the model f o r t h e remaining tests, The small believed t o be negligible f o r the present investigation.
differences i n c o ~ ~ i g x a t i ~ ~ s ( a z z e s s c q bodies tn or out or the irdet are indicated i n t a b l e I.
modification used) Various modifications t o the wing are shown i n figure 6. These modifications include leading-edge chord-extensions, wing trailing-edge f i l l e t s or extensions, inboard-upper-surface spoilers, and fences located on the lower surface of t h e i n l e t s . The 0.3% chord-extension was extended i n the wing-chord plane, whereas t h e 0.6% chord-extension (previously t e s t e d at CWT) w a s extended along the leading-edge camber l i n e . The 1/2-inch-lower surface i n l e t fences were constructed of 1/32-inch-thick material. For one t e s t a 1/8-inch-wide t r a n s i t i o n s t r i p (made with number 60 carborundum) was located at the 10-percent-chord l i n e of t h e wing upper and lower surface and a l/8-inch band was located on the fuselage 1 inch behind the nose.
Additional information concerning t h e model and the various tails Photographs of the model with the new vertical- i s presented i n t a b l e 11.
and horizontal-tail combination (WFVlH1) a r e presented i n figures 7 and 8.
"he i n l e t modifications are also shown i n these photographs.
TESTS The sting-supported model w a s tested i n t h e Langley high-speed 7- by 10-foot tunnel through a Mach nuuiber range of 0.60 t o 0.92 and through an angle-of-attack range that varied with Mach number because of load l i m i t s of t h e balance (the maximum range being about 0 ' t o 270). The Reynolds number (based on the m e a n aerodynamic chord) varied with Mach
number from about 1.50 x 10 6 t o 2.0 x 10 6 . The Mach range w a s limited
i n some cases by temperature and tunnel power.
Longitudinal s t a b i l i t y t e s t s were made f o r t h e complete model with the varioirs t a i l , w i n g , and i n l e t modifications. Stabilizer effective- ness tests and l a t e r a l derivative t e s t s ( j 3 = 24') were made f o r the com- p l e t e model with the new v e r t i c a l and horizontal tails (WFV1H1).
0 . 0.. . 0.. . 0 . 0 . . . . .e. 0 .
0 . 0 . 0 . . ... 0 . . e . .
b . e 0 . . 0 . . . . 0 . . . ....
0 . 0 . 0 . . . 0 . . . . 0 . .
e. 0.. e . e ... 0 . 0 . . .e
NACA RM L57A3i C&?IE&h .r * CORRECTIONS I Blockage corrections were applied to the results by the method of Jet-boundary corrections to the angle of attack and drag reference 2.
were applied in accordance with reference 3 . Corrections for effects of the longitudinal pressure gradient in the wind-tunnel test section have been applied to the data.
have not been applied except for a fuselage Model support tares base-pressure correction to reduce the drag to a condition of free- No corrections have been stream static pressure at the fuselage base.
From past experience, it is applied for the internal drag of the duct.
expected that the influence of the sting support on the model character- istics with tails %, Hi, and H2 is small with regard to the lift and pitching moment; however, for configurations with tails H3 and H4 the sting effects could be quite large.
The angle of attack has been corrected for deflection of the balance No attempt has been made to correct the data for and sting support.
aeroelastic distortion of the steel model.
PRESENTATION OF RESULTS The results are presented in figures 9 to 17 as follows: Figure Effect of several tail modifications on the longitudinal
aerodynamic characteristics of the model . . . . . . . . . . 9 to 11
Effect of duct-inlet modifications on the longitudinal aerodynamic characteristics of the model . . . . . . . . . .
Effect of wing modifications and fixes on the longitudinal
aerodynamic characteristics of the model . . . . . . . . 13 to 15
Effect of stabilizer deflection on the longitudinal aerodynamic characteristics of the model with the new vertical and horizontal tails (WFVlH1) . . . . . . . . .
Lateral stability derivatives of the model with the new
vertical and horizontal tails (WFVIH1) . . . . . . . . . . . 17
The results a r e presented about a center of gravity at the 28.37-percent ~a Small mean aerodynamic chord of the nominal w i n g (without f i l l e t s ) .
I variations i n model configurations (accessory bodies i n o r out and duct- 1 4 inlet configuration) are shown i n the tabulation of the detailed test program given i n t a b l e I.
DISCUSSION Pitching-Moment Characteristics Presented i n figure 9 are data showing the effect of several t a i l configurations on the longitudinal characteristics of the model over the angle-of-attack and Mach number range t e s t e d . For the complete model with t h e original v e r t i c a l and horizontal t a i l (Tdl?'V&,)r there is and an i n s t a b i l i t y at about a reduction i n s t a b i l i t y at a , = 5 0 t o 7 O Comparison of the tail-on a = 150 for Mach numbers of 0.60 t o 0.85.
and t a i l - o f f configurations shows t h a t the i n i t i a l reduction i n s t a b i l i t y is caused by the wing-fuselage configuration and that the abrupt insta- b i l i t y near the stall r e s u l t s primarily from the downwash at the tail.
Experience has shown t h i s tail-Cm i n s t a b i l i t y t o be peculiar t o configu- rations having t h e tail located above t h e wing-chord plane i n such a c way t h a t the t a i l traverses t h e wing wake and region of maximum downwash at high angles of attack. Tuft probe studies (at very low speed) i n t h e the i n i t i a l reduction i n s t a b i l i t y v i c i n i t y of the model showed t h a t .
may have been influenced by t h e l i f t i n g properties of t h e engine i n l e t s and that the i n s t a b i l i t y at t h e higher l i f t was associated with t h e t a i l entering the high downwash region, The t u f t survey showed t h a t a t high angles of attack t h e downwash angle a t t h e t a i l approached and possibly exceeded the angle of attack, a condition which would make t h e t a i l ineffective i n adding t o t h e s t a b i l i t y of t h e wing-body combination.
The new vertical- and horizontal-tail configuration (WFVlHl) was designed t o move the horizontal tail above the wing wake by reducing ' t h e sweep of the v e r t i c a l t a i l and mounting a modified horizontal t a i l (with reduced sweep and increased dihedral) on t h e v e r t i c a l - t a i l t i p "his combination raised the mean aero- chord i n a T - t a i l arrangement.
dynamic chord of the horizontal tail about 0.25F and moved it forward about 0.38s. This configuration improved t h e s t a b i l i t y a t low angles of a t t a c k (a = 50 t o 70) and delayed the occurrence of i n s t a b i l i t y by an increment of about 8 ' i n angle of attack or 0.15 i n l i f t coefficient The improvement i n l i f t or angle-of-attack range became a t M = 0.60.
Perhaps a more progressively smaller as the Mach number increased.
r e a l i s t i c reference for comparison of improvements i n lift or angle-of- attack range would be the point where the reduction i n s t a b i l i t y is first observed ( f o r example, see f i g . 9(a), M = 0.80), since i n some type of c maneuvers a sudden reduction i n s t a b i l i t y at moderate angles of a t t a c k might be f e l t as pitch-up. From t h i s p o i n t of view t h e improvement i n l i f t or angle-of-attack range would be larger than that previously indicated.
Another, b u t perhaps less p r a c t i c a l , modification w a s t h e a d d i t i o n of a fixed horizontal surface t o t h e lower surface of t h e engine-tail pipes t o provide a horizontal s t a b i l i z i n g surface below t h e w i n g wake.
This surface w a s t e s t e d i n combination with t h e T - t a i l configuration i n
a biplane t a i l arrangement (WFVlH1. + H2 + modified i n l e t 1). It w i l l be
shown later t h a t t h e change i n i n l e t modification f o r t h i s configuration had only a small e f f e c t on t h e r e s u l t s .
The biplane t a i l gave t h e b e s t o v e r a l 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 any of t h e configurations t e s t e d . Note, however, that t h e pitch-up at high angles of a t t a c k s t f l l e x i s t s b u t i s delayed t o a higher angle of a t t a c k and l i f t c o e f f i c i e n t . When % w a s t e s t e d i n combination
with the o r i g i n a l horizontal t a i l (configuration WFVoHo + H2 of
f i g . lo), a very s u b s t a n t i a l improvement over t h e o r i g i n a l configuration w a s obtained, although t h e r e s u l t s were somewhat i n f e r i o r t o those obtained with t h e biplane t a i l configuration having t h e higher hori-
zontal t a i l (configuration W F V l H l + H2 of f i g . 9). This condition
probably exists because t h e o r i g i n a l horizontal t a i l encounters t h e region of maximum downwash a t a much lower angle of a t t a c k i n such a way t h a t t h e t a i l n o n l i n e a r i t i e s do not cancel t h e wing n o n l i n e a r i t i e s .
The r e s u l t o f lowering t h e o r i g i n a l h o r i z o n t a l - t a i l plan form t o a position j u s t above t h e wing-chord plane w a s detrimental. (See configu- r a t i o n WFVoH3 + H2 i n f i g . 10.) The configuration with only t h e hori- zontal t a i l H2 experiences r a t h e r l a r g e l o s s e s i n s t a b i l i t y at moder- ate angles of a t t a c k ( f i g s . l O ( a ) and 1 0 ( b ) ) . A t t h e higher angles of a t t a c k , however, t h e s t a b i l i t y increases as t h e t a i l (H2) emerges from t h e wing wake. The addition of t r a n s i t i o n t o t h e leading edge of t h e wing did not appreciably a f f e c t t h e pitching-moment c h a r a c t e r i s t i c s ( f i g . 10). The e f f e c t s on s t a b i l i t y of t h e highly swept horizontal t a i l or s t r a k e H4 ( f i g . 11) were small.
t a i l assembly (WFViH1) w a s t e s t e d The configuration with t h e new with various wing and i n l e t modifications.
The e f f e c t s of lower-surface i n l e t fences were negligible ( f i g . 11). The i n l e t modifications of f i g u r e 5 were made i n an e f f o r t t o reduce t h e l i f t i n g e f f i c i e n c y of t h e i n l e t s i n the hope o f improving t h e t a i l - o f f c h a r a c t e r i s t i c s as w e l l as t h e downwash at t h e t a i l . Some s t a b i l i t y improvements r e s u l t e d from t h e modified i n l e t s , although they were r a t h e r small.
(See f i g . 12.)
Some improvement i n t h e pitching-moment-curve l i n e a r i t y w a s obtained, however, with t h e wing trailing-edge f i l l e t s . (See f i g . 13.) No improvement was obtained with t h e wing trailing-edge s p o i l e r o r with a. a a a 0 a a a. e . e e . . ..a a a e . . . a . . a . e e . a . . a a a a * . a * . a .
a a . e a . . .
a . 0 .
a a a a 0 . a .
. . a a * a 0 a .a. a.
a. .a. e . a a a e 0 CI RM L57A31 CONFIDENT&a .
I ' the tail assembly moved forward (fig. 1 4 ) .
A somewhat larger gain in pitching-moment-curve linearity was obtained with chord-extensions running from 0 . 6 9 to the wing tip (fig. 15); however, they provided .
little change i n the angle of attack at which instability occurs.
Since the new vertical- and horizontal-tail configuration (WFV1Hl) showed considerable improvement in longitudinal stability characteristics,
it ....,. +L.....-L+ r -
wc;L3 lrwwu t , u be ilesir~~ble to &temihie the effect of stabilizer deflection on the aerodynamic characteristics.
These results are shown in figure 1 6 . The small nonlinearity at M = 0 . 8 0 and M = 0.85 is magnified somewhat by the - 6 ' stabilizer deflection; otherwise, the results are typical of the usual stabilizer effects.
Lift and Drag Characteristics Use of the new tail assembly did not appreciably affect the VIHl lift characteristics; however, addition of the horizontal surface $ 9 and 1 0 ) extended the lift coefficients to below the tail pipe (figs.
higher values at subcritical m c h numbers. Addition of wing trailing- edge fillets or leading-edge chord-extensions gave small increases in c lift-coefficient range (figs. l l ( c ) , l3(c), and l?(c)) at some E i l a c h numbers.
The configuration with the new tail assembly had a mini- WFVIHl mum drag approximately equal to that of the basic configuration WFVo&.
Drag values obtained for configurations having H2 below the tail pipes are regarded as unrealistic since no attempt was made to obtain a clean installation of this tail on the model.
Small changes in drag were obtained when trailing-edge fillets and leading-edge chord-extensions were used.
Lateral Derivatives Lateral stability derivatives (from tests at j 3 = +bo) were obtained with the new tail assembly WFVlHl as shown in figure 17. These results show that the model is directionally stable through the angle-of -attack range tested, although there is considerable reduction in stability at the higher angles of attack. The effective dihedral increases with angle of attack.
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;al*.lMfl ........ .......
NACA RM L57A3l 1 0 SUMMARY O F RESULTS An investigation of the s t a b i l i t y characteristics a t high subsonic speeds (Mach numbers of 0.6 t o 0.92) of a model of a twin-engine f i g h t e r - type airplane indicate the following r e s u l t s : The r e s u l t s showed that a new horizontal t a i l raised t o the top of the v e r t i c a l t a i l and moved forward from the o r i g i n a l position gave sub- s t a n t i a l improvements i n s t a b i l i t y and delayed the angle of attack and l i f t coefficient at which pitch-up i n s t a b i l i t y occurred. Use of a fixed horizontal surface, attached below the engine t a i l pipes, i n combination with either t h e original or new horizontal t a i l (biplane - arrangement) a l s o provided substantial improvements i n s t a b i l i t y ; however, the com- bination that included the new horizontal t a i l had b e t t e r characteris- t i c s . Some improvement w a s obtained when wing leading-edge chord- extensions, modified wing trailing-edge f i l l e t s , and modified engine i n l e t s were used.
Langley Aeronautical Laboratory, National Advisory Committee f o r Aeronautics, Langley Field, V a . , January 7, 1957.
REFERENCES A Summary of t h e Longitudinal and Lateral Sta- 1. Mitcham, Grady L.: b i l i t y and Control Characteristics Obtained From Rocket-Model Tests of a Swept-Wing Fighter-Type Airplane a t Mach Numbers From 0.5 t o 1.9. NACA R M ~ 5 6 ~ 1 9 , 1957.
2. Herriot, John G.: Blockage Corrections f o r Three-Dimensional-Flow Closed-Throat Wind Tunnels, With Consideration of t h e Effect of Compressibility. NACA Rep. 995, 19%. (Supersedes NACA RM ~ 7 ~ 2 8 . ) 3 . G i l l i s , Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.: Charts f o r Determining Jet-Boundary Corrections f o r Complete Models NACA W R L-123, i n 7- by 10-Foot Closed Rectangular Wind Tunnels.
1943. (Formerly NACA ARR L5G31.)
d 1 1 TAgLE I.- MOD= CONFIGURATIONS TESTED ~~ ~ Accessory housing Duct inlet bodies Original out Originai out Modification 1 out Original In Original out WoHo Original out m'0H2 Original out wFVoH,+ H2 Original out W o H 3 + % W0% + 0.10~ transition Original out ~~ out Original W l H l In Original W l H l + H4 W I H l + H4 + Lower surface inlet fences In Original W I H l + Hb + Wing fillet 2 In Original Original out Modification 1 out Modification 2 out ~~ ~ Original In Original In W F + Wing fillet 2 Original out 13 WIHl Original In W I H l + Wing fillet 2 Modification 1 out W I H l + Wing fillet 3 Modification 1 out WIHl 14 (Tail moved forward 2 inches) out Original W I H l W I H l + Trailing-edge spoiler out Modification 1 ~ Original out W l H l W I H l + 0 . 6 9 chord-extension (CWT) Original out Original In
wlffi + 0.39 chord-extension
Original out ~ ~ 5 ; it = OO
Original out WFVIHl; it = - 6 '
I WFVIHl; it = -6; Lateral derivatives Original Out
1 7 .
NACA RM L 5 7 A j l ' bD d a a rlrl a k Ld 0 ) k sa, r - r - * * * rl rl rl t V 4 d .ri c, c, c, rl f N k k k v3 9 4 co rl f rl t a , 2 M N
H 2
i 2 N (u f cu t- a3 cu m N cv . . .
0 3 0 , .
....... ...............
. . . . . . . . . . . . . . . .
....
... : C O N F ~ " , : : ' ::* :: NACA RM L57A31
........................ *
.
Side force -- Drag Yowinq moment
1:'
Figure 1.- System of axes. (Positive values of forces, mments, and angles are indicated by arrows.)
..
e e. e c e e NACA R M L57A31 e . = e eee*cmII1GmIAc : *::: e. e.. e e .e e. e e e.. e. e.. e.
- 1040
1. I749 CG ut .2657E - I972 A l l dimensions are Figure 2 . - Two-view drawing of basic model (WFVJ3,).
in inches.
p 4 . 5 8 -
-7-7 f
I
I /
/
I 9.45
\I 7
"orig. tail 13.05 14.25 Figure 3.- Two-view drawing of a new v e r t i c a l and horizontal T - t a i l A l l dimensions are i n inches.
arrangement (WFVlH1).
............... .......
. . . . . . . . . . . . . . . .
NACA RM L57A31 f : .
k7 -IC
: .
I f r, I E : rl I al n rd P a .
.
z z .
a
s I
#- i o *
,r\
\
.. e . . . ..I .. *a. * e * .*
a * * ... * . . . * .. . *
NACA R M L57A31 a a , a I d V I t - a , k 2l .d Fr N k
-7
. .
-P k 0, > a , s c, 5 Y =.
x d P c, c2 C d NACA RM L57A31
.lo
.05 -05 -. /O , 2 5 T40 T45
0 5 lo /5 20 25
%de9 (a) C, against a.
Figure 9.- Effect of several tail configurations on the longitudinal aero- dynamic characteristics of the model. it = 0".
NACA RM ~ 5 7 ~ 3 1 (a) Continued.
Figure 9.- Continued.
-05 - * 10
0 5 lo /5 20 25
Q , deg (a) Continued.
FigLre 9.- Continued.
e. e.* e e e e- e. e .*e e * * e e * e . . . e * . e . e . e e . * e
e . .e
e . e c o I v l m ~ e e *. : :* : :* :: e. . . e .e .*e 0 .e e. e e e e.. e.
c -05
-. /o
2 5
-.m
0 5 /o 1 5 20 25
Q, de7 (a) Continued.
Figure 9.- Continued.
0 5 /o /5 20 25
a, deg (a) Concluded.
F i g 3 e 9 - Continued.
0 .2 4 . 6 .8 1.0 /.2 CL (b) C, against CL.
Figure 9.- Continued.
-D5 -. /O -./5 2 5 -.30 1 3 5 ~40 T 4 5 0 .2 4 . 6 .8 L O 1 2 C L Continued.
(b 1 9.- Continued.
Figure .
0 .2 4 . 6 .8 L O L 2 C L (b) Continued.
Figure 9.- Continued.
NACA RM L57A31 0 .2 4 . 6 .8 1 . 0 /.2 CL ( b ) Continued.
Figure 9.- Continued.
a. a a a a a a ,a. a m a a. a. a * a * . . a a a a . a . a a . a . a .
a a .a f = O N B n > ~ a ' 0 : :' : :' : : a a a a. .a. 0. 0.0 0 a a. a. 0 a .a. a.
.
i
o .2 4 . 6 -8 L O
(b) Concluded.
Figure 9.- Continued.
1.2 LO .8 . 6 .2 .8 . 6 .2 0 5 /O /5 20 25 a , deg (c) CL against a.
Figure 9.- Continued.
E . 6 C f L 2 L O -8 . 6 .2
0 5 /o /5 20 25
Q, d q ( c ) Continued.
l.2 L O
.8
CL
. 6
.2
0 5
Q, deg ( c ) Concluded.
Figure 9.- Continued.
-45 .40 .35 .30 -25 ./5 .
c
’D . 1 0 .05 0 .2 4 . 6 .8 L O L 2 CL (d) CD against CL.
Figure 9.- Continued.
NACA R M L57A51 0 .2 4 . 6 . 8 1 . 0 /.2 CL (d) Continued.
Figure 9.- Continued.
.45 .40 .35 . /5
. /o
.05 0 .2 4 . 6 .8 (a) Continued.
Figure 9 . - Continued.
.45 .40 .35 .30 .25 .05 0 .2 4 . 6 .8 L O /.2 CL ( d ) Continued.
Figure 9.- Continued.
-45 -40
.35
./5
. /o
.05
0 .2 4 . 6 .8 LO /.2
(a) Concluded.
Figure 9.- Concluded.
............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
.. 0 . 0 . . ..e . e *
........ .....
.&-&.. NACA RM L57A3i
-05 -. 10 -.I5
c m
:20 : 3 0 (a) C , against a.
Figure 1 0 . - Effect of several tail configurations on the longitudinal aerodynamic characteristics of the model. it = Oo.
(a) Continued.
Figure 1 0 . - Continued.
42 RM
Q , deg
(a) Concluded.
Figure 10.- Continued.
I -
=m
-.IO -.I5 C , ............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . . . . . 0 . . . . . .
........ ..c&@fD&.. .....
.05 -05
-. /o
-.30
.lo
.05 -05 -. /O 2 5 0 2 4 . 6 . 8 L O /.2 CL ( b ) Continued.
Figure 10. - Continued.
. /o
.05
-05
-, 10
-20
-.30
0 .2
4 . 6 .8
.
(b) Concluded.
Figure 10.- Continued.
............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . .
........ .......
.Co+rm NACA RM L57A3i
Q , dep ( e ) cL against a.
ed.
.
??ACA RMI LO
B
. 6
0 5 /O I5 20 25
QJdeP (c) Continued.
Figure 10.- Continued.
NACA LO
. 8
.6
.2
0 5 1 0 15 20 25
a, de7
( c ) Concluded.
Figure 10.- Continued.
.
H NACA RM L 5 7 A 3 1 I .
CL (d) CD against CL.
Figure 10.- Continued.
NACA RM L57A3f a- mea a a a ma a m a mea a a a a m a a a a m o a a a a a m a m a m a m a m a a a m a a a a a m e a a m m a a a .am a m o a a m m a a m a a a m a a a a a a a m .45 .# .35 .30 .25 .20 . /5
. /o
.05 ............... . . 0.. 0 .
. . . . . . . . . . . . . . . .
- . - - - - . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . 0 . .
........ . . C O ~ ~ ~ & . . 0 . . 0 .
NACA RM L57A3f (d) Continued.
Figure 10.- Continued.
0 0 0.0 0 0 0 0 . 0 0 0 0.- 0 0.0 0 0 0 . 0 0 . 0 0 . 0 0 0 . 0 . ..
o m * * 0 0 0 0 0 0 0 . 0 0 0 0 0 0 0 . 0
0 0 0.. ow*mgi$@mo 0: :: : : o . :o*
0 .2 4 . 6 .8 L O L 2 (d) Concluded.
Figure 10.- Concluded.
NACA R M L 5 7 A 5 1 Q , drg (a) C, against a.
Figure 11.- Effect of several tail and wing modifications on the longi- it = 0'.
tudinal aerodynamic characteristics of the model.
I 0. 0.0 . . . * * 0 . . 0.0 . 0.0 0 .
0 . 0 0 . 0 0 . . . . C 0 . 0 .
. . 0 . 0 . . . 0 . . o r . 0 .
...
0 0 e . . = . c & - . . ' : : : :.. : e '
(a) Continued.
Figure 1 1 . - Continued.
(a) Concluded.
Figure 11.- Continued.
m
.
IH .
NACA RM L57A31 l * (b) C , against CL.
Figure l l . - Continued.
............... . . 0.. 0 .
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
0 . 0 . : t '.iJ&p&&l&... . . *
...... ..... NACA RM L57A3i
0 .2 4 . 6 .8 l.0 l.2 CL (b) Continued.
Figure 11.- Continued.
0 .2 4 . 6 .8
CL .
(b) Concluded.
Figure 1 1 . - Continued.
1.2 L O . 8 . 6 .2 L 2 L O . 8 .6 .2 C 0 5 /O /5 20 25 a, deg ( c ) CL against a.
Figure LL.- Continued.
/2 1.0 -8 . 6 .2 1.2 .8 . 6 .2 (c) Continued.
Figure 1 1 . - Continued.
5 1 0 /5 20 25
Q, 0 % ( c ) Concluded.
Figure 11.- Continued.
e . a 0 0 0 0 0 . 0 0 e 0 0 0 0 0 0 - a.
0 . 0 e . . * o o .. 0 0 0 .
. . . e 0 0 0 0 . * e 0 .
e..
0 0 0.0 o&o&* 0 0 . : : : z.0 : : -45 .40 .35 .25 .45 .40 .35 ./5
.30 . /o
25 .05 ./5
. /o
.05 (d) CD against CL.
Figure 1 1 . - Continued.
NACA RM L57Ajl .
.45 .40 .35 .25 . /5
. /o
.05 0 .2 4 .6 .8 L O L 2 CL ( d ) Continued.
Figure 11.- Continued.
.
-45 .40
.35
. /5
. /o
.05
L O /.2
0 .2 4 . 6 . 8
CL (d) Concluded.
Figure 11.- Concluded.
(a) C , against a* Figure 12.- Effect of duct-inlet modifications on the longitudinal aero- it = Oo.
dynamic characteristics of the model.
0 WFL;H, +Mod inlet f -05 -. /O -./5 r 2 5 .IO .05 - 0 5 -. /O 7 2 5 .
(a) Cont h u e d .
Figure 12.- Continued.
NACA RM L57A31 6 8
Q, deg
(a) Concluded.
Figure 12.- Continued.
D5 -05
-. ro
a
(b) C, against CL.
Figure 72.- Continued.
0 .2 4 . 6 .8 1.0 /.2 CL ( b ) Continued.
Figure 12.- Continued.
7 1 I
, -
.05
-05
-. /O
e
-30
0 .2 4 . 6 .8 L O /.2
.
(b) Concluded.
Figure 12. - Continued .
............... . . 0.. 0 .
0 . 0 . 0 . . 0 . . 0 . . 0 . .
. . . . . . . . . . . . . . . . . .
0 . 0 . 0 . . 0.. . 0 . .
........ ..c;m~*&.. .....
NACA RM L57A51 Q , dep (e) CL against a .
Figure 12.- Continued.
.OK 'NACA RM L57A31 Q, d p s (c) Continued.
Figure 1 2 . - Continued.
e . e . . e e . . e e . e . e e e e . . e .
e . e . e . e e.. e.. e..
* e e . . e . . e e e . e e e . . .
* e . e e . .*e e : : o e e C b P J F m G N T a I a e e e . : ' e : NACA RM L57A31 L 2 .8
. 6
.2
5 1 0 15 20 25
%deg
( c ) Concluded.
Figure 12.- Continued.
.45 .# .35 .45 .4Q .35 .20 .
.30 ./5 .25
. /o
.05 ./5
. /o
.05 0 .2 4 . 6 .8 L O 1 2 C L .
( d ) CD against CL.
Figure 12.- Continued.
............... . . 0.. 0 .
. . . . . . . . . . . . . . . .
0 .2 4 . 6 .8 1.0 /.2 CL (d) Continued.
Figure 12.- Continued.
....... ...............
. . . . . . . . . . . . . . . .
..... ............
e . .
e* e . . *co&amu\r: . : : : : :.. : e * 77
.35
./5
.05
0 .2 4 . 6 .8
.
(d) Concluded.
Figure 12.- Concluded.
-05 -.IO -.I5 0 5 IO I5 20 25 Q, deg (a) C , against a .
Figure 13.- Effect of wing-trailing-edge fillets on the longitudinal aero- dynamic characteristics of the model.
a a a. a m a a a a a a .
. a .
m a .a. : $ P 1 s F w a : : : : :a* :aa
u -05 -. /O 2 2 0 -25
.05 a
0 -.35 -05 -. /O 7 . 2 0 2 5 -30 0 .2 4 . 6 .8 /.O 1 2 c, (b) C , against CL.
Figure 13.- Continued.
0 . 0.. . 0.. . 0 . 0 . . . 0.. 0 .
0 . 0 . .. . 0 . . 0 . . ...
0 . 0 . . ... . . 0 . . . . . . e
0 . 0 . 0 . . 0.. . 0 . .
0 . ... . . . ..;mIDE&. 0.. 0 .
Q , de7 ( e ) CL against a .
Figure 13.- Continued.
I - .45 .40 .35 . /5
. /o
.05 C , 0 .2 4 . 6 .8 L O L 2 CL (d) CD against CL.
Figure 13. - Concluded.
a2 (a) C , against a.
Figure 14.- Effect of t a i l p o s i t i o n and s p o i l e r s on t h e longitudinal aero- dynamic c h a r a c t e r i s t i c s of t h e model.
c .05 -05 -.IO -./5 ./o .05 -30 -05 l !
-. /O 2 5 7.35 0 .2 4 . 6 .8 L O /2 C L (b) C, against CL.
Figure 14.- Continued.
a4 a, deg ( c ) cL against a.
Figure 14.- Continued.
t CL t (a) cD against cL.
Figure 1 4 . - Concluded.
, . . . . . . . . . . . . . . . . ............... .......
. . . . . . . . . . . . ....
a . 0 . 0 . .
NACA RM L57A31 .......... ' . c W J P ~ T r p j * , : **: c r n ( a ) C, against a.
Figure 15. - Effect of chord-extensions on t h e l o n g i t u d i n a l aerodynamic
c h a r a c t e r i s t i c s of t h e model.
.
-05 -.IO f20 2 5 -30 -io5 -.IO -20 2 5 - 3 0 0 5 /O I5 20 25 Q, de?
(a) Continued.
Figure 13.- Continued.
, .I NACA RM L57A31 -D5
-. IO
cm
~ 3 0
20 25
0 5 1 0 I5
0 , deg (a) Concluded.
Figure 15. - Cont inued.
.05 .
-05
. / o
.05 -05
-. /o
.
Figure 15 .- Continued.
NACA m ~ 5 7 ~ 3 1
( b ) Continued.
Figure 15.- Continued.
CL ( b ) Concluded.
Figure 13.- Continued.
NACA RM L57A31 0 5 /O /5 20 25 Q I d e g ( c ) CL against a.
Figure 15. - Continued.
o o 0.0 0 0 0 o o oo 0 o o o a 0 0 0 0.
0 . 0 0 . 0 0 . 0 0 0 0 0 0 0 0
. o o o . 0 0 0 0 0 0 0 0 . . 0 0
0 . 0
a
C L . 6 I - /Z B . 6 ( c ) Continued.
Figure 15.- Continued.
94 NACA
I. 0
.8
.6
.2
( e ) Concluded.
Figure 15.- Continued .
.35 c ./5 . 1 0 .05 0 2 4 . 6 .8 L O l2 CL (a) CD against CL.
Figure 15.- Continued.
NACA RM L57A31
9 6
.45 .40 .35 .05 0 .2 4 .6 .8 1.0 /.2 CL (d) Continued. .
Figure 15. - Continued.
3 H NACA RM L57A31 L I
I
I .
.05
0 .2 4 . 6 .8
CL (d) Concluded.
F i g ~ r e 15. - Concluded.
............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . .
: : ... 'eCO!'RIDSPrT&.o: '0: NACA RM L57A31
98 ......
(a) C , against CL.
Figure 16.- E f f e c t of s t a b i l i z e r d e f l e c t i o n on t h e longitudinal aerody- of t h e model with t h e new v e r t i c a l and horizontal namic c h a r a c t e r i s t i c s t a i l . WFVIHl.
.
-20
. /5
. /o
.05
-05
-. /o
0 .2 4 . 6 .8 L O L 2
CL (a) Continued.
Figure 16.- Continued.
............... .......
. . . . . . . . . . . . . . . .
.........
. . . . . . . ........................ ,'cMIDE~& ': ::
0 .2 4 . 6 .8 l.0 l.2
CL (a) Continued.
Figure 16. - Continued.
0. e.* e e e. *e e e.. e e * * e.
e e . . e .e a%?mENRBG. -. : :. : :e :: .NACA RM L57A31 e - - e e eee e . . e . . e
e m e.. e. e.. e e. e. e e . e e e - e
.
0 .2 4 . 6 -8 l.0 l.2
CL (a) Continued.
I C Figure 16.- Continued.
\
\ NACA RM L57A31 .
1.0 /.2 0 .2 4 .6 .8 CL ( a ) Concluded.
Figure 16.- Continued.
i2 LO .8 . 6 .
1.0 .6 .2
0 5 lo 15 20 25
a, de7 (b) CL against a .
Figure 16.- Continued.
............... . . 0.. 0 .
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . .
........................
CONFIDENTIAL .
Q , de7 ( b ) Continued.
Figure 16.- Continued.
.
......................
. . . . . . . . . . . . . . . . . . 0.
... c(tNFI€.)ENm .... : :' : :' : : NACA 4H ........................
.
(b) Concluded.
Figure 1 6 . - Continued.
.45 .40 .35 .30 .25 .45 .40 CD . /5 .35 . /o .30 .05 .25 . /5
. /o
.05 0 .2 4 . 6 .8 l.0 l.2 L i f t coefficient,CL ( c ) CD against CL.
Figure 16.- Continued.
.
I .45 .40 .35 .30 .45 .40 .20 .35 .15 .30 . 1 0 .05 ./5
. /o
.05 0 .2 4 . 6 .8 l.0 l.2 L i f t coefficient, C , ( c ) Continued.
Figure 16.- Continued.
NACA RM L57A31 10 8 Y L i f t c o e f f i c i e n t , CL ( e ) Concluded.
Figure 16.- Concluded.
I - w NACA - Langley Field, Va.