National Aeronautics and Space AdministrationLangley Research Center • Hampton, Virginia 23681-0001

NASA Technical Memorandum 4726

Subsonic Aerodynamic Characteristics of aCircular Body Earth-to-Orbit VehicleRoger A. Lepsch, Jr. and George M. WareLangley Research Center • Hampton, Virginia

Ian O. MacConochieLockheed Engineering & Sciences Company • Hampton, Virginia

July 1996

Printed copies available from the following:

NASA Center for AeroSpace Information National Technical Information Service (NTIS)800 Elkridge Landing Road 5285 Port Royal RoadLinthicum Heights, MD 21090-2934 Springfield, VA 22161-2171(301) 621-0390 (703) 487-4650

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

Summary

The circular body configuration is a generic shapeapplicable to single- or multistage reusable Earth-to-orbittransports. The principal attribute of the configuration isits low structural weight for a given propellant loading.The low weight results from the utilization of simple,structurally efficient, circular cross-section tanks as partof the body. A thick, clipped delta wing is the major lift-ing surface. For directional control, three different verti-cal fin arrangements were investigated: a conventionalaft-mounted center vertical fin, wingtip fins, and a nose-mounted vertical fin. The test was conducted in the 7- by10-Foot High-Speed Tunnel at the Langley ResearchCenter at Mach number 0.3.

The results of the investigation indicated that theconfiguration was longitudinally stable about the esti-mated center-of-gravity position of 72 percent of bodylength. The model had sufficient pitch-control authoritywith elevators to produce stable trim for a wide range ofangle 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 directionallystable for all angles of attack tested. The wingtip andnose fins were not intended to produce directional stabil-ity, and did not. The rudder-like surfaces (controllers) onthe wingtip fins and the all-moveable nose fin weredesigned as active controls to produce artificial direc-tional stability. These controls were effective in pro-ducing yawing moment. Small rolling-moment valuesresulted from yaw control of the nose fin. Large adverserolling-moment increments resulted from wingtip-fincontroller deflection above an angle of attack of 13°.Flow visualization indicated that this was most likelycaused by the influence of wingtip-fin controller deflec-tion on wing flow separation.

Introduction

The National Aeronautics and Space Administration(NASA) is investigating concepts for future space trans-portation systems. The studies have included single- andmultistage-to-orbit designs (refs. 1–5). Structural weightis a critical factor in the performance and cost of thesedesigns. To minimize weight, circular cross-section pro-pellant tanks were integrated into a vehicle concept asload-carrying structure. The intent was to produce a sim-ple configuration with a low weight-to-volume ratio.Referred to as the circular body vehicle (CBV), thisdesign is a generic configuration that is an example of asingle-stage vehicle or an orbiter or booster element of amultistage system. The vehicle has an aft-mountedclipped delta wing and an estimated center of gravity at

72 percent of the body length. (The aft location resultsfrom the heavy rocket engines at the base and empty pro-pellant tanks in the forward body.) This center-of-gravitylocation limits the moment arm of a conventional verticaltail, resulting in the need for a relatively large tail sizewhen both stability and control are desired.

Alternate yaw-control devices were tested in aneffort to further reduce weight. The devices were wingtipfins and an all-moving nose-mounted fin. The wingtipfins house rudder-like surfaces (wingtip-fin controllers)that are designed to be continually deflected for stabilityaugmentation. These fins can be small because they arenot required to provide natural directional stability,thereby saving weight. (See ref. 6 for a description ofwingtip-fin controllers and their use.) The all-movingnose-mounted fin is designed to act in a manner similarto the wingtip fins. Sensors detect deviation from thedesired flight path and signal the nose fin to deflect todrive the CBV back on course or prevent the vehiclefrom diverging. A small fin is possible because of thelarge moment arm available. The structural characteris-tics and hypersonic heating characteristics of the CBVare shown in references 7 and 8, respectively. Summarysubsonic characteristics appear in reference 9, and super-sonic characteristics appear in reference 10.

The present investigation was made to determine inmore detail the subsonic aerodynamic characteristics ofthe CBV during unpowered entry from Earth orbit. Threeyaw control devices were tested: a conventional aft-mounted center vertical fin, wingtip fins, and a nose-mounted fin. The large aft center fin is the only fin con-figuration designed to give the CBV directional stability(positive ). Separate pitch- and roll-control surfaceswere mounted on the wing trailing edge, and a body flapextended aft of the fuselage. The test was conducted inthe Langley 7- by 10-Foot High-Speed Tunnel at Machnumber 0.3 and a Reynolds number of 8.7× 106, basedon body length.

Symbols

The axis system used in the investigation is pre-sented in figure 1. The data are normalized by the wingplanform area (including the body flap), the wingspan,and the length of the wing mean aerodynamic chord. Themoment reference center was located at the proposedvehicle center of gravity, which is at 72 percent of bodylength from the nose.

b wing span, in.

CD drag coefficient, Drag/qSref

CL lift coefficient, Lift/qSref

Cl rolling-moment coefficient,Rolling moment/qSrefb

Cnβ

2

Cm pitching-moment coefficient,

Cn yawing-moment coefficient,Yawing moment/qSrefb

CY side-force coefficient, Side force/qSref

wing mean aerodynamic chord, in.

L/D lift-drag ratio

l body length, in.

M Mach number

q free-stream dynamic pressure, lb/in2

Sref wing planform area (projected to bodycenterline including body flap), in2

X longitudinal body axis

Y lateral body axis

Z vertical body axis

α angle of attack, deg

β angle of sideslip, deg

∆Cl change in rolling-moment coefficient

∆Cn change in yawing-moment coefficient

∆CY change in side-force coefficient

δa aileron-control deflection angle(δa,L − δa,R)/2, deg

δa,L left wing aileron deflection angle (positivewhen deflected downward), deg

δa,R right wing aileron deflection angle (positivewhen deflected downward), deg

δe elevator deflection angle (positive whendeflected downward), deg

δn nose-fin deflection angle (positive whendeflected with trailing edge to right), deg

δr rudder deflection angle (positive whendeflected with trailing edge to left), deg

δSB speed-brake deflection angle, deg

δTF wingtip-fin controller deflection angle positivewhen deflected with trailing edge to left), deg

Stability derivatives:

∆Cl /∆β taken atβ = 0° and 4°, per deg

∆Cn/∆β taken atβ = 0° and 4°, per deg

∆CY /∆β taken atβ = 0° and 4°, per deg

Subscripts:

max maximum value

ref reference

Description of Model

A photograph of the circular body model in theLangley 7- by 10-Foot High-Speed Tunnel is presentedin figure 2(a), and a sketch of the CBV with the three finarrangements tested is shown in figure 2(b). Dimensionalinformation is given in figure 3 and table 1. The modelconsisted of a spherically blunted ogive nose blendedinto a large circular cross-section body with a clippeddelta wing with a 47° of sweep mounted on the far aftunderside. A body flap extended aft from the lower body.The wing, which had a National Advisory Committee forAeronautics (NACA) 0010-64 airfoil, was equipped withelevator surfaces on the inboard portion of the unswepttrailing edge and small aileron surfaces on the outboardportion. Three vertical fin configurations were investi-gated: (1) a large conventional center fin on the upper aftfuselage, (2) a small vertical fin near the fuselage nose,and (3) a small fin on each wingtip. The aft center fin hada wedge airfoil shape with a rounded leading edge and ablunt trailing edge, as did each wingtip fin. The nose finhad a modified NACA 0015 airfoil with a rounded trail-ing 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 notdeflected in these tests. The deflected yaw-control sur-faces on the center vertical fin were simulated by wedgesof 7.5° and 15° attached to the fin. This configurationrepresents a split control surface with one half deflected,while the other half remains undeflected. The wingtip-fincontrol surfaces, referred to as controllers, are flush-mounted panels designed to be deflected in an outwarddirection only. Controller deflections were simulated byangled plates of 20°, 40°, and 60°. The controllers act asyaw generators when a single surface is deflected out-ward. Deflection of the nose fin for yaw-control evalua-tion was accomplished by pivoting it about its 0.25-chordstation. In addition to pitch-, roll-, and yaw-controlsurfaces, various speed brake controls were investi-gated. (See fig. 2(b).) With the aft center fin, the brakesystem consisted of a flared split rudder with equaldeflection on each side of the fin. With wingtip fins, thebrakes were simultaneous outward deflections of bothwingtip-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-FootHigh-Speed Wind Tunnel, a closed-circuit subsonicatmospheric tunnel with a nominal 7-foot-high by10-foot-wide test section. A description of the facilityappears in reference 11. The test was conducted ata Mach number of 0.3 and a Reynolds number of8.7 x 106, based on body length. The model was sting

Pitching momentqSrefc⁄

c

Clβ

Cnβ

CYβ

3

mounted through its base, and forces and moments weremeasured with an internally mounted strain-gage bal-ance. Model angles of attack and sideslip were correctedfor sting and balance deflection under load by wind-offcalibration of the system with weights. Tunnel interfer-ence corrections caused by changes in jet boundary,buoyancy, flow angularity, and blockage by the presenceof the model were applied to the data. To fix the bound-ary layer laminar-to-turbulent-flow transition point, tran-sition grit was applied on the model according tomethods described in reference 12. Number 100 carbo-rundum grit was thinly sprinkled in 1/16-in. bands 1 in.aft of the nose. Number 80 grit of the same width wasapplied 1 in. downstream of the wing and fin leadingedges. The model test pitch range was nominally from−2° to 16°. The model was tested at angles of sideslip of0° and ±4° over the angle-of-attack range. Data weretaken as the model was moved from negative to positiveangles. Drag values are presented as measured with nobase corrections applied.

To aid in the analysis of the force and moment data,the flow over the model was defined by an oil-flow tech-nique. SAE 40 oil containing a pigment that fluoresceswhen exposed to ultraviolet light was brushed over themodel. The model was positioned in the tunnel at thedesired angle of attack. The tunnel was brought up tospeed, and after the oil had sufficient time to flow, thetunnel was shut down. The model was illuminated withan ultraviolet light and photographs were taken. The pho-tographs were then used to visualize flow characteristics.

Results and Discussion

Longitudinal Characteristics

Baseline characteristics.Lift, drag, and pitching-moment coefficients andL/D are plotted against angle ofattack in figure 4 for the model with fins off and witheach of the vertical fin arrangements. The lift and pitchdata for all configurations are linear up to aboutα = 12°.Above α = 12°, a decrease in the lift curve slope indi-cates the onset of flow separation. The highest lift valuesfor a given angle of attack were obtained with thewingtip-fin configuration. The wingtip fins acted as endplates to increase wing effectiveness. (See ref. 13.) Theadded wing lift also is seen as a slight increase in longitu-dinal stability over that of the other configurations. Themodel with each vertical fin combination or with fins offwas longitudinally stable about the 72 percent of bodylength reference. The very large aft center fin (68 percentof the exposed area of a single wing panel) producedmore drag than the other fin arrangements. As a result,the model with the center fin had lowerL/D values and apositive pitching-moment increment caused by the drag

of the fin acting above the model center of gravity. Themaximum untrimmedL/D ratio for the aft center-fin con-figuration was about 5, whereas the other configurationshad values of about 7.

Oil-flow photographs of the upper surface of theright wing without and with wingtip fins are presented infigures 5 and 6, respectively. The photographs show theonset of flow separation at angles of attack from 10°to 13°. Separation results from outflow toward the wing-tip reversing the flow on the wing upper surface near thetip (fig. 5(c)). Separation for the model with wingtip finson appears to be very similar to the fins-off case.

Pitch control characteristics.Data for elevatordeflections of 0°, −14°, and−20° are presented in fig-ure 7. For each fin arrangement, the negatively deflectedsurfaces produce moments that trim the model to higherangles of attack. Stable trim is possible for all finarrangements up to about 12° or 14° angle of attack. Forthe remainder of the test angle-of-attack range, the CBVis neutrally stable or unstable, with the exception of thewingtip-fin configuration, which may have a stable sec-ondary trim point at a higher angle of attack. The controlpower of the elevators is sufficient to allow stable trimover a wide angle-of-attack range regardless of fin con-figuration. As pointed out in the previous section, use ofthe center fin results in a positive increment in pitchingmoment relative to the other two configurations. Thetrim angle for the center-fin configuration with neutralelevators is aboutα = 2°, whereas the other configura-tions trim at α = 0° or −1°. The estimated trimmed(L/D)max is less than 5 for the center-fin model andsomewhat greater than 6 for the model with the other twofin 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 theonset of flow separation. (Compare figs. 5(c) and 8(c).)A decrease in flow circulation is consistent with the pres-sure rise on the wing upper surface that would resultfrom an upward elevator deflection.

Speed-brake effects.Three different speed-brakesystems were tested on the CBV model, and data fromthose tests appear in figure 9. Speed brakes are used by agliding unpowered spacecraft as an energy managementdevice to adjust cross range and target the landing site.For the model with the aft center-fin arrangement, thesplit rudder acted as a speed brake. Data are presentedwith the brake open 7.5° and 15° on either side from theclosed position (fig. 9(a)). The brake effectivelyincreased drag, but lift decreased somewhat because ofdownforce generated by the fin. The downforce is thevertical component of the pressure force acting on the

4

flared split-rudder surfaces, which have a significant pro-jected area in the horizontal plane. The increase in dragabove the model center of gravity again produced a largenose-up pitching moment. As a result, the trim angle ofattack shifted from about 2° with speed brakes unde-flected to 9° with speed brakes set at 15°. If speed brakesare used on the CBV in this fashion, a compensating ele-vator deflection will have to accompany brakedeflection.

The effect of wingtip-fin-mounted speed brakes ispresented in figure 9(b). Because the surfaces were rela-tively small, deflections up to 60° were tested. Thebrakes extended out from the wingtips and were effectivein producing drag increments linear with speed-brakedeflection. The line of action of the drag increment fromthe brakes was close to the estimated center of gravity, solittle change in pitching moment resulted. Their use alsodelayed the onset of flow separation, as can be seen byexamining the lift and moment curves in the 12° to 17°angle-of-attack range.

The speed brakes for the nose-fin model weremounted on the sides of the body over the wing.Figure 9(c) shows that deflecting the side-body brakesproduced drag increments linear with deflection. How-ever, the brakes were largely ineffective in the upper por-tion of the angle-of-attack range. Also, lift was reducedas the brakes were deflected. Deflection of the brakes insuch close proximity to the wing upper surface disturbedthe flow over the wing, resulting in a small loss of lift.This loss of lift caused the pitching-moment curves toshift slightly upward with speed-brake deflection. Theseeffects are quite different from those found in tests atM = 2.0 to 4.6 (ref. 9). In the supersonic speed range,deflection of the side-body brakes increased drag onlyslightly and introduced a more significant nose-up pitch-ing moment.

Lateral Characteristics

Lateral-directional stability.The lateral-directionalcharacteristics of the CBV appear in figure 10 in the formof the stability parameters , , and plottedagainst angle of attack. Data are shown for the modelwith all fin configurations and with fins off for an eleva-tor setting of−14°. This elevator setting trims the modelat an angle of attack that approximates a landingapproach condition.

The CBV with aft center fin was directionally stable,with the value of increasing slightly over the testangle-of-attack range. The model with the other finarrangements was, as expected, directionally unstable.The nose fin was directionally destabilizing, whereaswingtip fins generally added increments. Little dif-

ference occurred in the effective dihedral parameterbetween the fins off and nose-fin configurations.

The center-fin model had negative values of overthe test angle-of-attack range.

Yaw-control effects.Rudder effectiveness for eachof the vertical-fin arrangements is presented in figure 11.Yaw control for the aft center-fin configuration(fig. 11(a)) was accomplished by deflecting one half ofthe split rudder outboard and leaving the other half unde-flected. The rudder produced almost constant values ofyawing moment across the test angle-of-attack range.The retention of effectiveness at high angles of attackprobably was caused by the large size of the center fin,which placed it above the influence of the fuselage andwing. As expected, the size and placement of the finabove the vehicle center of gravity also caused largeadverse rolling moments with rudder deflection. Thevalue of the rolling moment, which was about one-thirdthat of the yawing moment at the trim angle of attack of14°, increased significantly with decreased angle ofattack.

The effect of deflecting a wingtip-fin controller isshown in figure 11(b). The data indicate that the yawingmoments produced by the controller decreased onlyslightly with increasing angle of attack up to about 13°.At higher angles, a large increase in adverse rolling-moment increment occurred. A clue to the adverse roll-ing moment is seen in figure 12 when one compares theflow pattern over the wing with the wingtip-fin controllerdeflected with the flow pattern with the controller unde-flected. The deflected controller reduced the outflow onthe wing thereby delaying flow separation near thewingtip. With one controller deflected, an adverse rollingmoment resulted from the asymmetric flow conditionsover the wings.

The nose fin, placed far forward to take advantage ofthe long moment arm created by the 72 percent of bodylength center-of-gravity location, was very effective ingenerating yawing moment (fig. 11(c)). The effective-ness of the deflected nose fin increased almost linearlywith increasing angle of attack. This increased effective-ness is probably due to the increase in cross flow aroundthe circular body as the pitch angle increased. Theseyawing moments were accompanied by only small roll-ing moments up to the wing flow-separation angle.

Roll-control effects.Roll control results from differ-entially deflecting the dedicated aileron-control surfaceson the outer wing trailing edge. The effectiveness valuesare presented in figure 13 for surfaces set at 5° on the leftand−5° on the right. Tests were made with wingtip finson and off. The wingtip-fin controllers were undeflected.

CYβ Cnβ Clβ

Cnβ

Cnβ

Clβ–Clβ

5

The data indicate that aileron roll effectivenessincreased linearly with angle of attack up to angles whereflow separation occurred on the wingtip. Small adverseyawing moments accompanied roll control. The controlauthority of the ailerons appears to be sufficient tocounter the adverse rolling moments noted for the center-fin and wingtip-fin arrangements near the trim angles ofattack desired for landing. Insufficient data exist to indi-cate whether the ailerons would have sufficient controleffectiveness to counter adverse roll caused by the centerfin at reduced angles of attack.

Concluding Remarks

Tests of a circular body spacecraft model have beenconducted in the Langley 7- by 10-Foot High-SpeedWind Tunnel. The design is an option considered forsingle- or multistage-to-orbit vehicles. The model had abody with a circular cross section and a clipped deltawing. Three vertical fin arrangements were investigated:a conventional aft-mounted center vertical fin, wingtipfins, and a nose-mounted vertical fin.

Results of the investigation indicated that the config-uration was longitudinally stable about the estimatedcenter-of-gravity position of 72 percent of body length.The model had sufficient pitch-control authority withelevators to produce stable trim over a wide range ofangle of attack regardless of fin arrangement. The maxi-mum trimmed lift-drag ratio for the aft center-fin config-uration was less than 5, whereas the other configurationshad values above 6.

The aft center-fin configuration was directionallystable for all angles of attack. The wingtip and nose finswere not intended to produce directional stability, anddid not. The rudder-like surfaces on the wingtip fins andthe nose fin were designed as active controls to produceartificial directional stability. These controls were effec-tive in producing yawing moment throughout the testangle-of-attack range. As expected, deflection of the rud-der on the aft center fin produced large adverse rollingmoments. Small rolling-moment values resulted fromyaw control of the nose fin, whereas large, adverserolling-moment increments resulted from wingtip-fincontroller deflection above an angle of attack of 13°.

Differential deflections of the aileron surfaces on theoutboard wing trailing edge were effective in producingrolling moments with wingtip fins on or off. Aileroneffectiveness was sufficient to counter the adverse rollingmoments noted for the aft center fin and wingtip fin

configurations in the angle-of-attack range desired forlanding.

NASA Langley Research CenterHampton, VA 23681-0001April 18, 1996

References1. 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 II. SAE Tech. Paper Ser. 871335, June1987.

4. Holloway, Paul F.; and Talay, Theodore A.:Space Transporta-tion Systems—Beyond 2000.IAF Paper 87-188, Oct. 1987.

5. Talay, Theodore A.; and Morris, W. Douglas: AdvancedManned Launch Systems. Paper presented atAAAF, DGLR,Royal Aeronautical Society, and ESA Second European Aero-space Conference on Progress in Space Transportation (Bonn,Federal Republic of Germany), May 22–24, 1989.

6. Powell, Richard W.; and Freeman, Delma C., Jr.: Applicationof a Tip-Fin Controller to the Shuttle Orbiter for ImprovedYaw Control. AIAA-81-0074, Jan. 1981.

7. MacConochie, I. O.; and Klich, P. J.:Technologies Involved inConfiguring an Advanced Earth-to-Orbit Transport for LowStructural Mass. SAWE Paper 1380, May 1980.

8. Wells, William L.; MacConochie, Ian O.; Helms, Vernon T.;and Raney, David: Heating Rate Distributions at Mach 10 ona Circular Body Earth-to-Orbit Transport Vehicle. AIAA-85-0974, June 1985.

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

10. Ware, George M.; Engelund, Walter C.; and MacConochie, IanO.: Supersonic Aerodynamic Characteristics of a CircularBody Earth-to-Orbit Vehicle. NASA TM-4533, 1994.

11. Peñaranda, Frank E.; and Freda, M. Shannon, eds.:Aeronauti-cal Facilities Catalogue. Volume 1—Wind Tunnels.NASARP-1132, 1985.

12. Braslow, Albert L.; Hicks, Raymond, M.; and Harris, Roy V.,Jr.: Use of Grit-Type Boundary-Layer-Transition Trips.Con-ference on Aircraft Aerodynamics,NASA SP-124, 1966,pp. 19–36.

13. Montoya, Lawrence C.; Jacobs, Peter F.; and Flechner, StuartG.: Effect of Winglets on a First-Generation Jet TransportWing. III—Pressure and Spanwise Load Distributions for aSemispan Model at Mach 0.30. NASA TN D-8478, 1977.

6

Table 1. Geometric Characteristics of Circular Body Model

Body:Length, in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.00Base area, in2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66.96

Wing:Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 0010-64Mean aerodynamic chord (reference length), in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16.00Span (reference span), in. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35.04Area to body centerline, reference area, in2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 486.60Area, exposed outside of body, in2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 305.08

Center vertical fin:Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Double wedgeArea, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100.00

Tip fins (each):Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Modified wedgeArea, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.60

Nose fin:Airfoil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Modified NACA 0015Area, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.36

Control surfaces (each):Elevator area, in2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.56Body flap area, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24.96Aileron area, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.84Wingtip-fin controller and speed brake area, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.08Aft center-fin rudder and speed brake area, in2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.80Body speed brake area, in2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.72

7

Figure 1. Axis system used in investigation with positive directions of forces and moments.

Side forceY

Wind direction

X

β

Z

α

Wind direction

Lift

Drag

Pitchingmoment

Rollingmoment

Yawingmoment

X

8

L-84-10,263(a) Model in Langley 7- by 10-Foot High-Speed Tunnel.

Figure 2. Circular body earth-to-orbit vehicle model.

9

(b) Model showing three fin arrangements investigated.

Figure 2. Concluded.

Speed brake(used with nose-fin configuration)

Elevator

Aileron

Nose-finconfiguration

Wingtip-fin configuration(with controllers and

speed brakes)

Center-fin configuration(with rudder and

speed brake)

10

(a) Wing body.

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

8.80

4.00

37.44

52.00

.50

16.24

1.00 rad.

37.10 rad.

Tangency point

4.49

1.45

2.88

Bodyflap

3.66

8.66 35.04

Sref = 486.60 in.2

c = 16.00 in.

b = 35.04 in.

l = 52.00 in.

23.28

47°

2.88

7°

1.56

8.66 dia.

11

(b) Wingtip fin.

Figure 3. Continued.

3.923.36

.48

4.00

30°

1.52

1.72

.50.36

10°

1.5° toe-in

Controller and speed brake

Wingtip-fin controllerand speed brake

12

(c) Aft center fin.

Figure 3. Continued.

43.84

63.06

10.16

4.74

2.50

1.88

4.06 13.50

7.50

13.00

3.20

45°

1.22

18.30

10°.156 rad.

Split rudderand speed brake

13

(d) Nose fin.

Figure 3. Concluded.

3.74

1.86

5.34

Speed brake

Speed brake

Nose fin

Hinge line

1.06

4.00

1.34

20°

35°

1.90

3.86

Pivot axis

3.72

.116 rad.

1.40 rad.

1.70

52.00

.92.05 rad.

.46

.86

.22 rad.

14

Figure 4. Longitudinal characteristics of circular body model with various fin arrangements.δe = 0°.

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.2

-.1

0

.1

.2

.3

.4

.5

.6

.7

.8

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.28

-.24

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

Cm

-2

-1

0

1

2

3

4

5

6

7

8

L/D

Configuration

Fins offCenter finWingtip finsNose fin

15

(a) α = 6.35°. (b) α = 10.64°.

(c) α = 12.80°. (d) α = 14.83°.

Figure 5. Oil flow on fins-off configuration atM = 0.3, showing upper surface of right-side wing withδe = 0°.

16

(a) α = 6.36°. (b) α = 10.72°.

(c) α = 12.86°. (d) α = 14.90°.

Figure 6. Oil flow on tip-fin configuration atM = 0.3, showing right-wing upper surface of right-side wing withδe = 0°.

17

(a) Center fin.

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

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.3

-.2

-.1

0

.1

.2

.3

.4

.5

.6

.7

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-4

-3

-2

-1

0

1

2

3

4

5

6

L/D

δe, deg

0-14-20

18

(b) Wingtip fins.

Figure 7. Continued.

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.4

-.2

0

.2

.4

.6

.8

1.0

1.2

1.4

1.6

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-6.0

-4.5

-3.0

-1.5

0

1.5

3.0

4.5

6.0

7.5

9.0

L/D

δe, deg

0-14-20

19

(c) Nose fin.

Figure 7. Concluded.

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.4

-.2

0

.2

.4

.6

.8

1.0

1.2

1.4

1.6

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-6.0

-4.5

-3.0

-1.5

0

1.5

3.0

4.5

6.0

7.5

9.0

L/D

δe, deg

0-14-20

20

(a) α = 6.35°. (b) α = 10.64°.

(c) α = 12.80°. (d) α = 14.83°.

Figure 8. Oil flow on fins-off configuration atM = 0.3, showing upper surface of left-side wing withδe = −14°.

21

(a) Center fin.

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

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.2

-.1

0

.1

.2

.3

.4

.5

.6

.7

.8

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-3

-2

-1

0

1

2

3

4

5

6

7

L/D

δSB

, deg

07.5

15

22

(b) Wingtip fins.

Figure 9. Continued.

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.1

0

.1

.2

.3

.4

.5

.6

.7

.8

.9

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-2

-1

0

1

2

3

4

5

6

7

8

L/D

δSB

, deg

0204060

23

(c) Nose fin.

Figure 9. Concluded.

-8 -4 0 4 8 12 16 20 24

α, deg

0

.04

.08

.12

.16

.20

.24

.28

.32

.36

.40

CD

-.1

0

.1

.2

.3

.4

.5

.6

.7

.8

.9

CL

-8 -4 0 4 8 12 16 20 24

α, deg

-.20

-.16

-.12

-.08

-.04

0

.04

.08

.12

.16

.20

Cm

-2

-1

0

1

2

3

4

5

6

7

8

L/D

δSB

, deg

0102030

24

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

-8 -4 0 4 8 12 16 20 24

α, deg

-.010

-.008

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

Clβ

Configuration

Fins offCenter finWingtip finsNose fin

-.10

-.08

-.06

-.04

-.02

0

.02

.04

.06

.08

.10

CYβ

-8 -4 0 4 8 12 16 20 24

α, deg

-.010

-.008

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

Cnβ

Stable

Stable

25

(a) Center fin.

Figure 11. Effect of rudder deflection as yaw control for circular body model.δe = −14°.

-8 -4 0 4 8 12 16 20 24

α, deg

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

.012

.014

∆Clδ

r, deg

7.515

-.05

-.04

-.03

-.02

-.01

0

.01

.02

.03

.04

.05

∆CY

-8 -4 0 4 8 12 16 20 24

α, deg

-.018

-.016

-.014

-.012

-.010

-.008

-.006

-.004

-.002

0

.002

∆Cn

26

(b) Wingtip fins.

Figure 11. Continued.

-8 -4 0 4 8 12 16 20 24

α, deg

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

.012

.014

∆Clδ

TF, deg

204060

-.05

-.04

-.03

-.02

-.01

0

.01

.02

.03

.04

.05

∆CY

-8 -4 0 4 8 12 16 20 24

α, deg

-.018

-.016

-.014

-.012

-.010

-.008

-.006

-.004

-.002

0

.002

∆Cn

27

(c) Nose fin.

Figure 11. Concluded.

-8 -4 0 4 8 12 16 20 24

α, deg

-.014

-.012

-.010

-.008

-.006

-.004

-.002

0

.002

.004

.006

∆Clδ

n, deg

51015

-.08

-.07

-.06

-.05

-.04

-.03

-.02

-.01

0

.01

.02

∆CY

-8 -4 0 4 8 12 16 20 24

α, deg

-.036

-.032

-.028

-.024

-.020

-.016

-.012

-.008

-.004

0

.004

∆Cn

28

(a) δTF = 0°.

(b) δTF = 20°.

Figure 12. Oil flow on tip-fin configuration atM = 0.3, showing right-wing upper surface,α = 12.86°, δe = 0°.

29

Figure 13. Effect of outboard ailerons as roll control for circular body model.δa = 5° andδe = 0°.

-8 -4 0 4 8 12 16 20 24

α, deg

-.010

-.008

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

∆C lConfiguration

Fins offWingtip fins

-.05

-.04

-.03

-.02

-.01

0

.01

.02

.03

.04

.05

∆CY

-8 -4 0 4 8 12 16 20 24

α, deg

-.010

-.008

-.006

-.004

-.002

0

.002

.004

.006

.008

.010

∆C n

Form ApprovedOMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of thiscollection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 JeffersonDavis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

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

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

6. AUTHOR(S)

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

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

11. SUPPLEMENTARY NOTES

8. PERFORMING ORGANIZATIONREPORT NUMBER

10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

14. SUBJECT TERMS

17. SECURITY CLASSIFICATIONOF REPORT

18. SECURITY CLASSIFICATIONOF THIS PAGE

19. SECURITY CLASSIFICATIONOF ABSTRACT

20. LIMITATIONOF ABSTRACT

15. NUMBER OF PAGES

16. PRICE CODE

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)Prescribed by ANSI Std. Z39-18298-102

REPORT DOCUMENTATION PAGE

July 1996 Technical Memorandum

Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-OrbitVehicle WU 242-20-08-01

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

L-17429

NASA TM-4726

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

A test of a generic reusable earth-to-orbit transport was conducted in the 7- by 10-Foot high-speed tunnel atthe Langley Research Center at Mach number 0.3. The model had a body with a circular cross section and athick clipped delta wing as the major lifting surface. For directional control, three different vertical finarrangements were investigated: a conventional aft-mounted center vertical fin, wingtip fins, and a nose-mounted vertical fin. The configuration was longitudinally stable about the estimated center-of-gravity posi-tion of 0.72 body length and had sufficient pitch-control authority for stable trim over a wide range of angleof attack, regardless of fin arrangement. The maximum trimmed lift/drag ratio for the aft center-fin configura-tion was less than 5, whereas the other configurations had values of above 6. The aft center-fin configurationwas directionally stable for all angles of attack tested. The wingtip and nose fins were not intended to producedirectional stability but to be active controllers for artificial stabilization. Small rolling-moment valuesresulted from yaw control of the nose fin. Large adverse rolling-moment increments resulted from tip-fin con-troller deflection above 13° angle of attack. Flow visualization indicated that the adverse rolling-momentincrements were probably caused by the influence of the deflected tip-fin controller on wing flow separation.

Aerodynamics; Subsonic; Spacecraft; Booster 30

A03

NASA Langley Research CenterHampton, VA 23681-0001

National Aeronautics and Space AdministrationWashington, DC 20546-0001

Unclassified–UnlimitedSubject Categories 15 and 02Availability: NASA CASI (301) 621-0390

Unclassified Unclassified Unclassified