i

NASA Technical Memorandum 4726

Subsonic Aerodynamic Characteristics of a

Circular Body Earth-to-Orbit Vehicle

Roger A. Lepsch, Jr., George M. Ware, and Ian O. MacConochie

July 1996

NASA Technical Memorandum 4726

Subsonic Aerodynamic Characteristics of a

Circular Body Earth-to-Orbit Vehicle

Roger A. Lepsch, Jr. and George M. Ware

Langley Research Center • Hampton, Virginia

Ian 0. MacConochie

Lockheed Engineering & Sciences Company. Hampton, Virginia

National Aeronautics and Space Administration

Langley Research Center • Hampton, Virginia 23681-0001

July 1996
Available electronically at the following URL address: http://techreports.larc.nasa.gov/ltrs/itrs.html

Printed copies available from the following:

NASA Center for AeroSpace Information

800 Elkridge Landing Road

Linthicum Heights, MD 21090-2934

(301) 621-0390

National Technical Information Service (NTIS)

5285 Port Royal Road

Springfield, VA 22161-2171

(703) 487-4650
Summary

The circular body configuration is a generic shape

applicable to single- or multistage reusable Earth-to-orbit

transports. The principal attribute of the configuration is

its low structural weight for a given propellant loading.

The low weight results from the utilization of simple,

structurally efficient, circular cross-section tanks as part

of the body. A thick, clipped delta wing is the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional

aft-mounted center vertical fin, wingtip fins, and a nosemounted vertical fin. The test was conducted in the 7- by

10-Foot High-Speed Tunnel at the Langley Research

Center at Mach number 0.3.

The results of the investigation indicated that the

configuration was longitudinally stable about the estimated center-of-gravity position of 72 percent of body

length. The model had sufficient pitch-control authority

with elevators to produce stable trim for a wide range of

angle of attack. The maximum trimmed lift-drag (L/D)

ratio for the aft center-fin configuration was less than 5,

whereas the other configurations had values above 6.

The aft center-fin configuration was directionally

stable for all angles of attack tested. The wingtip and

nose fins were not intended to produce directional stability, and did not. The rudder-like surfaces (controllers) on

the wingtip fins and the all-moveable nose fin were

designed as active controls to produce artificial directional stability. These controls were effective in producing yawing moment. Small roiling-moment values

resulted from yaw control of the nose fin. Large adverse

rolling-moment increments resulted from wingtip-fin

controller deflection above an angle of attack of 13° .

Flow visualization indicated that this was most likely

caused by the influence of wingtip-fin controller deflection on wing flow separation.

Introduction

The National Aeronautics and Space Administration

(NASA) is investigating concepts for future space transportation systems. The studies have included single- and

multistage-to-orbit designs (refs. 1-5). Structural weight

is a critical factor in the performance and cost of these

designs. To minimize weight, circular cross-section propellant tanks were integrated into a vehicle concept as

load-carrying structure. The intent was to produce a simple configuration with a low weight-to-volume ratio.

Referred to as the circular body vehicle (CBV), this

design is a generic configuration that is an example of a

single-stage vehicle or an orbiter or booster element of a

multistage system. The vehicle has an aft-mounted

clipped delta wing and an estimated center of gravity at

72 percent of the body length. (The aft location results

from the heavy rocket engines at the base and empty propellant tanks in the forward body.) This center-of-gravity

location limits the moment arm of a conventional vertical

tail, resulting in the need for a relatively large tail size

when both stability and control are desired.

Alternate yaw-control devices were tested in an

effort to further reduce weight. The devices were wingtip

fins and an all-moving nose-mounted fin. The wingtip

fins house rudder-like surfaces (wingtip-fin controllers)

that are designed to be continually deflected for stability

augmentation. These fins can be small because they are

not required to provide natural directional stability,

thereby saving weight. (See ref. 6 for a description of

wingtip-fin controllers and their use.) The all-moving

nose-mounted fin is designed to act in a manner similar

to the wingtip fins. Sensors detect deviation from the

desired flight path and signal the nose fin to deflect to

drive the CBV back on course or prevent the vehicle

from diverging. A small fin is possible because of the

large moment arm available. The structural characteristics and hypersonic heating characteristics of the CBV

are shown in references 7 and 8, respectively. Summary

subsonic characteristics appear in reference 9, and supersonic characteristics appear in reference 10.

The present investigation was made to determine in

more detail the subsonic aerodynamic characteristics of

the CBV during unpowered entry from Earth orbit. Three

yaw control devices were tested: a conventional aftmounted center vertical fin, wingtip fins, and a nosemounted fin. The large aft center fin is the only fin configuration designed to give the CBV directional stability

(positive Cns). Separate pitch- and roll-control surfaces

were mounted on the wing trailing edge, and a body flap

extended aft of the fuselage. The test was conducted in

the Langley 7- by 10-Foot High-Speed Tunnel at Mach

number 0.3 and a Reynolds number of 8.7 x 106, based

on body length.

Symbols

The axis system used in the investigation is presented in figure 1. The data are normalized by the wing

planform area (including the body flap), the wingspan,

and the length of the wing mean aerodynamic chord. The

moment reference center was located at the proposed

vehicle center of gravity, which is at 72 percent of body

length from the nose.

b

cL

Cl

wing span, in.

drag coefficient, Drag/qSre f

lift coefficient, Lift/qSre f

rolling-moment coefficient,

Rolling moment/qSref b
Cm

Cr

L/D

l

M

q

Sref

X

Y

Z

(t

act

acn

ACt

_a.R

_e

5,,

6r

5SB

pitching-moment coefficient,

Pitching moment/ qSref_

yawing-moment coefficient,

Yawing moment/qSre f b

side-force coefficient, Side force/qSre f

wing mean aerodynamic chord, in.

lift-drag ratio

body length, in.

Mach number

free-stream dynamic pressure, lb/in 2

wing planform area (projected to body

centerline including body flap), in 2

longitudinal body axis

lateral body axis

vertical body axis

angle of attack, deg

angle of sideslip, deg

change in rolling-moment coefficient

change in yawing-moment coefficient

change in side-force coefficient

aileron-control deflection angle

(_a,L -- _a,R) 12, deg

left wing aileron deflection angle (positive

when deflected downward), deg

right wing aileron deflection angle (positive

when deflected downward), deg

elevator deflection angle (positive when

deflected downward), deg

nose-fin deflection angle (positive when

deflected with trailing edge to right), deg

rudder deflection angle (positive when

deflected with trailing edge to left), deg

speed-brake deflection angle, deg

wingtip-fin controller deflection angle positive

when deflected with trailing edge to left), deg

Stability derivatives:

CIa ACI/A _ taken at _ = 0° and 4°, per deg

C% ACn/A _ taken at I_ = 0 ° and 4 °, per deg

Cr_ ACy/A_I taken at [3 = 0 ° and 4°, per deg

Subscripts:

max maximum value

ref reference

Description of Model

A photograph of the circular body model in the

Langley 7- by 10-Foot High-Speed Tunnel is presented

in figure 2(a), and a sketch of the CBV with the three fin

arrangements tested is shown in figure 2(b). Dimensional

information is given in figure 3 and table I. The model

consisted of a spherically blunted ogive nose blended

into a large circular cross-section body with a clipped

delta wing with a 47 ° of sweep mounted on the far aft

underside. A body flap extended aft from the lower body.

The wing, which had a National Advisory Committee for

Aeronautics (NACA) 0010-64 airfoil, was equipped with

elevator surfaces on the inboard portion of the unswept

trailing edge and small aileron surfaces on the outboard

portion. Three vertical fin configurations were investigated: (1) a large conventional center fin on the upper aft

fuselage, (2) a small vertical fin near the fuselage nose,

and (3) a small fin on each wingtip. The aft center fin had

a wedge airfoil shape with a rounded leading edge and a

blunt trailing edge, as did each wingtip fin. The nose fin

had a modified NACA 0015 airfoil with a rounded trailing edge truncated at 45 percent of the airfoil chord.

The elevators could be deflected from 14 ° to -20 °.

The ailerons were tested at +5 °. The body flap was not

deflected in these tests. The deflected yaw-control surfaces on the center vertical fin were simulated by wedges

of 7.5 ° and 15 ° attached to the fin. This configuration

represents a split control surface with one half deflected,

while the other half remains undeflected. The wingtip-fin

control surfaces, referred to as controllers, are flushmounted panels designed to be deflected in an outward

direction only. Controller deflections were simulated by

angled plates of 20 ° , 40 °, and 60 ° . The controllers act as

yaw generators when a single surface is deflected outward. Deflection of the nose fin for yaw-control evaluation was accomplished by pivoting it about its 0.25-chord

station. In addition to pitch-, roll-, and yaw-control

surfaces, various speed brake controls were investigated. (See fig. 2(b).) With the aft center fin, the brake

system consisted of a flared split rudder with equal

deflection on each side of the fin. With wingtip fins, the

brakes were simultaneous outward deflections of both

wingtip-fin controllers. For the nose-fin configuration,

aft side-body-mounted panels were deflected.

Apparatus, Tests, and Corrections

Tests were conducted in the Langley 7- by 10-Foot

High-Speed Wind Tunnel, a closed-circuit subsonic

atmospheric tunnel with a nominal 7-foot-high by

10-foot-wide test section. A description of the facility

appears in reference 11. The test was conducted at

a Mach number of 0.3 and a Reynolds number of

8.7 x 106, based on body length. The model was sting
mountedthroughitsbase,andforcesandmomentswere

measuredwith aninternallymountedstrain-gagebalance.Modelanglesof attackandsideslipwerecorrected

for stingandbalancedeflectionunderloadbywind-off

calibrationof thesystemwithweights.Tunnelinterferencecorrectionscausedby changesin jet boundary,

buoyancy,flowangularity,andblockagebythepresence

of themodelwereappliedtothedata.Tofix theboundarylayerlaminar-to-turbulent-flowtransitionpoint,transition grit wasappliedon the modelaccordingto

methodsdescribedin reference12.Number100carborundumgritwasthinlysprinkledin 1/16-in.bands1in.

aftof thenose.Number80grit of thesamewidthwas

applied1in. downstreamof thewingandfin leading

edges.Themodeltestpitchrangewasnominallyfrom

-2° to 16°.Themodelwastestedatanglesof sideslipof

0° and+4 ° over the angle-of-attack range. Data were

taken as the model was moved from negative to positive

angles. Drag values are presented as measured with no

base corrections applied.

To aid in the analysis of the force and moment data,

the flow over the model was defined by an oil-flow technique. SAE 40 oil containing a pigment that fluoresces

when exposed to ultraviolet light was brushed over the

model. The model was positioned in the tunnel at the

desired angle of attack. The tunnel was brought up to

speed, and after the oil had sufficient time to flow, the

tunnel was shut down. The model was illuminated with

an ultraviolet light and photographs were taken. The photographs were then used to visualize flow characteristics.

Results and Discussion

Longitudinal Characteristics

Baseline characteristics. Lift, drag, and pitchingmoment coefficients and L/D are plotted against angle of

attack in figure 4 for the model with fins off and with

each of the vertical fin arrangements. The lift and pitch

data for all configurations are linear up to about _ = 12 °.

Above ct = 12 °, a decrease in the lift curve slope indicates the onset of flow separation. The highest lift values

for a given angle of attack were obtained with the

wingtip-fin configuration. The wingtip fins acted as end

plates to increase wing effectiveness. (See ref. 13.) The

added wing lift also is seen as a slight increase in longitudinal stability over that of the other configurations. The

model with each vertical fin combination or with fins off

was longitudinally stable about the 72 percent of body

length reference. The very large aft center fin (68 percent

of the exposed area of a single wing panel) produced

more drag than the other fin arrangements. As a result,

the model with the center fin had lower L/D values and a

positive pitching-moment increment caused by the drag

of the fin acting above the model center of gravity. The

maximum untrimmed L/D ratio for the aft center-fin configuration was about 5, whereas the other configurations

had values of about 7.

Oil-flow photographs of the upper surface of the

right wing without and with wingtip fins are presented in

figures 5 and 6, respectively. The photographs show the

onset of flow separation at angles of attack from 10°

to 13°. Separation results from outflow toward the wingtip reversing the flow on the wing upper surface near the

tip (fig. 5(c)). Separation for the model with wingtip fins

on appears to be very similar to the fins-off case.

Pitch control characteristics. Data for elevator

deflections of 0% -14 °, and -20 ° are presented in figure 7. For each fin arrangement, the negatively deflected

surfaces produce moments that trim the model to higher

angles of attack. Stable trim is possible for all fin

arrangements up to about 12 ° or 14 ° angle of attack. For

the remainder of the test angle-of-attack range, the CBV

is neutrally stable or unstable, with the exception of the

wingtip-fin configuration, which may have a stable secondary trim point at a higher angle of attack. The control

power of the elevators is sufficient to allow stable trim

over a wide angle-of-attack range regardless of fin configuration. As pointed out in the previous section, use of

the center fin results in a positive increment in pitching

moment relative to the other two configurations. The

trim angle for the center-fin configuration with neutral

elevators is about ct = 2% whereas the other configurations trim at ot = 0° or-1 °. The estimated trimmed

(L/D)max is less than 5 for the center-fin model and

somewhat greater than 6 for the model with the other two

fin arrangements.

By comparing the flow pattern over the wing with

(fig. 8) and without (fig. 5) negative elevator deflection,

one sees that the upward deflected elevator delays the

onset of flow separation. (Compare figs. 5(c) and 8(c).)

A decrease in flow circulation is consistent with the pressure rise on the wing upper surface that would result

from an upward elevator deflection.

Speed-brake effects. Three different speed-brake

systems were tested on the CBV model, and data from

those tests appear in figure 9. Speed brakes are used by a

gliding unpowered spacecraft as an energy management

device to adjust cross range and target the landing site.

For the model with the aft center-fin arrangement, the

split rudder acted as a speed brake. Data are presented

with the brake open 7.5 ° and 15 ° on either side from the

closed position (fig. 9(a)). The brake effectively

increased drag, but lift decreased somewhat because of

downforce generated by the fin. The downforce is the

vertical component of the pressure force acting on the
flared split-rudder surfaces, which have a significant projected area in the horizontal plane. The increase in drag

above the model center of gravity again produced a large

nose-up pitching moment. As a result, the trim angle of

attack shifted from about 2 ° with speed brakes undeflected to 9° with speed brakes set at 15 °. If speed brakes

are used on the CBV in this fashion, a compensating elevator deflection will have to accompany brake

deflection.

The effect of wingtip-fin-mounted speed brakes is

presented in figure 9(b). Because the surfaces were relatively small, deflections up to 60 ° were tested. The

brakes extended out from the wingtips and were effective

in producing drag increments linear with speed-brake

deflection. The line of action of the drag increment from

the brakes was close to the estimated center of gravity, so

little change in pitching moment resulted. Their use also

delayed the onset of flow separation, as can be seen by

examining the lift and moment curves in the 12° to 17 °

angle-of-attack range.

The speed brakes for the nose-fin model were

mounted on the sides of the body over the wing.

Figure 9(c) shows that deflecting the side-body brakes

produced drag increments linear with deflection. However, the brakes were largely ineffective in the upper portion of the angle-of-attack range. Also, lift was reduced

as the brakes were deflected. Deflection of the brakes in

such close proximity to the wing upper surface disturbed

the flow over the wing, resulting in a small loss of lift.

This loss of lift caused the pitching-moment curves to

shift slightly upward with speed-brake deflection. These

effects are quite different from those found in tests at

M = 2.0 to 4.6 (ref. 9). In the supersonic speed range,

deflection of the side-body brakes increased drag only

slightly and introduced a more significant nose-up pitching moment.

Lateral Characteristics

Lateral-directional stability. The lateral-directional

characteristics of the CBV appear in figure 10 in the form

of the stability parameters Cy_, Cn_, and Clo plotted

against angle of attack. Data are shown for t_e model

with all fin configurations and with fins off for an elevator setting of-14 °. This elevator setting trims the model

at an angle of attack that approximates a landing

approach condition.

The CBV with aft center fin was directionally stable,

with the value of Cnl_ increasing slightly over the test

angle-of-attack range. The model with the other fin

arrangements was, as expected, directionally unstable.

The nose fin was directionally destabilizing, whereas

wingtip fins generally added Cnf_ increments. Little dif4

ference occurred in the effective dihedral parameter

-Ct_ between the fins off and nose-fin configurations. p

The center-fin model had negative values of Ct_ over

the test angle-of-attack range.

Yaw-controleffects. Rudder effectiveness for each

of the vertical-fin arrangements is presented in figure 11.

Yaw control for the aft center-fin configuration

(fig. 1 l(a)) was accomplished by deflecting one half of

the split rudder outboard and leaving the other half undeflected. The rudder produced almost constant values of

yawing moment across the test angle-of-attack range.

The retention of effectiveness at high angles of attack

probably was caused by the large size of the center fin,

which placed it above the influence of the fuselage and

wing. As expected, the size and placement of the fin

above the vehicle center of gravity also caused large

adverse rolling moments with rudder deflection. The

value of the rolling moment, which was about one-third

that of the yawing moment at the trim angle of attack of

14°, increased significantly with decreased angle of

attack.

The effect of deflecting a wingtip-fin controller is

shown in figure 11 (b). The data indicate that the yawing

moments produced by the controller decreased only

slightly with increasing angle of attack up to about 13°.

At higher angles, a large increase in adverse rollingmoment increment occurred. A clue to the adverse rolling moment is seen in figure 12 when one compares the

flow pattern over the wing with the wingtip-fin controller

deflected with the flow pattern with the controller undeflected. The deflected controller reduced the outflow on

the wing thereby delaying flow separation near the

wingtip. With one controller deflected, an adverse rolling

moment resulted from the asymmetric flow conditions

over the wings.

The nose fin, placed far forward to take advantage of

the long moment arm created by the 72 percent of body

length center-of-gravity location, was very effective in

generating yawing moment (fig. l l(c)). The effectiveness of the deflected nose fin increased almost linearly

with increasing angle of attack. This increased effectiveness is probably due to the increase in cross flow around

the circular body as the pitch angle increased. These

yawing moments were accompanied by only small rolling moments up to the wing flow-separation angle.

Roll-control effects. Roll control results from differentially deflecting the dedicated aileron-control surfaces

on the outer wing trailing edge. The effectiveness values

are presented in figure 13 for surfaces set at 5 ° on the left

and -5 ° on the right. Tests were made with wingtip fins

on and off. The wingtip-fin controllers were undeflected.
The dataindicatethat aileronroll effectiveness

increasedlinearlywithangleofattackuptoangleswhere

flowseparationoccurredonthewingtip.Smalladverse

yawingmomentsaccompaniedroll control.Thecontrol

authorityof the aileronsappearsto he sufficientto

countertheadverserollingmomentsnotedforthecenterfin andwingtip-finarrangements nearthetrimanglesof

attackdesiredforlanding.Insufficientdataexisttoindicatewhethertheaileronswouldhavesufficientcontrol

effectivenesstocounteradverserollcausedbythecenter

finatreducedanglesofattack.

Concluding Remarks

Tests of a circular body spacecraft model have been

conducted in the Langley 7- by 10-Foot High-Speed

Wind Tunnel. The design is an option considered for

single- or multistage-to-orbit vehicles. The model had a

body with a circular cross section and a clipped delta

wing. Three vertical fin arrangements were investigated:

a conventional aft-mounted center vertical fin, wingtip

fins, and a nose-mounted vertical fin.

Results of the investigation indicated that the configuration was longitudinally stable about the estimated

center-of-gravity position of 72 percent of body length.

The model had sufficient pitch-control authority with

elevators to produce stable trim over a wide range of

angle of attack regardless of fin arrangement. The maximum trimmed lift-drag ratio for the aft center-fin configuration was less than 5, whereas the other configurations

had values above 6.

The aft center-fin configuration was directionally

stable for all angles of attack. The wingtip and nose fins

were not intended to produce directional stability, and

did not. The rudder-like surfaces on the wingtip fins and

the nose fin were designed as active controls to produce

artificial directional stability. These controls were effective in producing yawing moment throughout the test

angle-of-attack range. As expected, deflection of the rudder on the aft center fin produced large adverse rolling

moments. Small rolling-moment values resulted from

yaw control of the nose fin, whereas large, adverse

rolling-moment increments resulted from wingtip-fin

controller deflection above an angle of attack of 13° .

Differential deflections of the aileron surfaces on the

outboard wing trailing edge were effective in producing

rolling moments with wingtip fins on or off. Aileron

effectiveness was sufficient to counter the adverse rolling

moments noted for the aft center fin and wingtip fin

configurations in the angle-of-attack range desired for

landing.

NASA Langley Research Center

Hampton, VA 23681-0001

April 18, 1996

References

1. Freeman, D. C., Jr.: The New Space Transportation Begins

Today. Astronaut. & Aeronaut., vol. 21, June 1983, pp. 37, 48.

2. Martin, James A.: Orbit on Demand: In this Century If Pushed.

Aerosp. America, vol. 23, no. 2, Feb. 1985, pp. 46-48.

3. Talay, T. A.: Shuttle 11. SAE Tech. Paper Ser. 871335, June

1987.

4. Holloway, Paul E; and Talay, Theodore A.: Space Transportation Systems--Beyond 2000. IAF Paper 87-188, Oct. 1987.

5. Talay, Theodore A.; and Morris, W. Douglas: Advanced

Manned Launch Systems. Paper presented at AAAF, DGLR,

Royal Aeronautical Society, and ESA Second European Aerospace Conference on Progress in Space Transportation (Bonn,

Federal Republic of Germany), May 22-24, 1989.

6. Powell, Richard W.; and Freeman, Delma C., Jr.: Application

of a Tip-Fin Controller to the Shuttle Orbiter for Improved

Yaw Control. AIAA-81-0074, Jan. 1981.

7. MacConochie, I. O.; and Klich, P. J.: Technologies Involved in

Configuring an Advanced Earth-to-Orbit Transport for Low

Structural Mass. SAWE Paper 1380, May 1980.

8. Wells, William L.; MacConochie, lan O.; Helms, Vernon T.;

and Raney, David: Heating Rate Distributions at Mach 10 on

a Circular Body Earth-to-Orbit Transport Vehicle. A1AA85-0974, June 1985.

9. Lepsch, R. A., Jr.; and MacConochie, I. O.: Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-Orbit Transport. AIAA-86-1801, June 1986.

10. Ware, George M.; Engelund, Walter C.; and MacConochie, Ian

O.: Supersonic Aerodynamic Characteristics of a Circular

Body Earth-to-Orbit Vehicle. NASA TM-4533, 1994.

11. Pefiaranda, Frank E.; and Freda, M. Shannon, eds.: Aeronautical Facilities Catalogue. Volume l--Wind Tunnels. NASA

RP-1132, 1985.

12. Braslow, Albert L., Hicks, Raymond, M.; and Harris, Roy V.,

Jr.: Use of Grit-Type Boundary-Layer-Transition Trips. Conference on Aircraft Aerodynamics, NASA SP-124, 1966,

pp. 19-36.

13. Montoya, Lawrence C.; Jacobs, Peter F.; and Flechner, Stuart

G.: Effect of Winglets on a First-Generation Jet Transport

Wing. I11 Pressure and Spanwise Load Distributions for a

Semispan Model at Mach 0.30. NASA TN D-8478, 1977.
Table I. Geometric Characteristics of Circular Body Model

Body:

Length, in ................................................................................. 52.00

Base area, in 2 .................. .66.96

Wing:

Airfoil .......................................................................... NACA 0010-64

Mean aerodynamic chord (reference length), in ................................................... 16.00

Span (reference span), in ..................................................................... 35.04

Area to body centerline, reference area, in 2 ..................................................... 486.60

Area, exposed outside of body, in 2 ............................................................ 305.08

Center vertical fin:

Airfoil ............................................................................ Double wedge

Area, in 2 ................................................................................. 100.00

Tip fins (each):

Airfoil .......................................................................... Modified wedge

Area, in 2 .................................................................................. 11.60

Nose fin:

Airfoil ..................................................................... Modified NACA 0015

Area, in 2 ................................................................................... 7.36

Control surfaces (each):

Elevator area, in 2 ........................................................................... 25.56

Body flap area, in 2 .......................................................................... 24.96

Aileron area, in 2............................................................................. 5.84

Wingtip-fin controller and speed brake area, in 2 .................................................... 5.08

Aft center-fin rudder and speed brake area, in 2 .................................................... 25.80

B • 2 ody speed brake area, m .................................................................... 9.72
Sideforce

Y

X

Lift

Rolling

X moment

Pitching

_ __"_ moment

Wi!ddirection _ _____-'- Drag

Yawing _-7_'_

moment7_

Z

Figure1.Axissystemusedininvestigationwithpositivedirectionsofforcesandmoments.

7
(a) ModelinLangley7-by 10-FootHigh-SpeedTunnel.

Figure2. Circularbodyearth-to-orbit vehiclemodel.

L-84-10,263

$
Center-finconfiguration

(withrudderand

speed

Speedbrake

(usedwithnosefinconfiguration)

Elevator

Aileron

Nose-fin

configuration

Wingtip-finconfiguration

(withcontrollersand

speedbrakes)

(b) Modelshowingthreefin arrangements investigated.

Figure2. Concluded.
37.10rad.

[_ 23.28

4.49

i.45

2.88

r Body

flap

t

8.80

4.00

;i--c

-5 7 °

--f-

- 8.66 35.04

37.44

52.00 _ 1.56

-_ _2.88

Sre f = 486.60 in. 2

= 16.00 in.

b = 35.04 in.

1= 52.00 in.

(a) Wing body.

Figure 3. Model used in investigation. All linear dimensions are in inches.

10
1.5° toe-in"

10°

Wingtip-fincontroller

andspeedbrake

.36

t

(

1.52

\

\

.48 _ '_

3.!6 3.92

L i

i

i

4.00 _'- Controller and speed brake

(b) Wingtip fin.

Figure 3. Continued.

11
,11 63.06 -I

43.84 Split rudder

and speed brake

•156 rad.

10 °

1.22

18.30

_-1!o

75o-.q 13.00

(c) Aft center fin.

Figure 3. Continued.

12
41 52.00

Nose fin

__.1 _9_

.05 rad. -__

__...4{ "22 rad"

4 ..--1, 35o /_/-- Piv°t axis

 ,9o .q

]_ 3.86

1.34

Hinge line

_--- .86

.I 16 rad. ]__

,_ 372 _ -_!

1.40 rad.

(d) Nose fin.

Figure 3. Concluded.

5.34

Speed brake

1.86

13
CL

.8

.7

.6

.5

.4

.3

.2

.!

0

-.2

/j"

L/D

8

7

6

5

4

3

2

1

0

-1

-2

/j

/

/

//" _'_

_/ _

// /- _

3

S

o

[]

<>

/x

Configuration

Fins off

Center fin

Wingtip fins

Nose fin

CD

.40 .12

.36 .08

.32

.28

.24

.20

.16

.12

.08

.04

.04

0 '_k. Lk"_ '7

-.04 \ _"_ 6%

\ _

Cm -.08 _\ '_N _2. c

-.12 "'Q "_

•

X2"- _

/_ -.16

-.20

-.24

0

-8

-.28

-4 0 4 8 12 16 20 24 -8 -4 0 4 8 12 16 20 24

_, deg cx, deg

Figure 4. Longitudinal characteristics of circular body model with various fin arrangements, fie = 0%

14
(a) _= 6.35 °. (b) or= 10.64 °.

(c) _ = 12.80 °. (d) c_ = 14.83 °.

Figure 5. Oil flow on fins-off configuration at M = 0.3, showing upper surface of right-side wing with 5e = 0°.

15
(a) _ =6.36°. (b) ot = 10.72 °.

(c) _t = 12.86 °. (d) a = 14.90 °.

Figure 6. Oil flow on tip-fin configuration at M = 0.3, showing right-wing upper surface of right-side wing with _e = 0°"

16
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i

-.16

-<3

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-4 0 4 8 12 16 20 24 -8 -4 0 4 8 12 16

ct, deg or, deg

(a) Center fin.

Figure 7. Effect of elevator deflection on longitudinal characteristics of circular body models.

20 24

17
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1.6

1.4

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16 20 24 -8

(b) Wingtip fins.

Figure 7. Continued.

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(c) Nose fin.

Figure 7. Concluded.

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or, deg

19
(a) o_= 6.35 °. (b) o_= 10.64 °.

2O

Figure 8.

(c) ot = 12.80 °. (d) ct = 14.83 °.

Oil flow on fins-off configuration at M = 0.3, showing upper surface of left-side wing with _e = --14 °.
CL

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0_, deg _, deg

(a) Center fin.

Figure 9. Effect of speed-brake deflection on longitudinal characteristics of circular body model. _e = 0°'

24

21
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(b) Wingtipfins.

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Figure 9. Continued.

h

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tx, deg

22
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(c) Nose fin.

Figure 9. Concluded.

F

I

i

E

i

i ° ° -_

q ....

i

I

J

-4 0 4 8 12 16 20 24

O., deg

23
.10 .010

.08 .OO8

.06 .006

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0

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0

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ix_ A__A

1

Stable -

12 16 20 24

04 deg

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3

_-----t_-t---_,2a 3__ 3.._ 3._ Up" u-_ q3.-

-.004

Stable

Configuration

o Fins off

[] Center fin

© Wingtip fins

A Nose fin

-.006

-.008

-.010

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_, deg

Figure 10. Lateral directional-stability characteristics of circular body model with various fin arrangements. _e = 0°"

24
ACy

.05

.04

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0

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_ / IS1--4--_ -[]--

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.012

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a.'a.

-0 \

¢1,deg

_r' deg

o 7.5

[] 15

-.002

-.004

-.006

-8 -4 0 4 8 12 16 20 24

or, deg

(a) Center fin.

Figure 11. Effect of rudder deflection as yaw control for circular body model. _e = -14°"

25
ACy

.05

.04

.03

.02

.01

0

-.01

-.02

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_--- 3---_ 3--- )-----(3-- o"_°_

.002

0

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Ck__ k _-,,.,.---)--'-_-. Ox

E _ -_ t-- .r--_&\ -1-_C_-_.J

Cf "_5 /

\

[]

\

-.04

-.05

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-.016

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-8 -4 0 4 8

Ct, deg

12 16 20 24

.012

.010

.008

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0

"---3--0" "

/:

_TF' deg

o 20

[] 40

<_ 60

-.002

-.004

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Ct, deg

12 16 20 24

(b) Wingtip fins.

Figure 11. Continued.

26
.02 .004

,01

0

-.01

-.02

ACy -.03

-.04

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\\

\

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5.

v

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AC n

0

-.004

-.008

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-.016

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-.024

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-8 -4 0 4 8 12 16 20 2_

.006

a, deg

.004

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0

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C_

r-

©

8n, deg

o 5

[] 10

O 15

-.010

-.012

-.014

-8 -4 0 4 8

ct, deg

12 16 20 24

(c) Nose fin.

Figure 11. Concluded.

27
(a) 8rF = 0°.

(b) _TF = 20°.

Figure 12. Oil flow on tip-fin configuration at M = 0.3, showing right-wing upper surface, o_= 12.86 °, _e = 0°-

28
ACy

AC l

.05

.04

.03

.02

.01

0

-.01

-.02

-.03

-.04

-.05

.010

.008

.006

.004

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0

-.002

-.004

-.006

-.008

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-8

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.OO4

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AC n 0

-.002

,,7,,,

- 3"

-.004

-.006

-.008

-.010

-8 -4 4 8

R, deg

12 16 20 24

No

[] Configuration

o Fins off

[] Wingtip fins

-4 0 4 8 12 16 20 24

a, deg

Figure 13. Effect of outboard ailerons as roll control for circular body model. 8 a = 5 ° and _e = 0°"

29
I

REPORT DOCUMENTATION PAGE I ro,_o_o_

I

OMB No. 0704-0188

Public reporting burden for this ¢oliection of _[,l_mat_on is estimated to average 1 hour per response, including the time for reviewing instru_k:ms, searching existing data sources

gathering and maintaining the data nee(led, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of thi_

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jeffersor

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budgat, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

July 1996 Technical Memorandum

4. TITLE AND SUBIIULE J5. FUNDING NUMBERS

Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-Orbit

Vehicle WU 242-20-08-01

8. AUTHOR(S)

Roger A. Lepsch, Jr., George M. Ware, and Ian O. MacConochie

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) a. PERFORMING ORGANIZATION

REPORT NUMBER

NASA Langley Research Center

Hampton, VA 23681-0001 L- 17429

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

National Aeronautics and Space Administration NASA TM-4726

Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES

Lepsch and Ware: Langley Research Center, Hampton, VA; MacConochie: Lockheed Engineering & Sciences

Company, Hampton, VA.

1211. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Categories 15 and 02

Availability: NASA CASI (301) 621-0390

12b. DISTRIBUTION CODE

13. AiSTRACT (Maximum 200 words)

A test of a generic reusable earth-to-orbit transport was conducted in the 7- by 10-Foot high-speed tunnel at

the Langley Research Center at Mach number 0.3. The model had a body with a circular cross section and a

thick clipped delta wing as the major lifting surface. For directional control, three different vertical fin

arrangements were investigated: a conventional aft-mounted center vertical fin, wingtip fins, and a nosemounted vertical fin. The configuration was longitudinally stable about the estimated center-of-gravity position of 0.72 body length and had sufficient pitch-control authority for stable trim over a wide range of angle

of attack, regardless of fin arrangement. The maximum trimmed lift/drag ratio for the aft center-fin configuration was less than 5, whereas the other configurations had values of above 6. The aft center-fin configuration

was directionally stable for all angles of attack tested. The wingtip and nose fins were not intended to produce

directional stability but to be active controllers for artificial stabilization. Small rolling-moment values

resulted from yaw control of the nose fin. Large adverse rolling-moment increments resulted from tip-fin controller deflection above 13 ° angle of attack. Flow visualization indicated that the adverse rolling-moment

increments were probably caused by the influence of the deflected tip-fin controller on wing flow separation.

14.SUBJECTTERMS

Aerodynamics; Subsonic; Spacecraft; Booster

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

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Unclassified

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30

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A03

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