Document
RESEARCH MEMORANDUM
INVESTIGATION O F AERODYNAMIC CHARACTERISTICS O F AN
AIRPLANE CONFIGURATION HAVING TAIL SURFACES
OUTBOARD O F THE WING TIPS AT MACH NUMBERS O F 2.30, 2.97, AND 3.51 By J a m e s D. Church, William C . Hayes, Jr., and William C. Sleeman, Jr.
Langley Aeronautical Laboratory Langley Field, Va.
CLASSIF!G:;TIC;i CANCELLED
This material contains informationaffecting the National DefeILse o f k e United States within the meaning of the espionage laws, Title 18, U . S . C . , Secs, 793 and 794, the transmission or revelation of which in any manner to an uuautbrized person is prohibited by law.
NATIONAL ADVISORY COM
FOR AERONAUTICS
WASHINGTON 0 0 0 0 0 0 b NATIONAL ADVISORY COMMITTEE FOR AERONAWCS I I d RESEARCH MEMORANDUM INVESTIGATION O F AERODYNAMIC CHARACTERISTICS O F AN AIRPLANE CONFIGURATION HAVING TAIL SURFACES OU'BOARD O F 'JIB WING TIPS AT MACH NUMBERS CIA? 2.30, 2.97, AND 3.51 By James D. Church, W i l l i a m C . Hayes, Jr., and W i l l i a m C. Sleeman, Jr.
SUMMARY 3333'
An investigation has been conducted a t t h e Langley Unitary Plan wind tunnel t o determine t h e drag, s t a t i c longitudinal and lateral sta- b i l i t y , and longitudinal trim c h a r a c t e r i s t i c s of an z i q h r c cc?yfigxa- t i o n having t a i l surfaces outboard of t h e w i n g t i p s . Data w e r e obtained 1 - a t Mach numbers of 2.30, 2.97, and 3.51 a t a Reynolds number of 2.03 x lo6.
Included i n t h e basic data are some e f f e c t s of Reynolds number, engine Y and wing t w i s t combined w i t h toe-out of t h e v e r t i c a l t a i l s . Values pack, of maximm l i f t - d r a g r a t i o a t a Mach number of 2.97 f o r t h e model with the engine pack i n s t a l l e d were about 3.85 and 5.60 f o r s t a b i l i z e r deflec- -0.10 and -&.go, respectively. These values would correspond t o t i o n s of t r i m conditions f o r low-lift s t a t i c margins of approximately 10 and 22 per- cent of t h e mean aerodynamic chord, respectively. With t h e 10 percent s t a t i c margin ( s t a b i l i z e r deflection of -0.lo), however, longitudinal i n s t a b i l i t y occurred above a l i f t coefficient of about 0.20. P o s i t i v e d i r e c t i o n a l s t a b i l i t y of t h e model was p r a c t i c a l l y invariant with angle of a t t a c k t o 12O.
INTRODUCTION Recent experimental and a n a l y t i c a l s t u d i e s (refs. 1 and 2) have indicated t h a t airplane configurations employing h o r i z o n t a l t a i l sur- faces outboard and rearward'of the wing t i p s should result i n an improve- Since t h i s ment i n performance c h a r a c t e r i s t i c s over conventional designs.
geometry l o g i c a l l y results i n twin v e r t i c a l tails, these performance .
gains m i g h t be achieved while retaining adequate d i r e c t i o n a l s t a b i l i t y .
Consequently, as p a r t of a programby t h e National Advisory Committee i
-
f o r Aeronautics t o investigate various configurations with high l i f t - drag r a t i o designed f o r sustained operation near M = 3.0, tests were bJ conducted i n t h e Langley Unitary Plan wind tunnel t o determine t h e drag, s t a t i c s t a b i l i t y , and longitudinal t r i m c h a r a c t e r i s t i c s of an outboard- t a i l model. Results from an investigation of a configuration representing a d i f f e r e n t approach t o t h e general problem of a t t a i n i n g high l i f t - d r a g r a t i o s are reported i n reference 3 .
Data f o r the present tests were obtained a t Mach numbers of 2.30, 2.97, and 3.51 f o r an les of a t t a c k from -4’ t o 16O and f o r angles Of
s i d e s l i p of 4 ’ and -4 . Included i n t h e basic data are some e f f e c t s of
Reynolds number, engine pack, horizontal s t a b i l i z e r , and wing t w i s t com- bined with toe-out of t h e v e r t i c a l t a i l s . These data are presented with- out analysis .
SYMBOLS The forces and moments are reduced t o c o e f f i c i e n t form and are referenced t o the following axis systems: The lateral components are presented about t h e body axes shown i n figure l ( a ) and t h e longitudinal components are oriented with respect t o t h e s t a b i l i t y axes i l l u s t r a t e d i n figure l ( b ) . Moment c o e f f i c i e n t s are taken about an assumed center of gravity located a t 65 percent of t h e mean aerodynamic chord of t h e wing alone (excluding the t a i l s ) .
b span of wing plus horizontal t a i l s , 24.00 i n .
L i f t
lift c o e f f i c i e n t , -
CL qs balance-chamber drag c o e f f i c i e n t ‘D, c ‘D,b engine-pack base-pressure drag c o e f f i c i e n t ‘D,d engine boundary-layer-diverter pressure-drag c o e f f i c i e n t engine-pack internal-flow drag c o e f f i c i e n t ‘D, i T o t a l drag e x t e r n a l drag c o e f f i c i e n t ,
- ‘D,c - ‘D,b - ‘D,i
cD qs Pitching moment pitching-moment coefficient, Cm qsc’ Rolling moment rolling-moment coefficient , s a Yawing moment yawing-moment coefficient, q a Side force side-force coefficient, qs acm
longitudinal-stability parameter, -
aCL stabilizer effectiveness parameter, (%)CL=o effective-dihedral parameter, (%)pEt40 directional-stability parameter,
(2) p*4O
I - I I
side-force parameter, (3)
I p*4O mean aerodynamic chord of wing plus horizontal tails, 12.95 in.
horizontal-tail incidence angle relative to center line of the bodies attached to the wing tips (positive when trailing edge is down), deg lift-drag ratio free-stream Mach number free-stream stagnation pressure, lb/sq ft abs free-stream dynamic pressure, lb/sq ft
-
Reynolds number based on c area of wing plus horizontal tails including wing-body inter- cept (wing-tip and tail-root chords are assumed to lie on the center line of the bodies attached to the wing tips), 1.7391 sq ft
n
a angle of a t t a c k r e f e r r e d t o fuselage reference l i n e , deg P angle of s i d e s l i p r e f e r r e d t o fuselage center l i n e , deg v e r t i c a l - t a i l incidence r e l a t i v e t o center l i n e of t h e bodies 6, attached t o t h e wing t i p s (positive when t r a i l i n g edge i s t o the l e f t ; f denotes toe-out wherein both t a i l s are deflected with trailing edge inboard), deg wing t w i s t of t h e o r e t i c a l t i p chord with respect t o t h e r o o t OW chord about t h e 50-percent-chord l i n e ( p o s i t i v e when t r a i l i n g edge of t i p chord is down), deg Subscripts: min minimum S s t a b i l i t y APPARATUS AND MODEL .
I The t e s t s were conducted i n t h e high Mach number t e s t section of t h e Langley Unitary Plan wind tunnel. J This tunnel i s of t h e variable- pressure, continuous-flow type with a t e s t section 4 f e e t square and approximately 7 f e e t i n length. Mach number may be varied continuously from about 2.3 t o 4.8 by means of an asymmetric sliding-block nozzle.
Sketches of t h e model and i t s engine pack are presented i n f i g u r e 2 and the geometric c h a r a c t e r i s t i c s are given i n table I. Photographs of t h e model a r e shown i n figure 3. The cross s e c t i o n of the b a s i c fuselage w a s semicircular from t h e nose rearward f o r about 22 inches, f a i r i n g smoothly from t h i s point t o a c i r c u l a r base. Mounted beneath t h e fuse- lage and extending t o a point f l u s h with t h e model base was a detachable engine pack. (See f i g . 2 ( a ) .) This pack consisted of a two-dimensional s p l i t i n l e t ducted t o exhaust through t h r e e choked nozzles. An i n t e g r a l p a r t of t h e pack was the wedge-type boundary-layer d i v e r t e r located on t h e upper surface of the inlet-duct housing.
The trapezoidal wing had TO0 of sweep a t t h e leading edge and was mounted with its t h e o r e t i c a l root chord on t h e fuselage reference l i n e .
This surface had an aspect r a t i o of 1.0000, a t a p e r r a t i o of 0.3919, a dihedral angle of - 5 . 3 O , and NACA @ A 0 0 4 a i r f o i l sections. Two d i f f e r e n t wings were t e s t e d on the model. One w a s t w i s t e d about the 0.50-chord .
l i n e SO t h a t the incidence between t h e t h e o r e t i c a l t i p and r o o t chords _ - _ - 0 0 0.0 0 0 L The other wing was untwisted (0, = 0'). A slender body was was -2.8'.
affixed t o each wing t i p and hence was inclined by the angle 0 , t o t h e w fuselage reference l i n e . The p r o f i l e of these bodies of revolution con- I s i s t e d of a short c y l i n d r i c a l section i n s e r t e d between an ogival nose and t a i l .
The v e r t i c a l and horizontal t a i l s had trapezoidal plan forms and were swept back 60° a t the leading edge. These surfaces were considered t o be undeflected when alined with the center l i n e s of the wing-tip bodies. All t a i l panels had an aspect r a t i o of 0.9185, t a p e r r a t i o of 0.3069, Oo of dihedral, and NACA 65A003 a i r f o i l sections.
Forces and moments f o r t h e model were measured by means of a six- component i n t e r n a l strain-gage balance. This balance was attached, by means of a s t i n g , t o the tunnel c e n t r a l support system. Included i n the model support system was a remotely operated, adjustable angle coupling which permitted tests t o be made a t various angles of a t t a c k simultane- ously with v a r i a t i o n s i n the angle o f s i d e s l i p .
TESTS T e s t s were conducted f o r a l l configurations through an angle-of- a t t a c k range of approximately -40 t o 16O a t an angle of s i d e s l i p of 0'.
L a t e r a l - s t a b i l i t y derivatives were determined from tests made through t h i s angle-of-attack range f o r angles of s i d e s l i p of about 4 ' and -bo, with and without the engine pack, a t M = 2.97. Tests t o determine sta- b i l i z e r effectiveness u t i l i z e d incidence a A l l l e s of -0.lo and -4.9'.
t e s t s except those with t h e untwisted w i n g $ 8 w = 0') were made with the v e r t i c a l t a i l s toed-out (6, = +1.5°).
Average Mach numbers, stagnation pressures, dynamic pressures, and Reynolds number a r e l i s t e d i n the following table: 9, R Pt' M lb/sq f t abs l b / s q f t (based on E )
2.30 1,434 425 2.03 x lo6
2.97 2,050 360 2.03 3 -51 2,726 304 2.03 Stagnation temperature was maintained a t 135O F f o r a l l Mach numbers.
.i A Pressure measurements were recorded during one of t h e tests w i t h t h e engine pack i n s t a l l e d i n order t o obtain t h e drag increments asso- c i a t e d with t h e engine-pack base pressure, i n t e r n a l flow, and boundary- .I l a y e r d i v e r t e r pressure. Data w e r e a l s o obtained a t M = 2.97 on t h e untwisted-wing configuration without t h e engine pack and with a l l t a i l surfaces a t a n e u t r a l s e t t i n g i n order t o e s t a b l i s h t h e e f f e c t of Reynolds number on minimum drag. This test w a s conducted near zero lift 6 6 over a Reynolds number range of 0.51 x 10 t o 6.4 x 10 .
Transition was fixed on a l l configurations by means of roughness s t r i p s placed around t h e fuselage and wing-tip bodies about 2 inches behind t h e noses, and along the 10-percent-chord l i n e s (upper and lower surfaces) of t h e wing and s t a b i l i z e r s .
The s t r i p s were 1/52 inch wide and were formed by embedding N o . 60 carborundum g r a i n s i n a p l a s t i c adhesive. Two d e n s i t i e s were employed: one t e s t u t i l i z e d about l5O g r a i n s per inch of s t r i p ( r e f e r r e d t o as heavy) and a l l o t h e r configurations u t i - l i z e d about 50 g r a i n s p e r inch of s t r i p ( l i g h t ) .
CORREXTIONS AND ACCURACY Tunnel pressure gradients i n t h e region of t h e model have been found .
t o be s u f f i c i e n t l y small so as not t o induce any measurable buoyancy e f f e c t s on t h e model. A l s o , a n g u l a r i t y surveys i n d i c a t e n e g l i g i b l e mis- alinement of t h e flow a t t h e t e s t Mach numbers. I n addition, a l l angles L/ of a t t a c k and s i d e s l i p have been corrected for d e f l e c t i o n of t h e balance and s t i n g due t o load.
!The balance-chamber drag (defined h e r e i n as t h e f o r c e t h a t r e s u l t s from the balance-chamber pressure a c t i n g over t h e e n t i r e cross s e c t i o n of t h e base, including the s t i n g ) has been subtracted from t h e drag r e s u l t s f o r a l l configurations. The following a d d i t i o n a l f o r c e s have been sub- t r a c t e d from the drag r e s u l t s f o r t h e configurations with engine pack: base-pressure drag (pressure f o r c e a c t i n g over a l l of detachable-pack base a r e a except f o r t h e t h r e e e x i t s ) and i n t e r n a l drag ( f o r c e computed from duct and e x i t pressures by using standard momentum-balance equation).
Accuracy of t h e presented data based on balance and t u n n e l c a l i b r a - limits: t i o n i s estimated t o be within the following
M . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . fo.015
a , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . f 0 . 2
p , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . f0.2
i t , d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . f O . l
6,,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . f O . l
.
Bw,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . to.1
V ....... 0 0 . e a e . . 0.
00 a 0 00 -0 m e . - * ..eve o m .............................. O W .
..................................
:*. t ...........
~ ~ o ~ ~ : : " " ........
NACA RM L58Cm ts
............
.
I
C L . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . to.003
C D . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ko.0008
c m . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . kO.003
c l . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *o.0005
C n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +0.001
c y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . f0.002
This t a b l e gives the accuracy of t h e absolute value of the q u a n t i t i e s f o r use i n evaluating the possible e r r o r i n i s o l a t e d data.
Experience w i t h r e p e a t a b i l i t y of data indicates t h a t probable e r r o r s can be con- sidered t o be roughly one-half a s large as the values i n t h e t a b l e .
PRESENTATION O F RESULTS The basic r e s u l t s of t h e investigation are presented i n figures 4 t o 10 and some summary r e s u l t s are contained i n f i g u r e s ll t o 15. A n abbreviated outline of f i g u r e content follows: F i g y Schlieren photographs . . . . . . . . . . . . . . . . . . . . .
Balance-chamber, diverter-pressure, internal-flow, and base-pressure d r a g s . . . . . . . . . . . . . . . . . . . . .
Effect of Reynolds number on minimum drag . . . . . . . . . . . 6
E f f e c t of t r a n s i t i o n density . . . . . . . . . . . . . . . . .
Effect of h o r i z m t a l s t a b i l i z e r With engine pack . . . . . . . . . . . . . . . . . . . . . .
Without engine pack . . . . . . . . . . . . . . . . . . . . .
E f f e c t of wing t w i s t and t a i l toe-out . . . . . . . . . . . . .
S t a t i c lateral s t a b i l i t y . . . . . . . . . . . . . . . . . . .
u .
S t a t i c longitudinal s t a b i l i t y . . . . . . . . . . . . . . . . .
S t a b i l i z e r effectiveness and minlrmun drag . . . . . . . . . . .
1 3 Maximum l i f t - d r a g r a t i o Model with engine pack . . . . . . . . . . . . . . . . . . .
Model without engine pack . . . . . . . . . . . . . . . . . .
SUMMARY OF RESULTS The main results of an investigation a t Mach numbers of 2.30, 2.97, and 3.51 of an outboard-tail configuration a t a Reynolds number of 6 are as foliows: 2.03 x 10
a
Values of maximum l i f t - d r a g r a t i o a t a Mach number of (L/D) max 2.97 f o r t h e model with t h e engine pack i n s t a l l e d were about 5.85 ..I and 5.60 f o r s t a b i l i z e r deflections of -0.lo and -4.9', respectively.
These values would correspond t o t r i m conditions f o r low-lift s t a t i c margins of approximately 10 and 22 percent of the mean aerodynamic With t h e 10 percent s t a t i c margin ( c o n t r o l deflec- chord, respectively.
t i o n of -0. lo), however, longitudinal i n s t a b i l i t y occurred above a lift coefficient of about 0.20. This i n s t a b i l i t y was due t o a s t a b i l i t y loss of t h e wing-body combination and t o an equal degree t o t h e reduction i n the s t a b i l i t y contribution of t h e t a i l surfaces. Twisting t h e wing -2.8' (and consequently deflecting t h e horizontal s t a b i l i z e r an equal amount) i n conjunction with f l . 5 ' toe-out of t h e v e r t i c a l t a i l s increased (L/D)- about 0.3 f o r the model without t h e engine pack a t a Mach number of 2.97.
An i d e n t i c a l increase i n t h i s parasleter r e s u l t e d from t h e addition of t h e horizontal t a i l s t o the configuration with twisted wing and toed-out t a i l s . I n both instances t h i s increase was due t o a decrease i n drag due t o l i f t .
The d i r e c t i o n a l s t a b i l i t y C of t h e model with engine pack was "P about 0.0015 and was p r a c t i c a l l y invariant with angle of a t t a c k t o 1 2 ' .
The values of C were reduced by a constant value of about O.OOO5 by t h e addition of t h e engine pack t o t h e model. The model with t h e engine pack had negative e f f e c t i v e dihedral f o r angles of a t t a c k less than 4.5'.
Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, V a . , March 14, 1958.
REFERENCES 1. Sleeman, W i l l i a m C . , Jr.: Preliminary Study of Airplane Configurations Having T a i l Surfaces Outboard of t h e Wing Tips. NACA RM ~ 5 8 ~ 0 6 , 1958.
2. Spearman, M. Leroy, and Robinson, Ross B.: Aerodynamic C h a r a c t e r i s t i c s of a Canard and an Outboard-Tail Airplane Model a t a Mach Number of 2.01. NACA R M L58I307, 1958.
3. Kelly, Thomas C . , Camel, Melvin M . , and Gregory, Donald T.: An Exploratory Investigation a t Mach Numbers of 2.50 and 2.87 Of a Canard Bomber-Type Configuration Designed f o r Supersonic Cruise F l i g h t . NACA RM L38B28, 1958.
. . .
........................
.
TABIS I.- GEOMETRIC CHARACTERISTICS OF M O D E L Wing p l u s h o r i z o n t a l t a i l s (used i n reduction of d a t a ) :
Area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . 1.7391
Span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . 2.000
Mean aerodynamic chord. ft . . . . . . . . . . . . . . . . . 1.0790
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 2.3000
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.1271
Wing:
Area. s q f t . . . . . . . . . . . . . . . . . . . . . . . . 1.3611
span. ft . . . . . . . . . . . . . . . . . . . . . . . . . . 1.1667
Mean aerodynamic chord. f t . . . . . . . . . . . . . . . . . 1.2409
Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 1 . 0 0 0 0
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.3919
A i r f o i l s e c t i o n . . . . . . . . . . . . . . . . . . . . . NACA 65~004
Twist. deg:
Root . . . . . . . . . . . . . . . . . . . . . . . . . . . 0
Tip . . . . . . . . . . . . . . . . . . . . . . . . . . . . -2.8
Dihedral. deg . . . . . . . . . . . . . . . . . . . . . . . . -5.3
LesdFng-edge sweepback. deg . . . . . . . . . . . . . . . . . 70.0
Volume. cu f t . . . . . . . . . . . . . . . . . . . . . . . . 0.02%
1 - Horizontal or v e r t i c a l t a i l (panel geometry) :
iirez. sq f t . . . . . . . . . . . . . . . . . . . . . . . . . 0.18%
Spm. ft . . . . . . . . . . . . . . . . . . . . . . . . . . 0.4167
Mem 3erodynamic chord. f t . . . . . . . . . . . . . . . . . 0.4962
A.cpect r z t i o . . . . . . . . . . . . . . . . . . . . . . . . 0.9185
Y q e r r a t i o . . . . . . . . . . . . . . . . . . . . . . . . 0.3069
A i r f o i l s e c t i o n . . . . . . . . . . . . . . . . . . . . . NACA 65A003
Tdist. deg . . . . . . . . . . . . . . . . . . . . . . . . . 0
Dihedral. deg . . . . . . . . . . . . . . . . . . . . . . . . 0
Leading-edge sweepback. deg . . . . . . . . . . . . . . . . . 60
0.0013 ‘Jclme (exposed). cu f t . . . . . . . . . . . . . . . . . . .
B._tsic fuselzge:
IEngth. f t . . . . . . . . . . . . . . . . . . . . . . . . . 2.8057
Fineness r a t i o . . . . . . . . . . . . . . . . . . . . . . . 12.5
Volume. cu f t . . . . . . . . . . . . . . . . . . . . . . . . 0.0694
Wing-tip body : . . . . . . . . . . . . . . . . . . . . . . . . .
h n g t h . f t 2 . m33
FLneness r a t i o . . . . . . . . . . . . . . . . . . . . . . . 16.6667
Volume. cu f t . . . . . . . . . . . . . . . . . . . . . . . . 0.0169
Engine pack:
Base a r e a (excluding t h e t h r e e e x i t s ) . s q f t . . . . . . . . 0.0178
Enclosed volume. cu f t . . . . . . . . . . . . . . . . . . . 0.0181
m a3 a c a3 k m k k
R
c n I ri a J .
a m z d : u I a J rl rn n
P rl 0
v R J
I
(a) Three-view drawing of model.
Figure 2.- General arrangement of o u t b o a r d - t a i l model. A l l dimensions are i n inches.
0 0 0 0 0 0 0 .
NACA RM ~ 3 8 ~ 2 3 0 , :oo .O0 : 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 0.0 0.
k 0 I cu c 4-r CONFIDENTIAL W i t h e n g l n e pack M z 3 . 5 1 W i t h o u t e n g i n e pack o L- 58- 175 (a) a = 0 ' ; 8 , = - 2 . 8 ' ; 6 , = t1.5O; it = -0.1 ; j 3 = 0'.
Figure 4.- Typical schlieren photographs.
NACA RM L58C25 CONFIDENTIAL W i t h e n g i n e pack M-3.51 W i t h o u t e n g i n e pack L-58-176 a = 6 . 0 ~ ; ow = -2.8'; 6 , = d~1.3~; it = -0.1'; p = 0'.
(b) Figure 4.- Continued.
CONFIDENTIAL NACA RM L58C25 CONFIDENTIAL CONFIDENTIAL 1 8 ' 0 , d C D , i o r C D , b Figure 7 . - Variation of balance-chamber, diverter-pressure, internal f l o w , and base-pressure drag coefficients with angle of attack.
8 , = -2.8'; 6 , = f l . 5 ' ; it = -0.1'; p = 0'.
L .OIO cD,min DO5 .OOl X O6 .5 I 2 3 4 5 6 8 1 0 R Figure 6.- Variation of minimum drag c o e f f i c i e n t with Reynolds number (based on F ) f o r the model without t h e engine pack.
M = 2.97; e, = oO; 6, = oO; it = -0.1'; p = 00.
.
-.2 -.1 0 .1 .2 .3 .4 .5 (a) M = 2.30.
.
Figure 7.- Effect of transition density on aerodynamic characteristics in pitch of model with engine pack. 8, = - 2 . 8 O ; 6 , = f 1 . 5 ' ; it = - 0 . 1 ' ; .
p = 00.
.. a * . 0 . . . a .. . .. . am. 0 .
(a) Concluded.
Figure 7.- Continued.
.. e . . . .. . L .. . . ..e e .
(b) M = 2.97.
Figure 7.- Continued.
..
L (b) Concluded.
Figure 7 . - Continued.
( c ) M = 3.51.
Figure 7. - Continued.
. e e . . . e . e e e e . . . e e .
e e e e . e .
e e e .
P C NACA RM ~ 5 8 ~ 2 5 : . * :,. . e e.
(c ) Concluded.
Figure 7.- Concluded.
0 . 0 . . . 0.. . 0 . 0 . . 0.. 0 .
NACA RM ~ 5 8 ~ 2 5 0 . . 0 .
0 . ... .
- : 2 -.l 0 .1 .2 .3 .4 . 5
CL
( a ) M = 2.30.
Figure 8.- E f f e c t of h o r i z o n t a l - t a i l incidence on aerodynamic character- i s t i c s i n p i t c h of model with engine pack.
8, = -2.8'; 6 , = f1.5°j p = 00.
.1 .o C .o .o .
(a) Concluded.
Figure 8.- Continued.
-.2 -.l 0 .1 . 2 .3 .4 . 5 C L (b) M = 2.97.
Figure 8.- Continued.
( c ) M = 3.51.
Figure 8.- Continued.
. e . * e e e
. . e . e . 0.1
. e . . e . 0 . . e e.
e * e .
e . .
NACA RM L38C25 *.
( e ) Concluded.
Figure 8.- Concluded.
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 0 . 0 0 0 0 . 0 0 0 0 . 0 0 0 0 0 0 0 . 0 " NACA R M ~ 5 8 ~ 2 5 (a) M = 2.30.
Figure 9.- Effect of horizontal stabilizer on aerodynamic characteristics in pitch of model without engine pack.
e, = -2.8'; 6 , = f1.5'; p = 0'.
G -.2 -.l 0 .1 - 2 .3 .4 .5 C L (a) Concluded.
Figure 9.- Continued.
( b ) M = 2.97.
Figure 9.- Continued.
( b ) Concluded.
Figure 9.- Continued.
( c ) M = 3.51.
Figure 9. - Continued.
I -
-
(c) Concluded.
Figure 9.- Concluded.
(a) M = 2.30.
Figure 10.- Effect of twist and toe on aerodynamic characteristics in = 0'.
it = - 0 . 1 ' ; pitch of model without engine pack.
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 .
.
(a) Concluded.
Figure 10. - Continued.
............... . . 0.. 0 .
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
(b) M = 2.97.
Figure 10. - Continued.
I C NACA RM L38C25 C .
(b) Concluded.
Figure 10.- Continued.
............... .......
. . . . . . . . . . . . . . . .
. . . . . . . . . . . . . . . . . .
( c ) M = 3.51.
Figure 10.- Continued.
NACA RM L58C25 C (c) Concluded Figure 10. - Concluded.
m m m m m am m. emm m m Figure 11.- E f f e c t of engine pack on t h e v a r i a t i o n of t h e s t a t i c l a t e r a l - s t a b i l i t y parameters with angle of a t t a c k . M = 2.97; 8, = - 2 . 8 ' ; 6, = f 1 . 5 ' ; it = - 0 . 1 ' .
C - L C - L C - L Figure 12.- Effect of engine pack and horizontal tail on the variatioi of the static longitudinal-stability parameter with lift coefficiei e, = - 2 . 8 O ; 6 , = k1.5'; p = o 0 .
0 . 0.. . 0.. . 0 . 0 . . . . ... 0 .
4 6 C mit .
#,mi CD .
M Figure 13.- Variation of stabilizer effectiveness and minimum drag coef- ficient with Mach number. p = 0'.
a . . a a .
. . . a . . * . a ......
a . a . a . a . a . a . .
a 0 . a . . . . . . a a a ..a a . 0 . 47 NACA RM L58C25 a . a a a . a . 0 .
Figure 14.- Variation with Mach number of maximum lift-drag ratio and of lift coefficient for maximum lift-drag ratio.
Model with engine pack; 0, = -2.aO; 6 , = *1.5O; p = o0.
0 . ... 0 .. 0 o r .. 0 0 0 0.0 0 0
M Figure 15.- Variation with Mach number of m a x i m lift-drag ratio and of lift coefficient for maximum lift-drag ratio.
Model without engine pack; p = O o .
NACA - Langley Field, vd.
~~