Wind-Tunnel Investigation a Canard-Controlled Missile · I ' NASA TP 1316 C. 1 NASA Technical Paper...

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

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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 configurations, 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 configuration 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 fineness 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 configurations, 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 characteristics 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 configuration 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 cruciform canard-controlled missile model. The effect of small canard roll- and yaw-control deflections was also investigated. The results of the investigation 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 conventional 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



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


- -

Low r o l l ra tes

Steady rol l ing


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 characteristics 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.


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


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


4

U

M = 2.36.


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


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


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 characteristics 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 . ..


1 2 1 4

4

3

* 'D

1

0

(a) Concluded.


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

d

I .,..

i,

/

1

' d )J

24 2 8

. a

. 4 c*

0


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


36

200

@ t a i l , rpm 0

- 200

. 2

C

-. i

Cm - . L

-.

- . I

-1.1

1

1

cN


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


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


. .

....

-

1 0 1 2 1 4 1 6

4

3

2 'D

1

0


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.


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


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.


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


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 characteristics 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.


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.


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.


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


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 characteristics 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,


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.


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,


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 characteristics 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.


. 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,


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.


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.


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.


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.


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.


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.


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


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

- _ -

..

* 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

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