NASA Technical Memorandum 84565 ,NASA-TM-84565 19830008012 1

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Effect of Modifications onAerodynamic Characteristicsof a Single-Stage-to-OrbitVehicle at Mach 5.9

Peter T. Bernot

JANUARY 1983

N/ A

https://ntrs.nasa.gov/search.jsp?R=19830008012 2019-08-14T13:11:36+00:00Z

NASA Technical Memorandum 84565

Effect of Modifications onAerodynamic Characteristicsof a Single-Stage-to-OrbitVehicle at Mach 5.9

Peter T. BernotLangley Research CenterHampton, Virginia

NI ANational Aeronauticsand Space Administration

Scientific and TechnicalInformation Branch

1983

SUMMARY

Longitudinal and lateral-directional characteristics have been determined for asingle-stage-to-orbit vehicle based on control-configured stability concepts. Theconfiguration had a large body with a small 50° swept wing. Two vertical-fin arrange-ments were investigated: a large center-line vertical tail and small wing-tip fins.The movable outward surfaces of the wing-tip fins (controllers) were designed toprovide yaw control. Tests were conducted in the Langley 20-Inch Mach 6 Tunnel overan angle-of-attack range from 0° to 40° at a Reynolds number of 2.5 x 106 based onfuselage length. Yaw-control effectiveness of the wing-tip-fin controllers wasobtained at a sideslip angle of 0°.

With the center of gravity at 71.5 percent of fuselage length, the model hadabout the same longitudinal characteristics with either the center-line vertical tailor the wing-tip fins. Both configurations could be trimmed longitudinally over anangle-of-attack range from approximately 13° to 40° with zero or positive controldeflection. Both configurations were directionally unstable over the angle-of-attackrange but had positive effective dihedral at angles of attack greater than 8°. Sig-nificant yaw-control effectiveness was indicated when the wing-tip-fin controller wasdeflected outward 40°.

INTRODUCTION

For the past several years, NASA and industry have been studying advancedorbital transportation systems that could provide greater cost and performance advan-tages than the current shuttle system. These studies have dealt mainly with wingedsingle-stage-to-orbit vehicles incorporating improved state-of-the-art technologiesin structures, heat protection systems, rocket propulsion, and guidance and controlschemes. Comprehensive overviews relating to many aspects of this study are dis-cussed in references I, 2, and 3.

In one phase of this study, NASA has focused its attention on two verticallylaunched single-stage-to-orbit vehicles designed for a payload mass of 65 000 lb.Both vehicles had large bodies for internal fuel and payload storage, 50° swept deltawings, a vertical tail, and a body flap. Using a computer-aided design process, thebaseline vehicle was designed to have a positive level of longitudinal stability,whereas the second vehicle was designed for more relaxed stability criteria based oncontrol-configured concepts recently used for transport and fighter aircraft, qhisdesign philosophy is discussed in reference 4, which also included wind-tunnelresults for the two vehicles for a Mach number range from 0.3 to 4.63. Hypersoniccharacteristics of both vehicles are presented in reference 5 at a Mach number of20.3. As indicated in reference 4, both vehicles yielded acceptable stability andperformance; however, the control-configured vehicle had a lower gross lift-offweight. Therefore, attention was refocused on this vehicle, and several modifica-tions were incorporated. The planform areas of the elevons and body flap wereincreased. Also, since the vertical tail was shielded by the body and wing atreentry angles of attack and was therefore ineffective, the tail was removed, andwing-tip fins were added for directional stability.

The purpose of this paper is to present the stability, control, and performanceresults on the single-stage-to-orbit control-configured vehicle at Mach 5.9. Twomodels were tested, one having a single center-line vertical tail and the other hav-ing wlng-tip fins with movable controllers. Static longitudinal and lateral-directional stability characteristics were obtained on the 0.006-scale models over anangle-of-attack range from 0° to 40°. Also determined were the effects on yaw androll of deflecting the wing-tip-fin controllers 20° and 40° at a sideslip angle of0°. This investigation was conducted in the Langley 20-Inch Mach 6 Tunnel at aconstant Reynolds number of 2.5 x 106 based on model fuselage length.

SYMBOLS

b reference wing span

Axial forceCA adjusted axial-force coefficient, qS ' CA,T - CA,B

CA,B base axial-force coefficient, IpB _SB-_qS/

Total axial forceCA,T total axial-force coefficient, qS

CD drag coefficient, Drag forceqS

Lift forceCL lift coefficient, qS

C rolling-moment coefficient, Rolling momentI qSb

CI_ effective dihedral parameter, ACI/A_, per deg

C pitching-moment coefficient, Pitching momentm qS_

Normal force

CN normal-force coefficient, qS

Cn yawing-moment coefficient, Yawing momentqSb

C directional-stability parameter, ACn/A_, per degn_

<C i_)n dyn dynamic directional-stabilityparameter,perdeq Cn_ cOs _ - CI_ x/Sin _,Side force

Cy side-force coefficient, qS

C rate of change of side-force coefficient with sideslip angle, ACy/A_,Y_ per deg

wing mean aerodynamic chord

Iz/Ix ratio of moments of inertia about yaw and roll axes

L/D lift-drag ratio

I fuselage reference length

M free-stream Mach number

q free-stream dynamic pressure

p pressure

R Reynolds number based on fuselage lengthI

S wing total planform reference area

x longitudinal distance measured from theoretical nose of fuselage

angle of attack, deg

angle of sideslip, deg

6 elevon deflection angle, positive for trailing edge down, dege

6F body-flap deflection angle, positive for trailing edge down, deg

6TF tip-fin controller deflection, positive when controller is deflectedoutward, deg

Subscripts:

B base area of fuselage

cg center of gravity

cp center of pressure

free-stream conditions

0.7151 moment center at 71.5 percent of fuselage reference length

DESCRIPTIONS OF MODELS

The two 0.006-scale models were modified versions of the control-configuredvehicle tested previously (refs. 4 and 5). The modifications consisted of reshapingthe upper portion of the fuselage to accommodate relocation of the payload bay andincreasing the planform areas of the elevons and body flap. Both models were fabri-cated from aluminum and were geometrically similar except one model incorporated a

3

single center-line vertical tail and the other had wing-tip fins. (See fig. I.) Themovable outer surfaces of the wlng-tip fins were designed for outward deflections

only. These movable controllers were machined from metal as wedgelike componentssimulating deflection angles of 20° and 40° and were attached to the wing-tip fin bysmall screws. Elevon deflections were set by using pre-bent brackets, whereas body-flap deflections were obtained by using separate interchangeable components for eachdeflection.

Models I and 2 were designated BWVF and BWTF, respectively, where model compo-nents are identified as follows:

B fuselage

W wing

V vertical tail

T tip fin

F body flap

Full-scale dimensions of both vehicles are presented in table I. Photographs of eachmodel mounted in the langley 20-Inch Mach 6 Tunnel test section are shown in figure 2.

APPARATUS AND TESTS

Wind Tunnel

This investigation was conducted in the langley 20-Inch Mach 6 Tunnel which is ablowdown type facility using air as the test medium. The tunnel has a two-dimensional contoured nozzle which is 2.27 m (89.4 in.) long and incorporates a testsection measuring 0.52 by 0.508 m (20.5 by 20 in.). The sting-support system isremotely controlled and can provide an angle-of-attack range from -5° to 50° and asideslip-angle range from 0° to -10°. Stagnation pressures can be varied from0.2 MPa to 3.6 MPa (30 to 525 psia). To prevent liquefaction, the air can bepreheated up to 567 K (560°F) by the use of an electrical resistance heater. Testrun times of 1 minute are possible when the flow exhausts into a vacuum sphere and upto 20 minutes when an annular type air ejector is used. Additional details includingMach number calibrations are presented in reference 6.

Tests

All tests were conducted at a stagnation pressure of 0.689 MPa (100 psia) and atemperature of 453 K (355°F). The average free-stream Mach number was 5.93 at a unitReynolds number of 6.24 × 106 per meter (1.9 x 106 per foot).

Force and moment data were measured by a sting-supported six-component strain-gage balance. Both models were tested over an angle-of-attack range from 0° to 40°in 5° increments at sideslip angles of 0° and -2°. Model attitudes were accuratelyset by using a small prism mounted flush in the fuselage which reflected light from apoint source located near the test section onto a calibration board. Free-streamMach number was determined by using an actuator-mounted pitot probe which was

inserted into the airstream only at the beginning and ending of each blowdown run.This procedure precluded the possibility of interference effects on the model testdata.

The estimated uncertainties in the measured data were based on ±0.5 percent ofthe balance design loads and are as follows:

CN .......... ±0.0115 C ......... ±0.000161

CA .......... +0.0029_ C ......... +0.00027n

Cm .......... +0.0027_ Cy .......... _+0.0024

The accuracy of the angles of attack and sideslip is estimated to be ±0.15°;free-stream Mach number is estimated to be accurate to ±0.02. Base pressure wasmeasured at one location for all tests and was used to adjust the axial-force coeffi-cients to free-stream pressure acting on the base.

RESULTS AND DISCUSSION

Flight Profile

This vehicle was designed for vertical take-off and would employ rocket enginegimballing for attitude control during ascent. For entry, the vehicle would glide atan angle of attack of approximately 36° over the hypersonic regime down to a Machnumber of 8. The angle of attack would then gradually be reduced to 10° at Mach 4and maintained at this attitude until just prior to horizontal landing.

I

Longitudinal Aerodynamic Characteristics

The static longitudinal characteristics of model I are presented in figure 3 fora center-of-gravity (c.g.) location at 71.5 percent of the fuselage length. Thecontrol system (elevons and body flap) was capable of trimming the configuration overan angle-of-attack range from approximately 13° to 40° for various control deflec-tions. These trim conditions were unstable up to approximately _ = 30° and werethen neutrally stable or stable at higher attitudes. As shown in figure 3(d), astable trim condition occurred at about _ = 40° with the body flap deflected 10°.Entry trajectory studies (ref. 7) indicated that the flight angle of attack would be26° for a Mach number of 6. For this angle of attack, neutral stability could beachieved without control deflections if the center of gravity were shifted forward to69.6 percent of the fuselage length. However, an unpublished weight analysis indi-cated that c.g. locations for the control-configured vehicle with the large verticaltail would be more rearward, at approximately 74 percent. Consequently, more forwardc.g. locations would appear unrealistic at the present time for model I. A maximumlift-drag ratio of 2.57 was obtained at _ = 12.5° without control deflections.At _ = 26° , the L/D was about 1.75 with the body flap deflected 10° .

The longitudinal aerodynamic characteristics of model 2 are presented in

figure 4 for various deflections of the tip-fin controllers at zero elevon and body-

flap settings. A comparison of the longitudinal aerodynamic characteristics of

model I and model 2 at zero control deflections shows negligible differences.

Consequently, the previous discussion of stability and control of model I isapplicable to model 2 for the same flight conditions and c.g. location.

Tip-fin controller deflection had only a small effect on the pitching-momentcoefficient (fig. 4(d)). When the right controller was deflected 40°, C at= 26° changed by only -0.003, which is approximately equal to estimate_ balance

accuracy. An even smaller increment in Cm was measured when both controllers weredeflected 20°. The only appreciable effect of tip-fin controller deflection was onthe maximum values of L/D which occurred at angles of attack from 12.5° to 15°, asshown in figure 4(g). At the operational angle of attack of 26°, the measured valueof L/D was approximately 1.77 for the various controller deflections.

Lateral-Directional Aerodynamic Characteristics

The lateral-directional aerodynamic characteristics of models I and 2 are pre-sented in figure 5 for zero deflections on all controls. The data were obtained atsideslip angles of 0° and -2°. Although both configurations were statically unstableat Mach 5.93, the results showed that removing the rather large center-line verticaltail and adding the smaller wing-tip fins actually slightly increased directionalstability. At angles of attack greater than 8°, both models exhibited positiveeffective dihedral /C h withabout the same levels of magnitude. Favorable

%JI _= were obtained only at higher angles of attack for(positive) values o n_/uyn

both configurations. At the flight angle of attack of 26°, the dynamic directional-stability parameter was zero for the center-line vertical tail configuration andslightly positive for the wing-tip-fin configuration.

Lateral-Directional Control Characteristics

In figure 6, the effects of tip-fin controller deflection of the right wing-tipfin on roll-, yaw-, and side-force control parameters were determined at zerosideslip. At the operational angle of attack of 26°, the magnitudes of roll- andyaw-control parameters were quite small for a controller deflection of 20°. However,an increase in the deflection to 40° did result in significant roll- and yaw-controlcapability, especially yaw control. At flight conditions, the ratio of AC tonAC of about 5.0 indicates that the wing-tip-fin controller is basically a yaw-control device. In-house calculations using the reaction-control system values fromfigure 1(b) of reference 7 indicate that the wing-tip-fin controller when deflected40° can produce a yawing-moment increment equal to approximately 1.2 yaw jets, or itcan provide a trim capability for a 1° perturbation in sideslip.

CONCLUDING REMARKS

Aerodynamic characteristics of a single-stage-to-orbit vehicle based on control-configured concepts have been obtained at Mach 5.9 in the Langley 20-Inch Mach 6Tunnel. The configuration had a large body for internal fuel and payload and a small50° swept wing. Two vertical-fin arrangements were investigated: a large center-line vertical tail and small wing-tip fins. Longitudinal and lateral-directionalcharacteristics were obtained over an angle-of-attack range from 0° to 40°.

With the center of gravity at 71.5 percent of the fuselage length, the aerody-

namic control system provided longitudinal trim conditions for both configurations

over an angle-of-attack range approximately 13° to 40 ° . _hese trim conditions were

unstable up to approximately an angle of attack _ = 30 ° and were then neutrally

stable or stable at higher attitudes. A stable trim condition was achieved at

= 40 ° with the body flap deflected 10 °. Both configurations had positive

effective dihedral at angles of attack greater than 8° . Directional instability

occurred over the entire angle-of-attack range because of the rearward center-of-

gravity location. At the flight angle of attack of 26°, the dynamic stability

parameter was zero for the vertical tail configuration and only slightlyTn

positive for the wing-tip-fin configuration. Large outward deflections of the wing-

tip-fin controller were effective for yaw control.

Langley Research Center

National Aeronautics and Space Administration

Hampton, VA 23665

November 30, 1982

REFERENCES

1. Henry, Beverly Z.; and Decker, John P.: Future Earth Orbit TransportationSystems/Technology Implications. Astronaut. & Aeronaut., vol. 14, no. 9,Sept. 1976, pp. 18-28.

2. Henry, Beverly Z.; and Eldred, Charles H.: Advanced Technology and Future Earth-Orbit Transportation Systems. AIAA Paper 77-530, May 1977.

3. Hepler, Andrew K.; Zeck, Howard; Walker, William H.; and Shafer, Daniel E.:Applicability of the Control Configured Design Approach to Advanced EarthOrbital Transportation Systems. NASA CR-2723, 1978.

4. Freeman, Delma C., Jr.; and Wilhite, Alan W.: Effects of Relaxed Static Longitu-dinal Stability on a Single-Stage-To-Orbit Vehicle Design. NASA TP-1594, 1979.

5. Bernot, Peter T.: Aerodynamic Characteristics of Two Single-Stage-To-OrbitVehicles at Mach 20.3. NASA TM X-3550, 1977.

6. Goldberg, Theodore J.; and Hefner, Jerry N. (appendix by James C. Emery):Starting Phenomena for Hypersonic Inlets With Thick Turbulent Boundary Layers atMach 6. NASA TN D-6280, 1971.

7. Freeman, Delma C., Jr.; and Powell, Richard W.: Impact of Far-Aft Center ofGravity for a Single-Stage-To-Orbit Vehicle. J. Spacecr. & Rockets, vol. 17,no. 4, July-Aug. 1980, pp. 311-315.

TABLE I.- GEOMETRIC DETAILS OF FULL-SCALE VEHICLES

Model I (BWVF) Model 2 (BWTF)Body:Length, m (ft) ......................... 66.79 (219.12) 66.79 (219.12)Maximum width, m (ft) ................... 15.93 (52.26) 15.93 (52.26)Base area, m2 (ft2) ...................... 142.61 (1535) 142.61 (1535)Fineness ratioa .......................... 4.955 4.955Planform area, m2 (ft2) .................. 815.21 (8774.9) 815.21 (8774.9)

Wing:Total area, m2 (ft2) 557.42 (6000) 557.42 (6000).eooo.ooo.ooeeoeoooeo

Span, m (ft) ............................. 37.18 (121.98) 37.18 (121.98)Root chord at center line, m (ft) ........ 26.07 (85.53) 26.07 (85.53)Tip chord, m (ft) ........................ 4.82 (15.84) 4.82 (15.84)Mean aerodynamic chord, m (ft) ........... 17.72 (58.14) 17.72 (58.14)Aspect ratio ............................. 2.48 2.48Dihedral angle, deg ...................... 7 7Incidence angle, deg ..................... 1.5 1.5Leading-edge sweep, deg .................. 50 50Trailing-edge sweep, deg ................. 0 0Airfoil section at root ................. NACA 0008-64 NACA 0008-64Airfoil section at tip .................... NACA 0012-64 NACA 0012-64Elevon chord, m (ft) ..................... 3.54 (11.61) 3.54 (11.61)Elevon span per side, m (ft) ............. 10.63 (34.86) 10.63 (34.86)Elevon total area, m2 (ft2) .............. 75.2 (809.4) 75.2 (809.4)

Vertical tail:Root chord, m (ft) ....................... 10.88 (35.71)Tip chord, m (ft) ........................ 5.41 (17.74)Area, m2 (ft2) .......................... 83.61 (900)Span, m (ft) ............................. 10.27 (33.68)Leading-edge sweep, deg .................. 45Trailing-edge sweep, deg ................. 25Airfoil section .......................... Wedge

Body flap:Chord, m (ft) ............................ 4.57 (15) 4.57 (15)Span, m (ft) ............................. 15.93 (52.26) 15.93 (52.26)Area, m2 (ft2) 72.83 (783.9) 72.83 (783.9)eeoooo.ooooeoooo.eoe.ooooe.

Tip fin:Root chord, m (ft) 3.81 (12.5)

Tip chord' m (ft) iiiiiiiiiiiiiiiiiiiiiiii 1.90 (6.25)Span, m (ft) .....[ 3.05 (10.0)

Area, m2 (ft2) ... 8.71 (93.75)Toe-in angle, deq 2.33Cant angle, deg .. [_ 10Controller chord, m (ft) ................. 1.52 (4.99)Controller span, m (ft) .................. 3.04 (9.99)Controller area, m2 (ft2 4.64 (49.92)eooooiooooo...o.

a Values of fineness ratio are obtained by dividing the fuselage length bythe diameter of a circle whose area eguals the fuselage base area.

8

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(a) Model I.

Figure I.- Model details. Dimensions are normalized by.referencefuselage length, t.

_D

Controller-_

\

Hingeline /

(b) Tip-fin details for model 2.

Figure 1.- Concluded.

10

L-80-8163

(a) Model 1.

Figure 2.- Model-sting arrangements in test section of Langley 20-Inch Mach 6 Tunnel.

L-80-8164

(b) Model 2.

Figure 2.- Concluded.

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Figure 3.- Effects of elevon and body-flap deflections on static lonqitudinal aerodynamiccharacteristics of model I. M = 5.93; R = 2.5 × 106.1

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Figure 3.- Continued.

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0 5 10 15 20 25 30 35 40

% deg

(f) CD plotted against 5.

Figure 3.- Continued.

18

6e,deg 6F,deg

0 0 0[] 0 10<_ 10 0

10 10{h 20 0

3.0

2.5

L/D2.O

1.5

1.0

.5

-.5 _

-i.0

0 5 I0 15 20 25 30 35 40

a,deJ

(g) L/D plotted against _.

Figure 3.- Continued.

19

1.0

'"'""""""' " iiiilj e'd_IIIIIIIIIIZIIlIII 11l II

Illlilllll]lllll] I I:

_lllllllIlllillll i II "".9 _,,'_'"']llJ"l'"" :I] ]',,i:I I O[] 00 I00_,_;::::_:::1::: rj ,,"II,_,,],,,,,,,,,1, ,, : 10 0

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t II I

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0 5 lO 15 20 25 30 35 40deg

(h) Xcp/l plotted against _.

Fiqure 3.- Concluded.

2O

•2O

15

10

•05

Cm,0.7151

0

-.05

-.10

(a) Cm plotted against CN-

Figure 4.- Effects of tip-fin controller deflection on longitudinal aerodynamic characteristics

of model 2. 6e = 6F = 0°; M = 5.93; R I = 2.5 × 106 .

Fo

''''''''' ........ I

6TF, deg , iII

1.8 __;...... 0 0/0 I

[] Ol 20 i I1 t

I

0/40 I II

I.6 _ 20/20 '

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14 '"_'"'"'"' _ I I• Illlllllllllfl I 'Z

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0 5 I0 15 20 25 30 35 40

% d_J

(b) CN plotted against 5.

Fiwure 4.- Continued.

22

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0 5 10 15 20 25 30 35 40

a, deg

(c) CA and CA,B plotted against _.

Figure 4.- Continued.

23

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lllllilllll

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15 _- , I I-. , lllllllllll II I111II I111

0 5 10 15 20 25 30 35 40

_ d_

(d) Cm plotted against _.

Figure 4.- Continued.

24

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a, deg

(e) CL plotted against _.

Figure 4.- Continued.

25

I ',, 6TF,degI

o o/o[] 0120

o/40

1.8 A 20/ 20I1111111!11 I1llllilllIIl II

I:lllllllIlII' IIlllllllIl 1

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a, deg

(f) CD plotted against 5.

Figure 4.- Continued.

26

30

20

1,5

L/D

1.0

(g) L/D plotted against e.

Figure 4.- Continued.

27

(h) Xcp/l plotted against e.

Figure 4.- Concluded.

28

•002

C

0

0

Cnp

-.002

•OO5

o

-.005

0 5 I0 15 20 25 )0 35 40%deg

Figure 5.- Lateral-directional aerodynamic characteristics of models 1 and 2.

6e = 6F = 6TF = 0°; M = 5.93; R I = 2.5 x 106.

29

o

-.002

.012

.008

.004

o

o

-.01

-.02o 5 10 15

a, deg

20 25 30

0TF' deg

o 20

o 40

35 40

30

Figurefin

~ =

6.- Effects of tip-fin controller deflectionon lateral-directional control parameters ofo = of = 0°; M = 5.93; R = 2.5 x 10 6 •e \

of right wing-tipmodel 2.

1. Report No. 2, Government AccessionNo. 3. Recipient's Catalog No.NASA TM-84565

4. Title and Subtitle 5. Report Date

EFFECT OF MODIFICATIONS ON AERODYNAMIC CHARACTERISTICS January 1983

OF A SINGLE-STAGE-TO-ORBIT VEHICLE AT MACH 5.9 6. PerformingOrganizationCode506-51-I3-02

7. Author(s) 8. Performing Organization Report No.

Peter T. Bernot L-15504

10. Work Unit No.

9. Performing Organization Name and Address

NASA Langley Research Center 11 ContractorGrantNo

Hampton, VA 23665

13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Technical Memorandum

National Aeronautics and Space Administration 14. Sponsoring Agency CodeWashington, DC 20546

15. Supplementary Notes

16. Abstract

Hypersonic stability, control, and performance characteristics have been determinedon a single-stage-to-orbit vehicle based on control-configured stability concepts.The configuration (0.006-scale model) had a large body with a small 50° swept wing.

Two vertical-fin arrangements were investigated which consisted of a large center-line vertical tail and small wing-tip fins. The wing-tip fins had movable surfaces

called controllers which could be deflected outward. Longitudinal and lateral-directional characteristics were obtained over an angle-of-attack range from 0°to 40°. The effects of tip-fin controller deflection on roll- and yaw-controlcharacteristics at a sideslip angle of 0° were obtained. This investigation wasconducted in the Langley 20-Inch Mach 6 Tunnel.

17. Key Words (Suggested by Author(s)) 18. Distribution Statement

Hypersonic stability and control Unclassified - UnlimitedEntry vehicle aerodynamics

Advanced space transportation systems

Subject Category 02

19. Security Classif. (of this report} 20. Security Classif. (of this page) 21. No. of Pages 22. Price

Unclassified Unclassified 31 A03

Forsale by theNationalTechnicalInformationService,Springfield,Virginia 22161NASA-Langl ey, 1983

NationalAeronauticsand THIRD-CLASS BULK RATE Postage and Fees Paid r A _

SpaceAdministration National Aeronautics andSpace Administration

Washington, D.C. NASA-45120546

Official Business

Penalty for Private Use, $300

III lIE "_X " POSTMASTER: If Undeliverable (Section 158Postal Manual) Do Not Return