I ' NASA TP 1316 C. 1
NASA Technical Paper 1316
Wind-Tunnel Investigation at Supersonic Speeds of a Canard-Controlled Missile With Fixed and Free-Rolling Tail Fins
A. B. Blair, Jr.
SEPTEMBER 197 8
NASA
https://ntrs.nasa.gov/search.jsp?R=19780024124 2018-06-21T15:16:28+00:00Z
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NASA Technical Paper 1316
Wind-Tunnel Investigation at Supersonic Speeds of a Canard-Controlled Missile With Fixed and Free-Rolling Tail Fins
A. B. Blair, Jr. Langley Research C e ~ t e r Hamptozz, Virgiltia
NASA National Aeronautics and Space Administration
Scientific and Technical Information Office
1978
r e f e r e n c e d i ame te r , 6.604 c m (2.600 in . )
1
M
drol l
Gyaw
@C
r e f e r e n c e body l e n g t h , 99.060 c m (39.000 in . )
Mach number
f r ee - s t r eam dynamic p r e s s u r e , N/m2 (Psf a )
a n g l e o f a t t a c k , deg
d i f f e r e n t i a l d e f l e c t i o n s of t w o c a n a r d s ( cana rds 2 and 4, shown i n s k e t c h ( a ) ) for r o l l c o n t r o l ; i n d i v i d u a l cana rds a r e d e f l e c t e d i n d i c a t e d amount: n e g a t i v e to p r o v i d e coun te rc lockwise r o t a t i o n when viewed from r e a r , deg
yaw-control d e f l e c t i o n of t w o cana rds ( cana rds 1 and 3 , shown i n s k e t c h ( a ) ) ; p o s i t i v e for l e a d i n g edge r i g h t when viewed from r e a r , deg
m o d e l r o l l a n g l e ; p o s i t i v e clockwise when viewed from r e a r ( f o r @c = Oo, c a n a r d s are i n v e r t i c a l and h o r i z o n t a l p l a n e s ) , deg
r o l l r a t e o f t a i l - f i n a f t e r b o d y ; p o s i t i v e c l o c k w i s e when viewed from r e a r , rpm
s t a t i c l o n g i t u d i n a l s t a b i l i t y para.meter
Canards
I 3
Rear v i e w
Sketch ( a )
3
I
APPAKATUS AND TESTS
per meter
6.6 x l o 6 6 .6 6 .6 6 .6
Wind Tunnel
per foot
2.0 x 1 0 6 2.0 2 . 0 2.0
The investigation was conducted in the low Mach number test section of the Langley Unitary Plan wind tunnel, which is a variable-pressure, continuous-flow facility, The test section is approximately 2.13 m ( 7 ft) long and 1 . 2 2 m (4 ft) square* The nozzle leading to the test section is of the asymmetric sliding-block type, which permits a continuous variation in Mach number from about 1 . 5 to 2 .9 . (See ref. 7 . )
56 .4 68 .5 7 5 . 7 9 8 . 4
Model
1 1 7 8 1 4 3 0 1 5 8 0 2056
Dimensional details of the model are shown in figure 1 (a) and a model photograph is shown in figure 2. tion that consisted of a cylindrical body with canards, aft tail fins, and a tangent ogive nose of fineness ratio 3 .0 . ness ratio of 1 5 . The canards and tail fins had slab cross sections with beveled leading and trailing edges. In order for the model to have a free- rolling tail-fin assembly, the tail-fin afterbody was mounted on a set of low- friction ball bearings and was free to rotate through 3 6 0 0 (lock screw out). For the fixed-tail configuration (lock screw in), the tail fins were locked in line with the canards, For both the fixed and free-rolling tail configura- tions, the canards were deflected to provide roll control and yaw control. The tail fins were not deflected (zero cant angle) and the tail-fin assembly had no braking system.
The model was a cruciform missile configura-
The complete model body had a fine-
339 339 339 339
Test Conditions
Tests were performed at the following tunnel conditions:
1 5 0 1 5 0 1 5 0 1 5 0
Mach number
1 . 7 0 2.1 6 2 .36 2 .86
Reynolds number
The dewpoint temperature measured at stagnation pressure was maintained below 239 K ( -300 F) to assure negligible condensation effects. All tests were performed with boundary-layer transition strips measured streamwise on both sides of the canards and tail fins and located 3.05 cm (1 .20 in.) aft of the body nose and 1 . 0 2 cm ( 0 . 4 0 in.) aft of the leading edges. The transition strips were approximately 0 , 1 5 7 cm wide (0 .062 in.) and were composed of No. 50 sand grains sprinkled in acrylic plastic. (See ref. 8.)
4
I
r e f e r e n c e d i ame te r , 6.604 c m (2.600 in . )
1
M
9
&yaw
r e f e r e n c e body l e n g t h , 99.060 c m (39.000 i n . )
Mach number
f r ee - s t r eam dynamic p r e s s u r e , N/m2 ( p s f a )
a n g l e o f a t tack, deg
d i f f e r e n t i a l d e f l e c t i o n s of t w o c a n a r d s ( cana rds 2 and 4 , shown i n s k e t c h ( a ) ) €or r o l l c o n t r o l ; i n d i v i d u a l c a n a r d s a r e d e f l e c t e d i n d i c a t e d amount; n e g a t i v e to p rov ide c o u n t e r c l o c k w i s e rotat ion when viewed from rear, deg
yaw-control d e f l e c t i o n of t w o c a n a r d s ( cana rds 1 and 3, shown i n s k e t c h ( a ) ) ; p o s i t i v e f o r l e a d i n g edge r i g h t when viewed from r e a r , de9
@C model r o l l a n g l e ; p o s i t i v e clockwise when viewed from r e a r ( f o r @c = 0°, cana rds a r e i n v e r t i c a l and h o r i z o n t a l p l a n e s ) , deg
@ t a i l r o l l r a t e o f t a i l - f i n a f t e r b o d y ; p o s i t i v e c l o c k w i s e when viewed from r e a r , rpm
aCJaC, s t a t i c l o n g i t u d i n a l s t a b i l i t y parameter
Canards
4
Rear v i ew
Sketch ( a )
3
I
APPARATUS AND TESTS
56.4 68,5 75.7 9 8 . 4
Wind Tunnel
11 78 1 4 3 0 1 5 8 0 2056
The investigation was conducted in the low Mach number test section of the Langley Unitary Plan wind tunnel, which is a variable-pressure, continuous-flow facility. The test section is approximately 2.13 m ( 7 ft) long and 1 . 2 2 m (4 ft) square. The nozzle leading to the test section is of the asymmetric sliding-block type, which permits a continuous variation in Mach number from about 1.5 to 2.9. (See ref. 7.)
Model
Dimensional details of the model are shown in figure l(a) and a model photograph is shown in figure 2, The model was a cruciform missile configura- tion that consisted of a cylindrical body with canards, aft tail fins, and a tangent ogive nose of fineness ratio 3.0. The complete model body had a fine- ness ratio of 1 5 . The canards and tail fins had slab cross sections with beveled leading and trailing edges. In order for the model to have a free- rolling tail-fin assembly, the tail-fin afterbody was mounted on a set of low- friction ball bearings and was free to rotate through 360° (lock screw out), For the fixed-tail configuration (lock screw in), the tail fins were locked in line with the canards. For both the fixed and free-rolling tail configura- tions, the canards were deflected to provide roll control and yaw control, The tail fins were not deflected (zero cant angle) and the tail-fin assembly had no braking system.
Test Conditions
Tests were performed at the following tunnel conditions:
Stagnation Machr- ~~ temperature
1.70 339 1 5 0 1.70 339 1 5 0 1 5 0 1 ::9” 1 5 0
2.1 6 1 2.36 2.86 1 339 1 1 5 0
. -.
Reynolds number
per meter
6 .6 x l o 6 6 .6 6 . 6 6 .6
per foot ~-
2.0 x 1 0 6 2.0 2.0 2.0
The dewpoint temperature measured at stagnation pressure was maintained below 239 K (-300 F) to assure negligible condensation effects. All tests were performed with boundary-layer transition strips measured streamwise on both sides of the canards and tail fins and located 3.05 cm (1 .20 in.) aft of the hotly nose and 1 . 0 2 cm (0 .40 in.) aft of the leading edges. The transition strips were approximately 0.157 cm wide ( 0 . 0 6 2 in.) and were composed of No. 50 sand grains sprinkled in acrylic plastic. (See ref. 8.)
4
i Y i The pr imary method f o r c o n t r o l l i n g t a i l - f i n r o t a t i o n a l speed was by l i m i t -
i ng t h e model a n g l e of a t tack- I n t h e e a r l y s t a g e s of t h i s test program, t a i l - f i n r o t a t i o n a l speed was nominal ly l i m i t e d to 200 rpm as a s a f e t y p r e c a u t i o n ; however, t h i s l i m i t w a s ex tended to 500 rpm as more conf idence was ga ined- o r d e r to s a t i s f y t h e s e l i m i t s , o n l y small canard d e f l e c t i o n s were made.
I n
Measurements
Aerodynamic f o r c e s and moments on t h e model were measured by means o f a six-component e lectr ical s t r a i n - g a g e ba lance which was housed w i t h i n t h e model. The ba lance w a s a t t a c h e d to a s t i n g which was, i n t u r n , r i g i d l y f a s t e n e d to t h e model s u p p o r t system. Balance-chamber pressure (base pressure) was measured by means o f a s i n g l e s t a t i c - p r e s s u r e o r i f ice l o c a t e d i n t h e v i c i n i t y of t h e ba l - ance. One l i g h t - e m i t t i n g d iode wi th a p h o t o - t r a n s i s t o r r e c e i v e r pick-up mounted on t h e s t i n g was used i n c o n j u n c t i o n wi th a color-coded r i n g a t t h e base of t h e model to r eco rd t a i l - f i n a f t e r b o d y r e v o l u t i o n s . The accu racy o f t h i s r e c o r d i n g sys tem was 520 rpm. N o a t t e m p t w a s made t o measure t h e a f t e r b o d y torque t h a t was produced by t h e i n t e r n a l b a l l - b e a r i n g f r i c t i o n , v i scous - l aye r s k i n f r i c t i o n , or aerodynamic damping -
Correct i o n s
The a n g l e s o f a t t a c k have been c o r r e c t e d f o r d e f l e c t i o n of t h e ba l ance and s t i n g due to aerodynamic loads . I n a d d i t i o n , a n g l e s o f a t t a c k have been cor- r e c t e d f o r tunnel-f low misal ignment . The d r a g and a x i a l - f o r c e c o e f f i c i e n t d a t a have been a d j u s t e d to f r ee - s t r eam s t a t i c p r e s s u r e a c t i n g over t h e model base. T y p i c a l measured v a l u e s of base a x i a l - f o r c e and d r a g c o e f f i c i e n t s a r e p r e s e n t e d i n f i g u r e 3.
PRESENTATION OF RESULTS
F i g u r e
E f f e c t of f r e e - r o l l i n g t a i l on 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 model wi th z e r o c o n t r o l d e f l e c t i o n a t - @ c = o 0 - . - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 @ , = 4 5 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
E f f e c t of c a n a r d s on 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 model wi th f r e e - r o l l i n g t a i l a t @c = 00 . . . . . . . . . . . . . . . . . . 6
E f f e c t of f r e e - r o l l i n g t a i l on l a t e r a l aerodynamic c h a r a c t e r i s t i c s o f model wi th z e r o c o n t r o l d e f l e c t i o n a t - @ c = o o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 @ , = 2 6 . 6 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 @ , = 4 5 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
5
I
Figure
Effect of canards on lateral aerodynamic characteristics of model with free-rolling tail at @c = 00 . e , , . . , . , , . . , . . , . , 1 0
Roll-control characteristics of model with fixed and free-rolling tail at - @ c = O O . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11 @,=450 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2
Yaw-control characteristics of model with fixed and free-rolling tail at @ C = O O . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 3
Table
Summary of test data from free-rolling tail configuration with - Zero control deflection . . . . . . . . . . . . . . . . . . . . . . . . I Canardoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I1 Two canards differentially deflected 0,5O each for negative roll control.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111
Vertical canards deflected 5O for positive yaw control * . . , . . IV
DISCUSSION
Longitudinal Aerodynamic Characteristics
The longitudinal aerodynamic characteristics of the model with zero control deflection are presented in figures 4 and 5 for In general, at low angles of attack (a 6 4O), both the fixed and free-rolling tail configurations have about the same lift-curve slope
level aG/aC,, At the higher angles of attack for @c = 0°, the free-rolling tail configuration has more nonlinear pitching-moment coefficient characteris- tics with a slight pitch-up tendency and, in general, less restoring moment than the fixed-tail Configuration. These aerodynamic differences between the two configurations for the @c = 45O case (fig. 5) are less pronounced, with the pitching-moment curves.becoming more nearly linear with increases in Mach number for the free-rolling tail configuration+ However, the fixed-tail con- figuration now exhibits the pitch-up tendency that characterized the free- rolling tail configuration at This pitch-up trend is typical for a missile with cruciform tail fins in the x-position (@c = 45O) at supersonic speeds. Flow-field effects, in conjunction with adverse panel-to-panel inter- ference between the windward and leeward tail-fin surfaces, result in a small overall reduction in tail lift capability. This loss of lift for the fixed-tail configuration (@c = 45O) can be seen in the lift-coefficient curves presented in figure 5 and for the free-rolling tail configuration at Visual observation has shown that for generally interdigitated to the canards (x-position) when rotation stops and are therefore in a similar flow environment as the fixed-tail case when This loss in tail lift would account for the pitch-up tendency,
@c = Oo and 45O, respectively.
CL, and stability
@c = Oo.
@c = Oo in figure 4. @c = Oo, the free-rolling tail fins are
aC = 45O*
yaw-control capability than the fixed-tail configuration. Again, the aero lockup is delayed to higher angles of attack. (See table IV.)
CONCLUSIONS
A wind-tunnel investigation was made at free-stream Mach numbers from 1.70 to 2.86 to determine the effects of fixed and free-rolling tail-fin afterbodies on the static longitudinal and lateral aerodynamic characteristics of a cruci- form canard-controlled missile model. The effect of small canard roll- and yaw-control deflections was also investigated. The results of the investiga- tion are as follows:
1. The fixed and free-tail configurations have about the same lift-curve slope and longitudinal stability level at low angles of attack,
2. For the free-rolling tail configuration, the canards provide conven- tional roll control with no roll-control reversal at low angles of attack,
3 . The free-rolling tail configuration reduced induced r o l l due to model roll angle and canard yaw control.
Langley Research Center National Aeronautics and Space Administration Hampton, VA 23665 August 9, 1978
9
REFERENCES
1. Sawyer, Wallace C. ; Jackson, Charlie M., Jr. i and Blair, A. B., Jr. : Aero- dynamic Technologies for the Next Generation of Missiles. Paper presented at the AIM/ADPA Tactical Missile Conference (Gaithersburg, Maryland), Apr. 27-28, 1977.
2. Schult, Eugene D.: Free-Flight Measurements of the Rolling Effectiveness and Operating Characteristics of a Bellows-Actuated Split-Flap Aileron on a 60° Delta Wing at Mach Numbers Between 0.8 and 1.8. NACA RM L54H17, 1954.
3. Regan, Frank J,; and Falusi, Mary E.: The Static and Magnus Aerodynamic Characteristics of the M823 Research Store Equipped With Fixed and Freely Spinning Stabilizers, NOLTR 72-291, U.S- Navy, Dec. 1, 1972. (Available from DDC as AD 751 658.)
4. Regan, F. J.; Shannon, J. H. W.; and Tanner, F. J.: The Joint N.O,L,/R.A.E./W.R.E, Research Programme on Bomb Dynamics. Part 111. A Low-Drag Bomb With Freely Spinning Stabilizers. WRE-Report-904 (WRbD) , Australian Def. Sci. Serv., June 1973.
5. Darling, John A,: Elimination of the Induced Roll of a Canard Control Configuration by Use of a Freely Spinning Tail. NOLTR 72-197, U.S. Navy, Aug. 16, 1972-
6. Mechtly, E. A.: The International System of Units - Physical Constants and Conversion Factors (Second Revision). NASA SP-7012, 1973.
7. Schaefer, William T., Jr.: Characteristics of Major Active Wind Tunnels at the Langley Research Center. NASA TM X-1130, 1965.
8. Stallings, Robert L., Jr.; and Lamb, Milton: Effects of Roughness Size on the Position of Boundary-Layer Transition and on the Aerodynamic Charac- teristics of a 55O Swept Delta Wing at Supersonic Speeds. NASA TP-1027, 1977.
9. Spahr, J. Richard; and Dickey, Robert R.: Wind-Tunnel Investigation of the Vortex Wake and Downwash Field Behind Triangular Wings and Wing-Body Com- binations at Supersonic Speeds. NACA RM A53D10, 1953.
10. Dillenius, Marnix F. E.; and Nielsen, Jack N.: Prediction of Aerodynamics of Missiles at High Angles of Attack in Supersonic Flow. NEAR TR 99 (Contract No. N00014-74-C-0050); Nielsen Eng. & Res., Inc., Oct. 1975. (Available from DDC as AD A078 680.)
11. Hardy, Samuel R.: Subsonic Wind Tunnel Tests of a Canard-Control Missile Configuration in Pure Rolling Motion. NSWC/DL TR-3615, U.S. Navy, June 1977. (Available frcm DDC as AD A044 957.)
10
M
.70
.70
.70
2.16
-
TABLE I.- SUMMARY OF TEST DATA FROM FREE-ROLLING TAIL CONFIGURATION
WITH ZERO CONTROL DEFLECTION
a, deg
-1.9 -. 8 0 1.2 2.2 4.4 6,6 8.9
11.1 13.5
J- 17.9
-2.0 -. 5 -. 1 1.1 2.1 4.5 6.6
-2.4 - * 9 0
.9 2.2 4.4 6.5 8.8 J.
17.8
-1.2 .l
1 .o 2.2 3.3 5.5 7.7 J. 24.7
15
0
la i l - f in r o l l ra te , rpmi
Counterclockwise
1 1 5 122 1 1 5 127 97 88 80
0 0 0
0
108 133 121 127 116 12
116
105 7 1 2 123 1 1 2 124
0 0
21
0
120 114 112 1 1 0 96 75 0
0
aWhen viewed from t h e rear.
Remarks
Stopped ro l l i ng 9ero lockup Jery small o s c i l l a t i o n angle
Rotated very slowly Roll r a t e apparently increasing w i t h a
Stopped ro l l i ng Very small o s c i l l a t i o n angle Rotated very slowly
Rero lockup
Stopped ro l l i ng ; aero lockup
1 1
TABLE I.- Continued
- 1 ---I Remar k s M a, @c, Tail-fin r o l l rate, rpma deg deg Counterclockwise
2.16 -1.0 26.4 121 -. 1 1 2 2 .9 1 3 0
2.1 107 3.2 96 5.4 0 Stopped rolling 7.5 0 7.8 1 9 9 Roll rate apparentljr increasing with a
-
2.16 -1.4 45 1 0 0 -. 1 1 0 4 1 .o 99 2.1 1 0 0 3.2 87 5.4 0 Stopped rolling 7.5 0 9.9 1 1 4 Started rolling
12.0 1 2 8 14.1 195 Roll rate increasing with a
2.36 -1 .5 0 1 4 3 -. 2 129 .9 83
2,o 78 2.9 72 5.2 37 7.3 27 9.6 0 Stopped rolling; aero lockup J-
23.7 0 Large oscillation angle
2.36 -1.5 26.6 80 0 94
.9 98 2.0 61 3.1 0 Stopped rolling 5 * 3 0
1 9 4 Roll rate apparently increasing with a I L - -
aWhen viewed from the rear.
1 2
TABLE I.- Continued
M
2.3t
2.8E
2.8t
3, de9
-1 .o .3
1.3 2,4 3.5 5.6 7.7 9.9 12.0 14.4 16.5 18.7
23.8
-2.9 -1 .6 -. 5 .7
1.8 3.8 J. 22.0
-2.8 -1.5 -, 6 .6
1.8 3,7 5.9 8.0
10.0 11.5
J.
IC I geg
45
0
26.5
rail-fin roll rate,
Counterclockwi:
56 70
100 56 54 0 33
118 161 167 122 0
0
23 71 64 62 36 0
0
33 49 51 0 50 0 0
131 0
230
SWhen viewed from the rear.
Remarks
Stopped rolling Started rolling Roll rate increasing with a
Stopped rolling; aero lockup
Low roll rates
Stopped rolling; aero lockup
Oscillated; 2 or 3 revolutions Started rolling Stopped rolling
Started rolling Stopped rolling Roll rate apparently increasing with C
13
TABLE I.- Concluded
- M
2.86
-
-2.5 -1.5 -. 5
.7 1 .7 3.9 5.9 8.1
10.3 12.6 14.6 17.0 19.1 20.2
~
45
T a i l - f i n r o l l rate, rpma
Counterc lockwise
27 51 93 50
0 0 0
75 120 1 2 4
0 0 0
1 5 7
aWhen viewed from t h e rear,
Low r o l l r a t
Remarks
3
Stopped r o l l i n g Small o s c i l l a t i o n a n g l e
S t a r t e d r o l l i n g S teady r o l l i n g
Stopped r o l l i n g
S t a r t e d r o l l i n g
1 4
I
M
70
!.1 6
TABLE 11.- SUMMARY OF TEST DATA FROM FREE-ROLLING TAIL CONFIGURATION
WITH CANARD OFF
3, deg
-2.0 -. 9 0 1 .o 2.1 4.0 6,O 8.0
10.0 12.1 14,2 16.4
-1 .6 -. 9 -. 1 1 .o 2.0 3.0 5.0 7.0 9.1
11.3 13.4
J. 23.2
)C leg
0
0
Tail-fin r o l l r a t e , rpma
Clockwise
54 37 39 46 47 31 31
0 0
30 28 26
33 34 31 47 20 30 23
0 0 0
26
27
Remar k s
Very l o w r o l l r a t e s
Stopped and s t a r t ed t o r o l l
Aero lockup
Very l o w and steady r o l l r a t e s
Stopped r o l l i n g Stopped; s t a r t e d for severa l revolu-
t ions a t a very slow r a t e Stopped; s t a r t e d ; o sc i l l a t ed
Rolled hes i t an t ly and i r r egu la r ly
aWhen viewed from the rear .
1 5
I
TABLE 11.- Concluded
I-
a , de9
-1.2 -. 3 . 8
1.8 2.8 4.9 6.9 9.0 J.
23.0
-2.5 J. 5.6 7.8 9.8 J.
21 .6
-
__
0
0
rail-fin r o l l ra te , rpma
Clockwise
74 47 86 52
1 0 2 88 45
0
42
39
28 0 0
0
3When viewed from the rear.
Remarks
Low r o l l ra tes
Stopped; s tar ted; and osci l la ted Rolled hesi tant ly and i r regular ly
Low r o l l ra tes
Stopped rol l ing 3scil lated through smal l angle
1 6
M
I .7
I .7
2,l
TABLE 111.- SUMMARY OF TEST DATA FROM FREE-ROLLING TAIL CONFIGURATION
WITH TWO CANARDS DIFFERENTIALLY DEFLECTED 0.50
EACH FOR NEGATIVE ROLL CONTROL
3 , deg
-2.2 -1.1 0 1.2 2.4 4.5 6.6 8.9
11.1 f 17.9
-2.3 -1 .3 -. 1 1.3 2,2 4.3 f
10.8
-1 * 2 0 1.1 2.2 3.3 5.5 7.6 J. 16.7 18.9
J. 24.8
C l o c k w i s e
98 96 90
100 114 123 131 97 0
0
102 81 83
105 104 128
207
93 97
109 122 136 154 164
138 0
0
' a i l - f i n r o l l ra te , rpma
3When viewed from t h e rear.
R e m a r k s
Stopped r o l l i n g ; aero lockup
Small o s c i l l a t i o n a l ig le
Roll r a t e i n c r e a s i n g w i t h (3;
(3 > 110; rpm > 500
Steady r o l l i n g
Stopped r o l l i n g ; aero lockup
17
TABLE 111. - Continued
-1.3 0 1 .o 2.1 3.3 5,4 7.6 J. 12.0 14.3
J. 24.5
-1 .3 -. 1 .8
2.0 3.1 5.2 7.3 9.6 J. 24.4
-1 .o -4
1.3 2.3 3.4 5.6 7.8 10.0 12.2
24.2 +
45
0
45
Pail-fin r o l l ra te , rpma
Clockwise
82 84 95 95
104 150 207
151 0
0
109 121 103 95
147 123 110
0
0
88 71
105 93
108 168 3 78 156 0
0
~ ~~
3When viewed from the rear.
Remarks
Steady rol l ing
Stopped rol l ing; aero lockup
Stopped rol l ing; aero lockup
Stopped rol l ing; aero lockup
18
TABLE 111.- Concluded
M
?. 86
2.86
I
e9
2.7 1.5 -. 4
- 7 2.1 3.8 5.9 J- 4.7 7.0 J. 2.6
2.6 1 . 5 -. 5
.6 1.7
J. 0.3 2.5 J. 12. E
3. a
bCI fieg
0
45
rail-fin r o l l ra te , rpma ~~
Clockwise
51 67 87
1 0 4 80 99
1 2 3
1 3 3 0
0
84 58 65 71 73
1 0 5
42 0
0
‘When viewed from the rear.
Remarks
Steady rol l ing
Stopped rol l ing; aero lockup
- -
Low r o l l ra tes
Steady rol l ing
Stopped rol l ing; aero lockup
1 9
-
M
- I .70
- 2,16
- 2.36
I . 86
TABLE IV,- SUMMARY OF TEST DATA FROM FREE-ROLLING TAIL CONFIGURATION
WITH VERTICAL CANARDS DEFLECTED 5O FOR POSITIVE YAW CONTROL
-2.2 -1 .1
0 1 .o 2.1
-1 .3 0 1.1 2.2 3 * 3 6.2
_ _ ~
-1 93 - * 2
.9 2,o 3.0 6.7
-2.7 -1 .5
- * 5 .6
1.7 3.9 5.8 8,3
10.3 12.5 14.1
J. 22.7
T a i l - f i n r o l l ra te , rpma
: l o c k w i s e
360 134
1 5 2
~
191
351 1 7 7
55
80 31 4 463
53 240 430 51 7 522
36 1 8 7 360 50 0 590
84 206 439 527 507 35 4
94 0
0
-
Counterclockwise
Roll d i r e c t i o n changed Roll r a t e i n c r e a s i n g wi th a Excessive r o l l r a t e
- ~
Low r o l l r a t e
Excessive r o l l r a t e .. - .. ~ ~
Jery l o w r o l l r a t e bo th d i r e c t i o n s Roll r a t e i n c r e a s i n g w i t h a
Xxcessive r o l l r a t e .-
3011 d i r e c t i o n changed 3011 r a t e i n c r e a s i n g wi th
Stopped r o l l i n g ; " s t a b l e " a e r o lockup
aWhen viewed from t h e r e a r .
20
.... N
n m Y m
a, m -d 3
0 m m a, 4 c 51
21
I T-
I
1 22
II t
23
... ~
. .
. . .
~~~
. . .
. . .
...
. . . . .
. . . , . . , , .
, . . , . I .
I I
m 0
24
!
T a i l
3 F i x e d 7 F r e e
1 2 - 4 0
(a) M = 1.70.
16 20 2 4
Figure 4.- Effect of free-rolling tail on longitudinal aerodynamic char- acteristics of model with zero control deflection at Oc = Oo.
25
I
2 0 0
- 2 0 0
. 2
0
-. 2
Crn
-. 4
-. 6
-. 8
2 0
1 6
1 2
a, d e g 8
4
0
0 - 4
- 2
i~. . ..
2 4 6 a 1 0
I 0 F i x e d 1 0 F r e e
4
3
'D
1
0
1 2 1 4 1 6
(a) Concluded.
F i g u r e 4.- Continued.
26
. 8
. 4 CA
0
a . deg
(b) M = 2.16.
F i g u r e 4.- Continued.
27
2 0 (
m t a i l , rpm (
- 201
.2
0
-. 2 Cm
-. 4
-. 6
-. 8
2 4
20
1 6
1 2
0 , d e g
8
4
0
- 4 - 2 0
zl b '1
'rl
$-
'I, a
#
2 4 6 8
(b) Concluded.
Figure 4.- Continued.
28
T a i l
0 F i x e d 0 F r e e
10 1 2 1 4 1 6
6
5
4
'D
2
1
0
200
@ t a i l . rpm o
- 200
. 2
0
-. 2
Cm
-. 4
-. 6
-. 8
1 2
10
8
6
cN
4
2
c
- 2
C n
D
T a i l
0 F i x e d 0 F r e e
4
U
M = 2.36.
F i g u r e 4.- Continued.
29
I
200
@ t a i l , rpm O
- 200
.2
0
-. 2
Cm
- . E
-. 1
2
2
1
0 2 4
I
(c) Concluded I
T a i l
Q F i x e d 0 F r e e
10 1 4
5
5
4
3 cD
2
1
0
1 6
F i g u r e 4.- Continued.
30
200
* t a i l I rpm 0
- 200
. 2
0
Cm - . 2
-. 4
-. 6
i CL
-t
f
.. ,
.I , . . I '
.. ! , ' I 1:: I , :../ ; /. :!I;
T a i l F i x e d
1 I I i
i I
0
'N
F i g u r e 4.- Continued.
31
- 4 O B ! ! ! - 2 0 2 4 6 a 10
T a i l
0 F i x e d F r e e
1 1 1 6
5
4
3
CO
2
1
3
(d) Concluded.
Figure 4.- Concluded.
32
200
-2oc
.2
0
-. <
Cm
-. 1
-. 1
1
CN
F r e e 1
/
D
- 4 16 20 2 4
I
8
‘A
0
Figure 5.- Effect of free-rolling tail on longitudinal aerodynamic char- acteristics of model with zero control deflection at @c = 45O.
33
200
@ t a i l , rPm o
- 200
.2
0
-. 2 Crn
-. 4
-. 6
-. 8
2 c
1 6
1 2
a , d e g
4
0
- 4
9
- 1
B P
P
d 6 8 1 0
I . ..
0 F i x e d 0 F r e e
1 2 1 4
4
3
* 'D
1
0
(a) Concluded.
Figure 5.- Continued.
34
a I
200
- 20
. 1
I
. #
c m - . 1
-. f
-. e
- 1 . 0
l i
1c
8
6
cN
4
2
0
- 2
3-
E
k
M
Q
D
B
3
T a i l
0 F i x e d 0 F r e e
0
d
I .,..
i,
/
1
' d )J
24 2 8
. a
. 4 c*
0
Figure 5.- Continued.
35
2 4 6
(b) Concluded.
1 0
0 0
1 2
T a i l
i x e d r e e
1 4 1 6
6
5
4
3 'D
2
1
0
Figure 5.- Continued.
36
200
@ t a i l , rpm 0
- 200
. 2
C
-. i
Cm - . L
-.
- . I
-1.1
1
1
cN
0 F i x e d 0 F r e e
20 24 28
~
8
'A
0
3 2
F i g u r e 5.- Cont inued.
37
20
- 201
. i
Cin -. 2
-. 4
-. 6
-. 8
2 4
2 0
1 6
4
0
- 4
, .
- .. . . . .
__ ..
f
.
.. .
.
--d
- 2 0 2 4 6 8 L O
(c) Concluded.
5
4
3
CO
?
I
12 14 16
Figure 5.- Continued.
38
200
*tai l rpm 0
- 200
. 2
I
Cm -.:
1
cN
T a i l
F i x e d F r e e
si r’
8 1 2 a . deg
(a) M = 2-86,
8
‘A
0
2
F i g u r e 5.- Continued,
39
I
@ t a i l
2 4 t a
(d) Concluded.
T a i l
0 F i x e d 0 F r e e
. .
....
-
1 0 1 2 1 4 1 6
4
3
2 'D
1
0
Figure 5.- Concluded.
40
200
‘ t a i l rpm
-200
. 2
0
-. 2
-. 4
-. 6
Cm
-. 8
- 1 . 0
-1 .2
- 1 . 4
I (
t
t
cN ‘
- <
C a n a r d
0 On 0 O f f
- 4 0 8 12
( a ) M = 1.70.
II ! i ! I i
!
.i
+
/
A
~
!
16 20
j .L
2 4
. 8
. 4 CA
0
Figure 6.- E f f e c t of cana rds on l o n g i t u d i n a l aerodynamic cha r - a c t e r i s t i c s o f model wi th f r e e - r o l l i n g t a i l a t @c = Oo.
41
I
200
@ t a i l s rpm 0
- 200
. 2
0
- . 2
- _ 4
cm - . 6
-. 8
- 1 . 0
-1 .2
- 1 . 4
20
1 6
1 2
8
a . d eg
4
0
- 4
- E
!
I 1 i
1 I
I i I
I
2 4 6 8
C t
I I -
i I I
I C a n a r d
0 O n 0 O f f
4
3
I
0
6
(a) Concluded.
Figure 6.- Continued.
42
200
O t a i l . rpm 0
-2oc
. 2
0
-. 2
Cm
-. 4
-. 6
-. 8
- 1 . 0
- 1 . 2
- 1 . 4
l i
I C
8
6
CN
4
2
0
- 7
C a n a r d ,
I O n 5 O f f
- 8 - 4 0 4 8 20 2 4
(b) M = 2.16.
Figure 6. - Continued. 43
200
% a i l . r p m 0
- 200
. 2
0
-. 2
- . 4
-. 6
Cm
- _ 8
-1.0
-1.2
- 1 . 4
2 0
1 6
1 2
e a .
d e g 4
0
- 4
- e
E-l -E
I.
pl
I
Y I
i I
I
i i I- !
i 1
I '
I I : I I
I 10 10
I I ~
I
I
I I I
I
I I
I
I 1
i
i I I i
: a n a r d '
O n O f f
~
I I
i
i
i
I
! I
I 1
I
i
1 I I
I
1
8 10 1 2 14 1 6
Concluded.
Continued.
5
5
4
3
2 cD
1
0
4 4
200
Q t a i l s r p m o
- 200
. 2
0
-. 2
-. 4
-. 6 Cm
-. 8
-1.0
- 1 . 2
-1.4
l i
1c
8 1 4
I : C a n a r
O n O f f
I r -.a - 4 0 4 8
I I ! f I I I
I 1
b i
.I-
C
I ~
I
4
~
i I
! I
i ~
1 2 1 6 2 0
( C ) M = 2.36-
2 4
Figure 6.- Continued.
4 5
. i!l
< E
I I I
I
I
4
0 6
F i g u r e
8 I O
Concluded.
6.- Continued.
~
C a n a r d
~ 0 O n I , 0 O f f
I
I I I 1 i
! i
I I I
j
I
I ~
I
I i 1
i
I
~ I
1 2 1 4 16
5
4
3
cD
1
3
46
C a n a r d 1
0 O n 0 O f f
(d) M = 2.86.
Figure 6.- Continued.
47
1
4
4
j !
i I
I !
h, !
J
$5 i i I j
\ 1
1 ,
I I i
C a n a r d
On 0 O f f
I !
I I
i I I
i i . 1
Y
1
~
~
I I I
I I
4 6 8 CL
I
I
I I
I
1 :
i I I
j I
+
I. I
I -1 -
I !
I
- 1 - 1
I
I
I
c
. .
4
3
cD
1
0
6
(d) Concluded.
F i g u r e 6.- Concluded.
48
- . ..
[ .. .
i !
i
I I
i j I i j I I I I I I. - i
i i ~
I
I 1 T a i l
0 F i x e d 0 F r e e
I i . 2
0
.. 2
-c
T !
-rr, i I
I .I_.
a. d e g
M = 1.70.
Figure 7.- Effect of f ree-rol l ing t a i l on l a t e r a l aerodynamic character- istics of model w i t h zero control deflection. a t aC = Oo.
49
1 2 16 20 2 4 2 8
a. d e g
(b) M = 2-16.
F i g u r e 7.- Continued.
50
20(
1
C" 0
- 1
CY ' .. i
a 1 2 a , deg
16 20 2 4
___
1 -- ' I tr I - -
L
- -t-
(e) M = 2.36.
Figure 7.- Continued.
51
3 4 - -c I
.. *. a
M I
I
(d ) M = 2.86.
F igu re 7 . - Concluded.
I
.. i I
-4
c
I -t-
i -
I I .-
I i.
! ~
j t - - I I
_ -
I I
- I
i
. - ! i . ..
. 2
0
- . 2
52
I - .-.
2 0
@ t a i l . rpm
- 20
C" (
- 1
2
0
- 2
!- I
i I
i I i I
i ' I
1 i I i -
I I
I.
. i ' 1. i i I i !
. I ' ..
I
! .. ..
I
.1.
E/; T a i l
F i x e d F r e e i
I
+ I
I
.. .
6
3
I <.
0
0 4 a. d e g
(a) M = 1.70 .
. 2
0
-. 2
2
Figure 8.- Effect of f ree-rol l ing t a i l on l a t e r a l aerodynamic character- i s t i c s of model w i t h zero control deflection a t Oc = 26.6O.
53
200
@ t a i l , rpm 0
- 2 0 0 ,
I
0
C"
-1
- 2
2
- ? 8 3 2
(b) M = 2.16.
Figure 8 , - Continued.
54
20(
*ta i l . rpm [
- 20(
1
C
c" -1
- 2
I ..
I + -- .. . . I I --c -
I - - .I- -
1 - -L
I - I -
.. .L-
1 i -r I I I .- -* I I - I --
-p , . . -
3-L i
I
I
. 2
~ O C 1
I - . 2 I . . .
I I
I
0
I a 1 2
a , d e g 2 4 - 4 4 16 20 28 3 2
(c) M = 2.36.
Figure 8.- Continued.
55
200 t
I * t a i l , rpm a
- 20c
T a i
0 F i x e i 0 F r e e
1
0
c" -1
-2
2
0
- 2
j - -I
I -- t-
d
0-
. 2
0
. 2
1 - - I j ... I
L+
I !.:.*: j ' I
I I
" I '-
! '.. - t
i I _- .A . I .
- 4 0 2 4 2 8 32
M = 2.86.
Figure 8. - Concluded.
56
- -. - .. , .. ... , ,
20
* t a i l I rpm
- 2 0
I
c" c
- 1
' - t
I & I 1- T a i l
0 F i x e d 0 F r e e
3
n
i i
KLn
1 I I
. .
I - 1
I 1
1 !
I ! I
1 I I 1
I
!
i i I i I
I !
I
I . + +
I
!
. 2
0
-. 2
1
i
E c c !--
I I
I i
v- I
I
. I I
I QP-
I I I I I
- 4 0 4 8 1 2 a. d e g
16 2 0 2 4 2 8 3 2
(a ) M = 1 .70 ,
Figure 9.- Effect of f ree-rol l ing t a i l on l a t e r a l aerodynamic character- istics of model w i t h zero control def lect ion a t @c = 45O.
57
2 0 0 ,
@ t a i l , rpm o
- 200
1
c" 0
-1
2
0
-2 I i
- 4 0 8 12
.. .
16 20 24 28
. 2
0
-. 2
(b) M = 2.16,
Figure 9.- Continued.
58
20
* t a i l , rpm
- 20
c"
'
T a i l
0 F i x e d 0 F r e e -
G
t- I--
+
B F€
rc
Q
. 2
0
-. 2
€2 i I c
-
D
I
1 ! i
?- i
3
4 8 12 a, d e g
20 2 4 2 8 3 2
( c ) M = 2.36.
Figure 9.- Continued.
59
.I^.. ~
. 4- . 1 . . i
I
. _.._~ I
I
Lr
i - 1
I
I +
I
.. I I + I 1 - .....
...
i i
-1
I
I I
i i t
I .. I ! I
! . L
I
1
~1
I
I I
- . ,
_L
(a) M = 2.86,
Figure 9.- Concluded.
60
-2 - 8
0 0
I I I I I
I I k I I I I
I I - 4
O n O f f
I I I I I 1 I I k
I I I I I
L I
1
I' Y'
I
0
I 4 4
I I I I
I I P I I I I I I I .I,
L
1
r 4
I d
I I I 1 I
I 7 E
T
I
I
I I $ T I
I I I L I r
E 1 1 2 1 6 L"
(a) M = 1.70.
. 2
0
-. 2
!4 2 8
Figure 10.- Effect of canards on l a t e r a l aerodynamic character- istics of model w i t h a f ree-rol l ing t a i l a t ' $c = Oo.
61
- 4 0 4 8 1 2 16 20 2 4 2 8 0 . d e g
(b) M = 2.16.
F i g u r e 10.- Continued.
. 2
0 cz
. 2
62
f
2 0 0
@ t a i l , rpm 0
- 200
1
C n 0
-1
2
CY 0
m 4rl ULdi
C a n a r d
-2 4 a
. 2
O Cz
.. 2
1 2 1 6 2 0 2 4 2 a
(c) M = 2.36,
F i g u r e 10.- Continued.
63
64
- a - 4 0 4 a 1 2 1 6 2 0 2 4 a , d e g
(a) M = 2.86.
Figure 10.- Concluded,
2 8
40
20
@ t a i l , rpm
-20
-40
C" 1
CY
T a i l d r o l l . d e ! 0 F i x e d 0 0 F r e e 0 0 F i x e d - 0 . 5 A F r e e - 0 . 5
( a ) M = 1.70.
2 4 2 8 32
. 2
0 C l
-. 2
F i g u r e 11. - R o l l - c o n t r o l c h a r a c t e r i s t i c s of model wi th f i x e d and f r e e - r o l l i n g t a i l a t $c = Oo. Two c a n a r d s deflected.
65
400
200
otai l I rPm o
-200
-400
CY
(b) M = 2.16.
Figure 11.- Continued.
66
40
20
@ t a i l I rpm
- 2 0
-40
C" I
- 1
C Y (
-,
T a i l broil , d e g 0 F i x e d 0 0 F r e e 0 2 [!;;d -0 . 5
-0. 5
- 4 0 4
i
i A -- rp-
I
I
c
I --
(c) M = 2.36.
Figure 11.- Continued.
67
T a i l , d e g 0 F i x e d 0 0 F r e e 0 0 F i x e d - 0 . 5 A Free - 0 . 5
1 2 a . d e g
(a) M = 2.86.
Figure 11 . - Concluded*
68
400
200
@ t a i l . rpm o
-200
- 4 0 0
1
C" 0
- I
2
CY 0
- 2
T a i l 0 F i x e d 0 F r e e 2 F i x e d
F r e e
0
(a) M = 1.70.
. 2
0 C l
-. 2
3 2
Figure 12.- Roll-control characteristics of model with fixed and free-rolling tail at $c = 45O. Two canards deflected.
69
I
L
(b) M = 2.16.
Figure 12.- Continued.
70
I .'
I '
v
1 I. . .. I - i. ! -i-
I .
--r
.... . + -i-
T a i l tiroll . d e < 0 F i x e d 0 0 F r e e 0 2 LIx,;d -0. 5
- 0 . 5
0 4
M = 2.36.
Figure 12.- Continued.
71
2 8 32
. 2
0 Cl
. 2
(d ) M = 2.86.
F i g u r e 1 2. - Concluded.
72
(a) M = 1 .70.
a
. 4
.2
0
-. 2
-. 4 F r e e 0 F i x e d 5 F r e e 5
-. 6
2 4 28 32
Figure 13.- Yaw-control characteristics of model with fixed and free-rolling tail at @c = Oo. Vertical canards deflected,
73
I
- 8 - 4 a. deg
(b) M = 2.16.
F i g u r e 13.- Continued.
74
J
401
201
I
m t a i l . rpm
-201
-401
-601
1
C"
I
4 8 20 2 4 28
( c ) M = 2.36.
Figure 13.- Continued.
75
(a) M = 2.86.
I'l
T a i l 0 Fixed 0 F r e e 0 Fixed A F r e e
20 2 4 28 32
Figure 13.- Concluded.
76
- .- ~~ . .
1. Report No. 1 2TGoiernmeni Accession No.
. .. - .. . _ ~
NASA TP-1316 . .. ~
4. Title and Subtitle WIND-TUNNEL INVESTIGATION AT SUPERSONIC SPEEDS OF A
TAIL FINS CANARD-CONTROLLED MISSILE WITH FIXED AND FREE-ROLLING
~- . . . . _. . - - .. . .~
7. Author(s)
A. B. Blair, Jr. .. - - . - - -_
9. Performing Organization Name and Address NASA Langley Research Center Hampton, VA 23665
.- - - . - - . - - - - - 7. Key Words (Suggested by Author(s1)
Cruciform missile Canard con tr ol Free-rolling tail
. _ ~ . __ 2. Sponsoring Agency Name and Address
National Aeronautics and Washington, JX 20546
-_ - _ - __-_ 18. Distribution Statement
Unclassified - Unlimited
. . . I -3?-ecipient*r catalog NO.
September 1978 5. Report Date
6. Performing Organization Code
8. Performing Organization Report No.
L-12297 10. Work Unit No.
13. Type of Report and Period Covered Technical Paper
~ . -
14. Sponsoring Agency Code Space Administration
I . - _ _ __- 1 -- ._
5. Supplementary Notes
~.- . . . - . _ _ _ _ _ _ ~ . - _ _ . . -_. . ~~ ~~
6. Abstract
A wind-tunnel investigation was made at free-stream Mach numbers from 1.70 to 2.86 to determine the effects of fixed and free-rolling tail-fin afterbodies on the static longitudinal and lateral aerodynamic characteristics of a cruciform canard- controlled missile model. The effect of small canard roll- and yaw-control deflections was also investigated. The results indicate that the fixed and free- rolling tail configurations have about the same lift-curve slope and longitudinal stability level at low angles of attack. For the free-rolling tail configuration, the canards provide conventional roll control with no roll-control reversal at low angles of attack. The free-rolling tail configuration reduced induced roll due to model roll angle and canard yaw control.
I . . - - -. - .
20. Security Classif. (of this page)
Unclassified .-.- ~
~ _ . -.
~ ~~~~ ~.
9. Security Classif. (of this report)
Unclassified
Subject Category 02 - - . - - -
. .. . ___ $6.00
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* For sale by the National Technical Information Service, Springfield. Virginia 22161 NASA-Langley, 1978
National Aeronautics and Space Administration
Washington, D.C. 20546 Official Business
Penalty for Private Use, $300
THIRD-CLASS B U L K R A T E Postage and Fees Paid National Aeronautics and Space Administration N A S A 4 5 1
USMAIL
I
i 8 1 IU,A, 090878 S00903DS DEPT OF THE AIR F O R C E I A F WEAPONS LABORATORY ATTN: TECHNICAL LIBRARY [SOL) K I R T L B N D AFB N M 87177
NASA poSTMASTER: If Undeliverable (Section 1 5 8 Postal Manual) Do Not Return
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