AERODYNAMIC DESIGN OF THE SPACE SHUTTLE ORBITER by

W. E. Bornemann T.E. Surber Manager, Space Shuttle Aerodynamics Supervisor, Orbiter Aerodyanmics Rockwell International Corporation and Rockwell International Corporation

Space Systems Group Space Systems Group 12214 Lakewood Boulevard 12214 Lakewood Boulevard

Downey, CA 90241 Downey, CA 90241

The Space Shuttle Vehicle i s being developed by the NASA to provide capability for lower cost space operations i n the 1980's and beyond. This paper describes Shuttle Orbiter aerodynamic design conducted by Rockwell International under contract to NASA. Aerodynamic cr i ter ia key to establishing the external configuration are discussed together with evolution of the design including effects of wing-body blending on high angle of attack aerodynamics.

An overview of the wind tunnel program i s given and aerodynamic characteristics of the final config- uration are described. Aerodynamic parameters cr i t ical to definition of Orbiter entry control and performance are identified. During entry, the Orbiter f l i es over an angle of attack range from 50 to zero degrees. Trim capability and s tabi l i ty and control characteristics are discussed a t crit ical regions in the entry trajectory. Methods are described to define reaction control rocket effectiveness and aero- dynamic interactions during the ini t ia l portion of entry. A t hypersonic speeds, wind tunnel results of viscous interaction effects a t high angles of attack are discussed. In the supersonic region where transi- tion from high to low angle of attack occurs, cr i t ical s tabi l i ty and control parameters and wind tunnel results are described. A t subsonic speeds, comparisons are shown between predicted aerodynamic character- istics and data from the approach and landing flight t es t program.

NOMENCLATURE

Acronyms

AEDC ALT ARC ATP CDR EAFB ET ETR FCF FMO F JSC KSC LaRC OMS OV PDR PRR RCS SRB SSME TAEM WR

Symbols

Arnold Engineering Development Center Approach and Landing Test NASA Ames Research Center Authority to Proceed Critical Design Review Edwards Air Force Base External Tank Eastern Test Range First Captive Flight First Manned Orbital Flight NASA Johnson Space Center NASA Kennedy Space Center NASA Langl ey Research Center (a1 so LRC) Orbital Maneuvering System Orbital Vehicle Prel iminary Design Review Program Requirements Review Reaction Control System Sol id Rocket Booster Space Shuttle Main Engine Terminal Area Energy Management Western Test Range

Span Axial force coefficient Drag force coefficient Lift coefficient Rolling moment coefficient Rolling moment coefficient due to sidesl ip (per degree) Rolling moment coefficient per degree aileron deflection Roll ing moment coefficient per degree rudder deflection Pitching moment coefficient Normal force coefficient Yawing moment coefficient Yawing moment coefficient due to sidesl ip Yawing moment coefficient per degree aileron deflection

Symbols (continued)

Yawing moment coefficient per degree rudder deflection Factor of proportionality in 1 inear viscosity-temperature relation, equation Center of gravity Altitude Lift-to-drag ratio Mach number Mean aerodynamic chord, a1 so c RCS jet'mass flow ratio, equation (4)

mm - - q Dynamic pressure = 1/2pV

2 RE Reynolds number, also Re S Reference area Se Standard error of estimate VD Design touchdown speed VA Viscous parameter, equation (2)

Viscous interaction parameter, equation (1 ) Angle of attack Angle of sidesl ip (positive nose-up) aileron deflection (positive for positive roll ing moment) Body flap deflection (positive for nose- down pitching moment) Elevator deflection (positive for nose- l e f t yawing moment) Standard deviation RCS j e t momentum ratio, equation (3)

Sweep angle Taper ratio Mass, density of a i r

Subscripts

LB Body 1 ength 03 Frees tream

INTRODUCTION

The Space Shuttle Vehicle i s being developed by the NASA to provide capability for lower cost space operations in the 1980's and beyond. The f l ight vehicle consists of a reusable orbiter, an expendable external propellant tank, and two reusable solid rocket boosters. Space Shuttle will be capable of launching a variety of payloads into earth-orbit from either the Eastern Test Range (ETR) a t Kennedy Space Center or the Western Test Range (WTR) a t Vandenberg Air Force Base. Maximum payload capabilities will be 29,480 kg for an easterly launch from ETR and 14,515 kg for launch into polar orbit from WTR.

The orbiter development contract, under the direction of NASA's Johnson Space Center, was awarded to Rockwell International in August 1972. Under this contract, two orbiter vehicles are being built-0V101 was delivered to the f l ight t es t center in February 1977, and approach and landing f l ight testing completed i n October 1977. OV102, the f i r s t orbital fl ight vehicle, i s in assembly and i s scheduled for roll-out in la te 1978. Orbital f l ight testing will begin in 1979.

Aerodynamic considerations have played a significant role in the vehicle design process. The Shuttle must f ly satisfactorily with predicted aerodynamic characteristics; i t i s not feasible to approach f l ight testing by incremental expansion of the altitude and velocity envelope. Consequently, the NASA and Rockwell have given careful attention to the development of an extensive data base derived largely from wind tunnel t es t s , with detailed attention being given to defining uncertainties through s tat is t ical analysis of wind tunnel data and by comparisons of wind tunnel predictions with flight data from previous programs. In addition, the f l ight control system i s being designed to minimize i t s sensitivity to uncertainties i n aerodynamic parameters.

The objectives of this paper are to: (1) briefly describe the Shuttle mission in order to identify key aerodynamic design cr i ter ia ; (2) summarize aerodynamic development of the orbiter; (3) describe the key aerodynamic parameters and their relationship to design and performance of the entry flight system; and (4) sumnarize recent f l ight results which verify the aerodynamic estimates for the approach and landing phase of the Shuttle mission.

VEHICLEIMISSION DESCRIPTION

The Shuttle Vehicle consists of four major elements: the orbiter; main engines (SSME); external tank (ET); and two solid rocket boosters (SRB). Overall vehicle configuration i s i l lustrated in Figure 1. The external tank contains the liquid oxygenlliquid hydrogen propellants used by the main engi?es during ascent. Liquid oxygen i s located in the forward tank to maintain an acceptable center of gravlty for the combined vehicle. Nozzles on each booster are gimballed to augment control during ascent.

The orbiter, Figure 2, i s a double-delta wing configuration comparable in size to a modern transport aircraft . Normally , the orbiter carries a crew of four-commander, pi lo t , mission special i s t , and payload specialist-with provision for as many as seven persons. The orbiter can remain in orbit nominally for seven days (up to 30 with special payloads), return to earth with personnel and payload, land like an airplane, and be refurbished for a subsequent flight in 14 days. Three main rocket engines mounted in the a f t section of the orbiter provide propulsive thrust during ascent. These two million Newton thrust 1 iquid oxygen11 iquid hydrogen engines are gimbal led in pitch and yaw to provide thrust vector control. Smaller orbital maneuvering system (OMS) rocket engines are also located in the a f t section to provide final impulse for orbit insertion, orbital maneuvers, and deorbit. Reaction control rockets (RCS) are located in both the forward and a f t section of the orbiter to provide attitude control and three-axis translation during orbit insertion and on-orbit operations. The a f t reaction control rockets are used in combination with aerodynamic surfaces for control during entry. Aerodynamic surface controls include sp l i t elevons along the wing trailing edge; a s p l i t rudder in the vertical f in which can also be flared open to serve as a speed brake during descent; and a hinged body flap located a t the lower a f t end of the fuselage to augment control during descent and landing approach. The body flap also shields the exposed main engine nozzles from aerodynamic heating during entry.

The entire external surface of the orbiter, except the windows, i s protected by reusable insulation to maintain acceptable structural temperatures under entry heating environment. Figure 3 i l l ustrates the application areas for the materials used in the thermal protection subsystem. Application i s as follows:

1. Coated Nomex f e l t i s used in areas where temperatures are less than 67Z°K for entry and 716OK for ascent; i .e., upper cargo bay door, mid- and aft-fuselage sides, upper wing, and OMS pod.

2. Low-temperature reusable surface insulation i s used in those areas where temperatures are below 922OK and above 67Z°K under design heating conditions.

3. High-temperature reusable surface insulation i s used in those areas exposed to temperatures below 1533OK and above 922OK under design heating conditions.

4. Reinforced carbon-carbon i s used on areas such as wing leading edge and nose cap where predicted temperatures exceed 1533°K under design heating conditions.

5. Thermal window panes are used in the crew compartment and high temperature metal i s used for forward reaction control system fairings and elevon upper surface rub seal panels.

6. Thermal barriers are instal led around operable penetrations (main egress hatch, landing gear doors, etc.) to protect against aerothermal heating.

The thermal protection system i s a passive system. I t has been designed for ease of maintenance and for flexibili ty of ground and f l ight operations while satisfying i t s prPmary function of maintaining acceptable airframe outer skin temperatures.

Mission Profi le

Mission performance capabil i ty is sumnarized i n Table 1. A typical mission prof i le is shown i n Figure 4. The Shutt le is launched w i t h the main engines and solid rocket boosters burning in parallel . A maximum dynamic pressure of 31,100 ~ / m ' is experienced approximately 62 seconds a f t e r launch a t 11,280 meters a l t i tude . Booster separation occurs a t 122 seconds a t an a l t i t ude of 43,280 meters, 46.3 kilo- meters downrange from the launch s i t e . The solid rocket boosters descend on parachutes, a re recovered a f t e r water impact, and are refurbished for subsequent reuse.

After booster separation, the orbi ter continues to ascend w i t h main engine cut-off and external tank separation occurring 479 seconds a f t e r 1 if t-off when the orbi ter has reached an a l t i t ude of 115,700 meters. The orbital maneuvering system engines, which provide the additional velocity needed fo r orbital insertion are cut-off approximately 600 seconds a f t e r launch.

After completion of the orbital operations phase, deorbit i s accomplished by retro-fire of the orbital maneuvering engines, and the orbi ter descends to the atmospheric entry interface (nominally, an a l t i tude of 121,920 meters). A typical entry trajectory is shown i n Figure 5. The i n i t i a l entry phase extends to a dynamic pressure level of 957.6 N/m2 (approximately 76,200 meters a l t i tude) during which a t t i tude control from two a f t pods i s blended w i t h aerodynamic surface controls, the l a t t e r gaining i n effectiveness a s dynamic pressure increases. Entry, from a dynamic pressure level of 957.6 N/m2 to a Mach number of less than f ive , i s accomplished a t a high angle of at tack ( i n i t i a l l y 38 degrees) during which the blanketing e f fec t of the wing essential ly precludes any rudder control . Coordinated la tera l - directional control i s provided by combined yaw reaction control j e t s and aileron control. The terminal phase occurs a s angle of at tack i s reduced below 18 degrees. As the orbi ter descends to a l t i tudes where winds can resul t in re la t ive ly large errors in ine r t i a l ly derived a i r data, probes are extended (M=3.5) to provide a i r data r e l a t ive to the vehicle. During a typical normal entry, range control is achieved by bank angle while angle of at tack follows a predetermined schedule to achieve ( a t approximately M = 1.5) an angle somewhat smaller than that corresponding to maximum LID. A downrange capability of up to 7,960 kilo- meters w i t h a cross range capability of 1,815 kilometers may be realized. Subsonic f l i g h t i s achieved a t an a l t i tude of approximately 12,190 meters. Range control during the gliding descent i s obtained by angle of at tack modulation with velocity control maintained by the speed brake. The approach and landing inter- face occurs a t 3,048 meters above ground level and a preflare i s i n i t i a t ed a t an appropriate a l t i tude , followed by a deceleration f loa t and touchdown. The i n i t i a l approach target and f l a re a l t i tude will be scheduled t o provide a minimum of 25 seconds between f l a r e i n i t i a t e and touchdown. Touchdown occurs a t an angle of at tack of about 15 degrees. The nominal touchdown velocity i s 88 meters/sec, and maximum landing speed w i t h a 14,515-kilogram payload is about 106 meters/sec including dispersions fo r hot-day ef fec ts and tailwinds.

Orbiter Aerodynamic Criteria

Aerodynamic c r i t e r i a , Reference 1 , for the orbi ter vehicle require the configuration to perform as both a spacecraft and an a i r c ra f t . Because of t h i s , the external features must be carefully configured to provide the protection and ve r sa t i l i t y required for orbital and atmospheric f l igh t , and the aerodynamic performance and control necessary for unpowered descent and landing. The aerodynamic l ines must ensure performance that i s acceptable over the hypersonic to subsonic speed range, and provide the required cross range capabil i t y and touchdown velocity . Aerodynamic requirements, Table 2, were developed from analysis of the entry phase of the mission. Landing requirements are shown i n Figure 6. S ta t i c s t ab i l i t y was not required since the design c r i t e r i a allowed reliance on the f l i g h t control system to meet flying qual i t ies c r i t e r i a . Early simulations identif ied a f l i g h t control requirement fo r s t a t i c longitudinal s t a b i l i t y to be no more than two percent body length (5.45 percent mean aerodynamic chord) unstable so the pitching moment curve established the a f t center of gravity l imi t a t 67.5 percent of body length. Payload c r i t e r i a established a center of gravity range of 2.5 percent, thus establishing the forward l imi t .

The selected configuration, Figure 2 , evolved from a ser ies of program and technical refinements directed to achieve the vehicle yielding the best combination of performance and cost . This evolution i s discussed further i n a l a t e r section. The double-delta planform combined with a moderately low fineness ra t io (approximately f ive) body minimizes interference heating - effec ts , provides the required cross range requirements, and possesses an acceptable trim and s t a b i l i t y range, Figure 7, over the f l igh t Mach number range.

The orbi ter wing was sized to provide a 88 meters/second touchdown speed (VD) a t a 15-degree angle of at tack ( t a i l scrape a t t i t ude fo r main gear s t r u t compressed, t i r e f l a t ) with body f l ap retracted and the center of gravity a t the forward l imit . The leading edge sweep (45 degrees) and aspect ra t io (2.265) were selected on the .basis of aerothermodynarnic trade studies t o provide the design touchdown speed for a center of gravity a t the forward l imi t with minimum wing s ize and to optimize the wing leading edge thermal protection system fo r a reuse cycle of 100 f l igh t s prior to major rework.

The fuselage was designed to accomnodate a variety of payloads and house the crew and maneuvering control systems. Nose camber, cross section, and upward sloping forebody J d e s were selected to improve hypersonic pitch trim and directional s t ab i l i t y and i n conjunction w i t h wind-body blending, to reduce entry heating on the body sides. Propulsion units fo r entry a t t i t ude control and orbital maneuvering have been incorporated in pods located in the a f t body fair ings. The body f lap i s used to protect the Shuttle main engine during entry and to provide trim capability t o rel ieve elevon loads.

The vert ical t a i l has been sized to provide a low-speed Cng of 0.0013 a t an angle of attack of 13 degrees about a center of gravity located a t the a f t l imit . I t has a reference area of 38.,39 m2 including the rudderispeed brake. The rudder is s p l i t along the orbi ter buttock plane t o provide

directional s t a b i l i t y augmentation i n the hypersonic/supersonic f l i g h t regimes and to apply drag modula- tion for the subsonic f l i g h t phases, approach and landing. The section profi le i s a five-degree, half- angle, 60-40 double-wedge a i r f o i l .

Aerodynamic and aerothermodynamic c r i t e r i a , Reference 2, regarding surface discontinuit ies, thermal protection system t i l e steps and gaps, and waviness a re shown in Figure 8. These c r i t e r i a are based on aerodynamic efficiency requirements of 1 i f t i ng surfaces and the prevention of premature transit ion from laminar to turbulent boundary layers in the high heating portion of entry. Aerodynamic efficiency i s affected to a much greater extent by surface conditions of the forward rather than a f t regions of com- ponents. Hence, tolerance c r i t e r i a are generally more res t r ic t ive for forward regions of the vehicle surfaces and somewhat relaxed a t a f t portions.

DEVELOPMENT APPROACH

Development Schedule

Major program milestones are i l l u s t r a t ed in Figure 9, s tar t ing with authority to proceed (ATP) i n 1972 and culminating w i t h i n i t i a l operational capability in 1980. The orbi ter concept a t ATP was a blended delta wing vehicle based on precontract studies and configured to meet i n i t i a l Shuttle Program requirements. As a resul t of a continuing assessment of system requirements and technical refinements, early in the contract the orbi ter concept was modified to reduce weight and decrease program and operating costs (Reference 3). As discussed in more detail l a t e r , refinements in the aerodynamic configuration led to a double-delta planform incorporating a more e f f i c i en t l i f t i n g surface than the blended delta. The System Requirements Review i n August 1973 finalized technical requirements f o r the Space Shuttle systems ( i .e . , the to ta l vehicle, i t s elements, and the i r ground systems) and approved the design approach of the vehicle and associated support equipment. Preliminary Design Review (PDR) of the f i r s t orbi ter (Orbiter 101) vehicle and subsystems for the approach and landing f l i g h t t e s t program was completed in February 1974, followed by the Preliminary Design Review of the second orbi ter (102) in March 1975. Orbiter 101 roll-out from f inal assembly i n Palmdale, California, took place i n September 1976. The vehicle was mated to the Boeing 747 carr ier a i r c ra f t a t the Dryden Flight Research Center, Edwards Air Force Base, and the f i r s t captive f l i g h t was com leted in February 1977. The f i r s t airlaunch of Orbiter 101 fo r the approach and landing f l i g h t t e s t (ALT~ took place on August 12, 1977, and the f inal f l i g h t was completed on October 26, 1977. Delivery of OVlOl to the Marshall Space Flight Center, Alabama, fo r ground vibration test ing took place i n March 1978. Fabrication and assembly of Orbiter 102, the f i r s t orbital vehicle, began in 1975. Rollout i s scheduled for l a t e 1978, followed by delivery to Kennedy Space Center, Florida, and f i r s t manned orbital f l i g h t i n 1979. The f i r s t six orbital f l i gh t s of the Shuttle are development t e s t f l ights , and the seventh f l i g h t i n 1980 i s considered the i n i t i a l operational capabil i ty f l igh t .

Aerodynamic Design Approach

I t i s conventional i n an a i r c ra f t program to approach f l i g h t demonstration by incremental expansion of the f l igh t envelope. This i s not feasible with the Shuttle vehicle. Once Shuttle i s launched, i t is comnitted to f l i g h t over the complete mission profi le from ascent to orbi ter insertion, deorbit, entry, and landing. Flight characterist ics must be based on aerodynamic data derived from ground test ing and analysis. Careful at tention has been given t o the interactions between f l i g h t control systems design and aerodynamic characterist ics, and allowance has been made fo r uncertainties in basic aero data in f l i g h t control design. Predicted aerodynamic characterist ics have been derived from extensive wind tunnel t e s t s which have included a systematic investigation of data uncertainties, nonlinear ef fec ts , and ef fec ts of wind tunnel instal 1 ation, blockage, and shock wave ref1 ections . The Langl ey Research Center conducted detailed wind tunnel investigations of control surface characterist ics and nonlinear aerodynamic ef fec ts , (I .e., Refer- ence 4) to support development of the data base.

Wind Tunnel Program

Key t o Space Shuttle development has been the acquisition of wind tunnel t e s t data t o support design and evaluation by providing a continuously maturing data base reflecting configuration and subsystem updates. By f i r s t orbital f l i gh t in 1979, approximately 40,100 to ta l wind tunnel t e s t hours will have been conducted fo r aerodynamics, heat transfer, and structural dynamics, consisting of approximately 20,200 fo r the orbi ter vehicle, 16,100 fo r the mated launch configuration, and 3,800 for the ca r r i e r a i r c ra f t program, Table 3. A to ta l of 94 models have been built-38 aerodyamic, 36 heat transfer, and 20 structural dynamic, Table 4. All wind tunnel test ing i s coordinated with and approved by NASA management a t JSC.

In order t o accurately simulate f l igh t conditions in a wind tunnel, Reynolds number and Mach number must be matched. Problems in flow simulation (Reference 5, NASA CP-2009) occur when the geometric scaling of viscous flow i s important, or when coupling between the viscous surface'.flow and the external flow f i e ld i s strong. In the f i r s t case, the boundary layer can be considered separately from the inviscid flow f i e ld , and viscous ef fec ts can be scaled. This holds for Mach numbers up to about 10. I t i s well known, fo r example, t ha t skin f r ic t ion varies w i t h Reynolds number in a predictable manner and can be scaled to f l igh t conditions from suitable wind tunnel results .

For Mach numbers greater than about 10, a pressure interaction resul ts from the outward streamline deflection induced by a thick boundary layer, and the viscous-inviscid interaction can no longer be neglected. For th i s case, there are two classical simulation parameters comnonly considered:

(1) ym, the viscous interaction parameter introduced by Hayes and Probstein (Reference 6)

(2) TL, the viscous parameter introduced by Whitfield and Griffi th (Reference 7)

where M i s the freestream Mach number, C; i s the factor of proportionality i n the l inear viscosi?y-temperature relat ion (Reference 8 ) , and Rb.. i s the freestream Reynolds number based on x. The parameter Ym i s the relevant parametgr fo r the "pressure" in both the strong and weak interaction cases; whereas V' i s the relevant parameter in terms of "pressure coefficient" ( i .e . , xm/qm). For s h t t l e , i t has been observed that vL correlates total aerodynamic coefficients bet ter than Xm, and consequently, has been used as the hypersonic simulation parameter.

Figure 10 shows a comparison between f l i g h t Re and vL and the simulation capabil i ty of typical wind tunnels used to develop the Orbiter aerodynamic data base. I t i s seen tha t the tunnel capabil i t ies closely match f l ight simulation requirements .

Orbiter aerodynamic t e s t hours are summarized in Figure 11 which i l l u s t r a t e s the phasing and relation- ship to program milestones, and the distr ibution by speed range. Approximately 38 percent of the hours were uti l ized in subsonic t e s t , 44 percent in the transonitlsupersonic range, and 18 percent in the hyper- sonic test ing. Four t e s t phases will be completed by f i r s t orbital f l i gh t . The f i r s t was a configuration definition phase to develop wing a i r fo i l section and planform geometry in support of design trades to reduce orbi ter weight which led to selection of the double-delta arrangement. This phase was completed by System Requirements Review in August 1973. The second phase extending essential ly to OV102 Prel imi nary Design Review i n March 1975 was dedicated t o refinement of the orbi ter vehicle and development of the ferry f l igh t configuration. A t a i l cone configuration was developed to improve ferry performance and pro- vide longer duration orbi ter f l i gh t s during the approach and landing t e s t program. Orbiter t e s t s were conducted to determine basic s t a b i l i t y and control capabil i ty over the complete entry speed range. Control surface effectiveness and hinge moments were measured to support preliminary design and sizing of actuators. In i t ia l RCS interaction t e s t s were conducted to support entry control analysis. Tests were also performed to measure Reynolds number and viscous interaction ef fec ts and to identify wind tunnel s t ing t a r e correc- tions to t e s t data. In addition, the Langley Research Center conducted test ing t o measure orbi ter damping derivatives.

The th i rd phase of the t e s t program was implemented following OV102 Preliminary Design Review in March 1975 to provide more detailed OV102 design data and to verify the aerodynamic characterist ics of OV101 prior t o f i r s t captive f l igh t in February 1977. The Langley Research Center program t o investigate nonlinear aerodynamic and control surface interaction characterist ics was continued during th i s phase. In addition, extensive test ing was conducted to develop the orbi ter a i r data system and provide sensor ca l i - brations for both OVlOl and 102. OVlOl verif ication test ing was also completed during t h i s period u t i l i z - ing a 0.36-scale model i n the Ames Research Center 40x80-ft (12.2x24.4 m) wind tunnel. This model was an accurate replica of the actual OVlOl f l i g h t vehicle and incorporated simulated thermal protection system t i l e s , outer moldline protuberances, and main engine and reaction control system exhaust nozzles.

The final phase, in i t ia ted in early 1978, i s s t i l l i n progress and i s directed toward verif ication of OV102 characterist ics prior to f i r s t orbital f l i gh t . Two models are employed (0.05 and 0.02 scale) to cover the speed range from Mach 16 t o 0.2. Figure 12, i l l u s t r a t e s the deta i l s of protuberances, cavit ies, and thermal sea ls simulated on the mdels .

Aerodynamic Uncertainties

Allowance has been made fo r uncertainties in basic aerodynamic data used i n design of the Shuttle Vehicle, subsystems, and early mission profi les, Reference 9. Two categories of uncertainties have been defined: 1) Tolerances, which account fo r wind tunnel data accuracy and manufacturing tolerances; and 2) Variations, which account fo r unknowns i n extrapolation of model data to free-fl ight . "Tolerances" are used i n subsystem design, and were derived from a s t a t i s t i c a l analysis of wind tunnel data in which t e s t s were conducted using the same models in several different wind tunnels and using different scale models in the same wind tunnel. "Variations" are used in establishing f l igh t t e s t ppans and constraints, and were determined from comparisons between predicted aerodynamics and f l ight t e s t resul ts from 1 i f t ing entry vehicles and selected high-speed a i r c ra f t . The f l i g h t data will allow reductions of the variations and removal of corresponding f l igh t placards to achieve operational capability.

A mu1 t i p l e regression analysis computer program, Reference 10, was used t o determine the "to1 erances" on the derivatives CL, CD, Cm versus a; Cn, Ca versus 8, and 6,. Utilization of the program involved inputting available se t s of wind tunnel data for a specified coefficient versus a , B, 6a or 6, a t given conditions of Mach number, control surface se t t ing , e tc . , with a proposed fowh of curve-fit; e.g., CL = KO + Kcc + K2a2 - . . . K5a5. The regression program s t a t i s t i c a l l y determines which terms of the proposed

curve-fit equation are significant and eliminates those which are not significant by performing a least-square curve-fit of the tes t data. Subsequent to selecting a "best" curve-fit, the deviation of each test point from the curve i s computed and the Standard Error of estimate, Se, (which i s a measure of the standard deviation, a ) i s multiplied by three to estimate the three-sigma (30) tolerance of the aero- dynamic coefficient, C, being analyzed. The three-sigma tolerance i s an increment or band about a nominal value of the aerodynamic coefficient, C, for any given Mach number, body flap deflection, etc., where the probability that a measured coefficient a t the specified condition l i es within C +3u i s 99.73 percent. An example of the procedure for determining the tolerance on l i f t coefficient i s shown in Figure 13. The regression program yields a polynominal expression for 1 i f t coefficient in terms of angle of attack:

The standard error of estimate becomes Se = 0.0146 and the corresponding three-sigma tolerance on l i f t coefficient a t Mach 5.0 becomes 0.0438.

"Variations" were developed for three speed regimes from comparisons of f l ight and predicted values based on wind tunnel t es t data for representative vehicles, constructing the bounds of the data poilits and applying the larger bound as a plus or minus value. An example of the procedure used to develop variation uncertainties i s displayed in Figure 14. The figure presents a comparison between predictions based on wind tunnel results and f l ight measured values of normal force for selected aircraft and space vehicles. The data bands were selected on the basis of engineering ~udgment and weighting "Shuttle-1 i ke" configura- tions more heavily than the 1 ifting bodies. The speed regime groups were M ( 0.8, 0.8 5 M ( 1.2 and M 2 1.2. The ratio of variation to tolerance a t Mach 10.0 was assumed to apply throughout the viscous interaction speed region.

Later in this report where Orbiter flight data from the approach and landing tes t program are dis- cussed, comparisons are shown between estimated tolerances and variations for several aerodynamic param- eters. The measured flight tes t data points are seen to be distributed about the nominal value and to fall within the predicted tolerance band, and well within the estimated variations. I t i s anticipated that further correlation with f l ight data will permit reduction of the variations and removal of corre- sponding f l ight placards t o achieve full operational capability.

CONFIGURATION EVOLUTION

Stability, control, and performance requirements for aerodynamic configuration design of the orbiter vehicle are, for the most part, established by the entry and recovery phases of flight. Consequently, i t i s these phases of flight which were key in determining aerodynamic requirements for the orbiter external arrangement. On the other hand, design airload conditions are primarily determined from the ascent phase.

Design issues key to achieving the proper aerodynamic balance to provide s tabi l i ty , control, and center of gravity range capability across the entrylrecovery flight regime are wing design, wing-body integration, and integration of aerodynamic and f l ight control requirements. Wing design was key because of i t s influence on vehicle weight, thermal environment, aerodynamic s tabi l i ty , buffet characteristics, and gliding and landing performance capability. Ning-body integration was important in obtaining a balanced aerodynamic configuration capable of trim and control over the entire speed range, and in mini- mizing thermal envi ronment due to interference flow effects. Fuselage dimensions were 1 argely fixed by payload size and packaging efficiency while aerodynamic and aerothermodynamic considerations establ ished forebody shape and local contours. Integration of aerodynamic control requirements was of major importance in meeting flying qua1 i ty goals in a l l fl ight regimes, and minimizing vehicle weight as affected by control surf ace arrangement, size, and actuator requirements .

Prior to Shuttle Program go-ahead in August 1972, Rockwell International participated in extensive NASA-funded Shuttle System studies during which numerous trades (Reference 1) were conducted to determine Shuttle operational cost effectiveness, desired configuration and geometry, major subsystem definition, and identification of major design drivers for the orbiter configuration. Design requirements found to be key configuration drivers are landing speed; payload size, weight, and center of gravity envelope; entry cross range and aerodynamic heating; stability and control requirements; and flyin qualities. From these studies emerged a basepoint configuration a t Shuttle Program authority to proceed TATP). Following ATP, further trade studies were conducted a t NASAIJSC and Rockwell to refine the basepoint design. Essentially four aerodynamic basepoints were evaluated in arriving a t the final selected design, as summarized in Figure 15.

ATP Configuration

For the ATP orbiter aerodynamic configuration, Rockwell selected a blended delta wing-body design t o meet NASA mission requirements. Selection of the external arrangement was based on results of previous investigations a t the NASA centers, and Rockwell design studies, supported by 4300 hours of wind tunnel testing. The orbiter aerodynamic shape incorporated a 50-degree swept delta wing planfom sized to provide 77.2 mlsec design touchdown speed with 18,100 kilograms return payload. Hypersonic LID was 1.3 a t 34-degree angle of attack, and maximum subsonic LID was 5.7. Elevons were sized t o provide trim a t hyper- sonic speeds over an angle of attack range from 20 to 50 degrees with an operational center of gravity range of three percent body length. The cargo bay provided a 4.57 meter diameter by 18.2 meter long volume for a wide variety of payloads. Cargo deploymentlretrieval manipulators were stowed in a dorsal fairing along the top of the payload bay doors. Provision was made for installing four airbreathing engines in the a f t portion of the payload bay for early development flights. Three main propulsion system rocket engines were located a t the base of the a f t fuselage, and on-orbit propulsion engines were installed

in two removable pod modules alongside the a f t fuselage. Reaction control rocket engines were located in the a f t pods and in the forward fuselage compartment.

PRR - PDR Configuration

Upon initiation of Shuttle go-ahead, the development aerodynamic wind tunnel program was implemented, and further trade studies were conducted a t NASA/JSC and Rockwell to refine the ATP basepoint. The ATP and PRR orbiters (Figure 15) were both blended delta wing configurations which were externally similar. The most obvious changes were: (1) a redesigned forebody to accomodate internal packaging revisions; (2) the movement of the OMSIRCS pod from the side of the a f t fuselage to the shoulder location; and (3) deletion of airbreathing propulsion for landing assis t following orbital fl ights.

A series of wind tunnel tes ts conducted over the Mach number range from 0.26 to 7.4, indicated revised wing twist, camber, and fuselage blending would improve low-speed 1 i f t capability . Results showed that reducing the wing-body f i l l e t radius and changing from a faired to a straight f i l l e t (wing-glove) increased the trim C significantly. In addition, significant system requirement changes were made by NASA to reduce vehicke weight and cost. Orbiter down payload weight was reduced from 18,100 to 11,300 kilograms, and the vehicle resized from a design dry weight of 77,100 kilograms to 68,000 kilograms. The minimum subsonic s tabi l i ty requirements were reduced from three percent to 0.5 percent body length s ta t ic margin a t the forward center of gravity. For the PDR configuration, Reference 3, direction was received from NASA to modify the wing planform to a double-delta design, and the wing was resized to meet the reduced dry weight and payload requirements. A 45/79 degree wing planform with reduced glove leading edge radius was incorporated for improved subsonic performance, Figures 16 and 17. Improved low-speed perform- ance and the reduced s ta t ic margin requirement permitted a reduction in wing size from 299 to 250 square meters and resulted in rebalancing the orbiter vehicle to meet s tabi l i ty and control requirements.

CDR Configuration

Wind tunnel investigations of the PDR configuration revealed a need for further configuration refinement. Aerodynamic tests showed a difficulty in providing trim capability a t the forward center of gravity in the supersonic f l ight regime. Aeroheating tests indicated the blunt fuselage nose resulted in early transitional flow and high temperatures along the lower body surface. Also, wing incidence, camber, and thickness distributions designed for maximum subsonic performance 1 ed to local fairings on the 1 ower wing and fuselage surfaces which caused high local heating.

Changes incorporated in the fuselage nose section are illustrated in Figure 18. The blunt nose shape was modified to a cross section which was basically parabolic in plan and side-view. Winglfuselage fairings along the bottom of the orbiter were modified to provide a thermodynamically aceptable smooth lower surface with minimum reverse curvature, Figure 19. Leading edge sweep of the glove was slightly changed (from 78 to 81 degrees) as a result of refairing into the modified fuselage nose.

To achieve the best combination of performance and cost, further configuration refinements were made. The down payload requirement was increased to 14,500 kilograms and the design center of gravity range established a t 2.5 percent body length. The OMS pod forebody fairing which extended onto the cargo bay door was shortened to reduce weight and simplify the door-to-fuselage seal design. In addition, the manipulator a m dorsal fairing along the top of the payload bay doors was deleted, and the manipulator was stowed inside the payload bay, Figure 20. Aerodynamic cr i ter ia for the final configuration are listed previously in Table 2.

ORBITER AERODYNAMIC CHARACTERISTICS

Aerodynamics characteristics of the final orbiter vehicle, Reference 2, are sunmarized in this section. These characteristics were derived from wind tunnel t es t results adjusted to account for scale effects or differences between model configurations and the final orbiter vehicle. Aeroelastic corrections have been estimated by standard methods. Wind tunnel skin friction drag and reaction control rocket plume interaction data have been corrected t o free-flight conditions using Reynolds number scaling for skin friction, and jet-to-freestream-momentum ratio for plume effects.

Criteria for aerodynamic design of the Orbiter have been determined from analyses of the entry flight phase, considering requirements for vehicle trim, control , performance, and aerodynamic heating. A typical entry profile i s i l lustrated in Figure 21. Trajectory guidance i s accomplished by flying an angle of attack/veloci ty profile preselected to meet thermal design cr i ter ia , and using roll comnands for range control. Flight control i s accomplished in two modes termined spacecraft and aircraft . The spacecraft mode applies from iinitial through mid-entry phases where the Orbiter i s a t high angle of attack making the vertical fin and rudder ineffective. The aircraft mode includes mid-entry through approach and 1 anding. Switching from spacecraft to aircraf t modes i s performed as a function of angle of attack and velocity. Transition begins a t approximately Mach 5 and i s completed a t about Mach 1.5. In the spacecraft mode, control in a l l three axes i s ini t ia l ly provided by the a f t reaction control system jets mounted a t the base of the Orbiter on either side of the vertical t a i l . As control authority of the aerodynamic surfaces becomes sufficient, the jets are deactivated. Utilization of the control surfaces and jets during entry i s illustrated in Figure 22. A t a dynamic pressure of 95.8 N / m 2 , the elevons are used to supplement the jets in pitch and rol l . The roll jets are turned off a t a dynamic pressure of 478.8 N/m2 a t which point the yaw jets are used to ini t ia te roll maneuvers (as well as yaw control) with the ailerons providing turn coordination until switchover to the aircraft mode. A t a dynamic pressure of 957.6 N/m2, the pitch jets are turned off and the elevons provide pitch control. Transition to the aircraft mode i s initiated a t approximately Mach 5 when the rudder i s activated. The yaw jets are turned off a t about Mach 1, and the

rudder provides control until landing. The speed brake i s programed t o a s s i s t pitch trim and augment la tera l s t a b i l i t y during t rans i t ion from spacecraft t o a i r c r a f t control. During approach and landing, the speed brake se t t ing i s modulated fo r speed control. Additional pitch trim i s provided by the body f lap which i s programed as a function of velocity.

Significance of the aerodynamic parameters res ts i n t he i r ef fec ts on vehicle performance, control , and airloads. Those parameters most sensit ive to meeting entry mission requirements a re l i s t e d i n Table 5. Li f t , drag, and pitching moment are the primary aerodynamic parameters governing the entry trajectory, range capability, and thermal system design requirements in terms of heat ra te and load. Heating r a t e influences maximum surface temperature and af fec ts material reuse capability. Heat load establishes material thickness to maintain structural temperatures and, therefore, a f fec ts thermal protection system weight. Pitching moment determines the elevon se t t ing required fo r trim. Design areas sensit ive to trim se t t ing are elevon heating during i n i t i a l entry, and control surface actuator s t a l l limits a t transonic speeds. In addition, there i s an interaction between elevon se t t ing and lateral-directional control capability because of the change i n ro l l and yaw effectiveness of the ailerons with elevon position. Lateral-directional trim and control capability i s governed by the ai leron and rudder control derivatives. Above Mach 5 the aileron i s used fo r both ro l l and yaw trim before the rudder become effective. Between Mach 5.0 and 1.5, the rudder provides both yaw and rol l trim w i t h the aileron providing t u r n coordination. Below Mach 1.5, ro l l trim is provided by the rudder. The derivatives Cng. Cag9 Cn6a, C R ~ , , CnSr, Cggr are

key to establishing control capabil i ty, reaction control system propellant usage, and the-switch-over- point from spacecraft t o a i r c ra f t control modes.

High Altitude Aerodynamics

The entry interface, defined as the upper l imi t of the sensible atmosphere, begins a t approximately 120,000 meters a l t i tude . In th i s high a l t i t ude region, say 70,000 to 120,000 meters, rarefied gas flows are encountered by the orbi ter as i t enters the atmosphere. Aerodynamic design issues i n t h i s region involve determining the effectiveness of the control j e t s and t h e i r influence on the Orbiter flow f i e ld , i n addition to defining viscous interaction ef fec ts associated with low Reynolds number/high Mach number f 1 ows .

In i t i a l entry aerodynamic characterist ics, Figure 23, are highly influenced by interactions between the reaction control system j e t plumes and the local flow f i e ld over the Orbiter. The to ta l j e t e f fec ts are comprised of three factors:

r J e t thrust r Surface impingement r Flow f i e ld interaction

Impingement and interaction ef fec ts are inter-related and have been obtained from wind tunnel test ing. Coupling i s present between the plume effec ts and aero surfaces, and between the j e t s themselves.

A ser ies of model nozzles with d i f ferent expansion ra t ios were employed during the wind tunnel t e s t program. General Dynamics/Convair, under contract t o the NASA (NAS9-14095), Reference 11, developed a method whereby the experimentally measured induced plume effec t (surface impingement plus flow f i e ld interaction) could be separated into two component parts and the impingement term extrapolated to f l igh t conditions. To obtain a correct modeling of the reaction control system plume effec ts i n the wind tunnel, i t was necessary to observe certain scaling c r i t e r i a . The primary factors for consideration, aside from dimensional scaling, are plume shape and jet-to-freestream momentum ra t io , @j/@m. In some instances, namely, yaw thrus ter f i r ings , mass flow ra t e r a t io , i j / l i m , scaling was found t o be a s l ight ly bet ter modeling parameter than momentum ra t io . The scaling parameters are defined as:

and

m. @. vm -s= - 1 (sin 13.)"~ = 1.300 x (>)

@a V j n J qm

where

m = J e t mass flow ra te j

im = Freestream mass flow ra t e

v = E x i t velocity - meters/sec j

- = Dynamic pressure - FE/m2 qm 9. = Nozzle half-angle a t ex i t J

n = Number of thrusters

Sref = Reference area (24.9 m2)

In scaling from wind tunnel to f l ight , certain adjustments to the data base are required to account for real exhaust plume effects since cold a i r was used as the je t media in the tunnel. Model plume impinge- ment effects were theoretically extracted from the measured tunnel data and the remaining j e t plume inter- action effects correlated against +j/&,, or hj/im. Prototype impingement effects were then theoretically generated. Examples of the data correlation for pitching moment are presented in Figure 24.

The application of the reaction control system data to a typical entry flight condition of &,, = 478.8 N/m2 a t an altitude of 79,250 meters are presented in Figure 25 for three a f t l e f t downfiring reaction control system thrusters. I t i s to be noted that adverse effects to control authority result from the impingement and flow interaction terms for roll and pitch; whereas, in yaw these terms tend to increase the je t moment.

Viscous interaction effects gre scaled from wind tunnel t es t data to f l ight conditions by means of the hypersonic viscous parameter V A discussed earlier:

where

Mm = Freestream Mach number

%"LB = Freestream Reynolds number (based on body length, LB)

CL = Proportionality factor for the linear viscosity-temperature relationship (Reference 8)

with Monahan's empirical relationship given by

where

T1 = Reference temperature, degrees Kelvin

Tao = Freestream s ta t ic temperature, degrees Kelvin

Tw = Wall temperature, degrees Kelvin

y = Specific heat ratio

K = Empirical constant = 0.5 for a i r

j = Empirical constant = 1.0 for a i r

NOTE: A constant wall temperature of 1367OK and specific heat ratio of 1.15 have been assumed for the flight conditions analyzed.

The primary viscous interaction effects are in shear forces w i t h essentially no effect on normal force. Variation of V; along the nominal entry trajectory i s illustrated in Figure 26. High values of v; correspond to low values of Reynolds number which i s associated with the thickening of the hypersonic laminar boundary layer causing increased shear on the lower supface of the Orbiter. Evidence of this i s seen as an increase in axial force coefficient with increasing VA with no change in normal force, Fig- ure 27. Pitching moment a t zero degree control deflection, Figure 28, becomes slightly more negative with increasing V A due to increased shear forces on the lower surface of the Orbiter. A t negative ( t rai l ing edge-up) control deflections, the movement of the control surface has-lit t le effect on the boundary layer on the lower surface of the Orbiter, and consequently, the effect of VA on ,pitching moment i s similar to the zero degree deflection case. For positive (trailing edge-down) deflections, however, the pitching moment effectiveness of the control surface decreases with increasing VA. A t high VA (corresponding to low Reynolds number) a thickening of the boundary layer results with a separation point which moves forward with increasing control deflection. This causes a net forward movement of the center of pressure, result- ing in reduced pitching moment effectiveness with increasing v;, Figure 28.

Effects of FA on aerodynamic performance characteristics are indicated in Figures 29 and 30 for a nominal entry trajectory. The decrease in lift-to-drag ratio caused by the increase in axial force i s accounted for in design of the entry trajectory. Reduced elevon effectiveness increases the control sur- face deflection required to trim, Figure 30.

Longitudinal Characteristics

Longitudinal s tabi l i ty and control characteristics for hypersonic to low speed Mach numbers are illustrated in Figures 31 and 32. These data have been determined as a result of extensive wind tunnel tests (representative data are shown on the curves) w i t h hypersonic theory being used to bridge the gap between high supersonic data and the hypersonic wind tunnel data. Low-speed longitudinal characteristics shown in Figure 32 demonstrate stall-free characteristics over the operating angle of attack range. The predicted characteristics are compared with tes t data obtained with a 0.36-scale model in the Ames Research Center 40x80-ft (12.19x24.38 m) wind tunnel. The changes in s tabi l i ty evidenced in Figure 32 by the large changes in pitching moment a t high angles of attack are due to leeside separation on the orbiter wing induced by vortices from the wing/fuselage junction.

The leeside flow separation influences the supersonic stabil i ty characteristics a1 so. Referring to Figure 31, i t can be seen that for M = 10 and 5, the variation of pitching moment with normal force coefficient for zero and positive elevon deflection follows the classical Newtonian "sine squared" relation- ship. This relationship between pitching moment and normal force coefficient does not follow the "sine- square" variation for negative elevon deflections. The change in characteristics i s due to the change in flow pattern on the leeside of the Orbiter wing as influenced by negative elevon deflections.

The surface flow patterns on the leeside of the Orbiter wing consist of three distinct flows. A t low angles of attack, the flow which i s ini t ia l ly perpendicular to the leading edge i s turned parallel to the freestream by the presence of the fuselage (Figure 33A). When the angle of attack i s great enough to cause the wing leadin edge shock to detach, the trailing edge shock will become strong enough to separate the boundary layer ?Figure 338). This separation i s the result of subsonic flow a f t of the detached shock expanding around the leading edge and reattaching a t supersonic speeds. The flow must s t i l l be turned into the freestream direction as before. The turning i s accomplished by a strong shock that causes the boundary layer to separate. The wake begins to affect the flow pattern a t higher angles of attack causing a secondary type of separation (Figure 33C). Leeside flow boundaries for M = 6.0 are shown in Figure 34. The relationship between spanwise location of the shock induced separation bs, and Mach number was obtained -

b from a correlation of delta wing data. The shock detach boundary was obtained from oil flow photographs.

The effect of leeside separation on wing pitching moment i s shown in Figure 35. The subsonic leading edge suction that occurs when the bow shock detaches results in a more stable pitching moment slope. The change to a more stable slope i s the result of leading edge suction when the wing bow wave detaches and a reduction of l i f t over the wing area a f t of separation line (Figure 35). The center of pressure i s more a f t for the l i f t gain (due to leading edge suction) than for the l i f t loss due t o shock-induced pressure a f t of the separation line. The wing pitching moment becomes more stable, thus accounting for the increased s tabi l i ty shown in Figure 31 for +lo-degrees elevon deflection.

Elevon effectiveness i s also influenced by leeside separation. Loss in elevon effectiveness a t high negative ( t rai l ing edge up) deflection can be attributed to the effect of back-pressure on the leeside flow field. Flap type controls will often cause boundary layer separation, especially in hypersonic low- density flows. Such back-pressure effects are of practical concern since i t i s desirable to control the Orbiter with leeward control deflection (trailing edge up) in order to minimize control surface heating. Figure 36 shows elevon effectiveness data obtained from the AEDC Tunnel A a t M = 5 for an elevon deflec- tion -35 degrees. The measured elevon effectiveness i s seen to be less than shown by shock expansion theory. This i s probably due to shock-induced separation. The separation extent increases with angle of attack. After the angle of attack for shock detachment i s reached, the back-pressure effect from the elevon will affect the wing flow. A t high angles of attack, the positive l i f t produced by the wing vortices outweighs the negative l i f t generated by the elevon-induced flow separation over the inner wing surface. The result i s a loss of elevon effectiveness below the shock expansion value. Adjusting the theory for leeside separation results in reasonable agreement between theory and experiment.

Static trim capability for the elevon and body flap positioned for trim to the forward and a f t center of gravity positions i s displayed in Figure 37. The control schedules presented on the figure are for determining maximum obtainable center of gravity trim limits. A reserve for maneuvering, trimming span- wise center of gravity offset, manufacturing misal ignments , and aerodynamic uncertainties has been added to the limits of the elevon effectiveness data to establish the limits shown on the figure. The a f t center of gravity 1 imits are based on a positive elevon deflection of 15 degrees for Mach numbers less than or equal to ten. A positive elevon deflection of ten degrees was used for Mach numbers greater than ten due to thermal protection system design limits during maximum heating conditions. Forward center of gravity trim limits are based on an incremental pitching moment coefficient reserve of 0.015 for Mach numbers less than or equal to ten and 0.02 for Mach numbers greater than ten. Figure 37 indicates a slightly reduced forward center of gravity trim margin a t Mach 5.0 in the angle of attack range rrom 20 to 45 degrees. This i s attributed to the loss in elevon effectiveness due to leeside separation. Center of gravity trim limits for the entry angle of attack schedule have been shown earlier in Figure 7. Both Figures 7 and 37 indicate that a wide trim margin exists across the Mach number range.

Elevon control power in conjunction with the body flap and speed brake provide trim capbility between the design center of gravity limits. The elevon schedule, shown in Figure 38, i l lustrates the nominal and the most positive and negative settings for trim a t forward and a f t center of gravity positions. The extreme settings account for control margin and uncertainties in aerodynamic characteristics. The speed brake i s ini t ia l ly opened during entry a t Mach 10, and i s programed as a function of velocity to

approximately Mach 1. Opening the speed brake a t Mach 10 assists in longitudinal trim during the transi- tion from high to low angle of attack. Below Mach 1 , the speed brake setting i s modulated to provide speed control during approach and landing. The body flap i s used as a trim device to keep the elevons operating in an effective range, and, during the high heating portion of entry, to keep the elevons from overheating. Body flap deflection varies during entry. For entry a t the forward center of gravity, the ini t ia l body flap position i s normally full-up. For the a f t center of gravity case, the ini t ia l position i s approximately 16 degrees down.

Lateral-Directional Characteristics

Lateral-directional s tabi l i ty and control characteristics along the nominal entry trajectory are illustrated in Figures 39, 40, and 41. As shown in Figure 39, the Orbiter exhibits positive dihedral effect (negative Cgg) across the complete Mach range during both the spacecraft and aircraft control modes. During the spacecraft mode, and during transition to the aircraf t mode, the directional s tabi l i ty deriva- tive Cn8 i s negative. Cns becomes positive indicating s ta t ic s tabi l i ty in yaw a t approximately Mach 1.7, '

and retains positive values throughout the aircraf t mode (Mach numbers below approximately 1.5). Aileron and rudder control effectiveness characteristics are il lustrated in Figures 40 and 41. Because the ailerons provide control authority across the complete Mach range, and the rudder i s essentially ineffective above Mach 8, the ailerons are used in conjunction with the yaw jets to provide for roll control in the space- craft control mode.

Early analytical studies predicted an elevon deflection interaction effect on the lateral-directional characteristics. Studies showed that the relatively large sized elevon in the presence of the deep, f la t - sided fuselage could induce a change in the pressure distribution in the a f t region of the fuselage. The resulting change in pressure distribution resulted in an incremental change in side force, yaw, and roll- ing moment when the vehicle was yawed. The effect of elevon interaction i s illustrated for the yawing and rolling moment derivatives in Figures 42 and 43. The control derivatives Cgda and Cn6 are also affected

a

by elevon position. The influence of elevon position on the control derivatites i s shiwn in Figures 44 and 45. The sensitivity of the derivatives to elevon position influences vehicle control boundaries.

The nature of the control derivatives define the Mach regions where aileron and/or rudder i s used for lateral trim. Aileron-alone i s used for YCG trim above Mach number 4.5 and rudder-alone i s used for YCG trim below Mach number 3.5. A combination of aileron and rudder control i s used for trim in the Mach number region between 4.5 and 3.5. The yaw reaction control system (RCS) jets are used to augment the aerodynamic controls where required.

The interrelation between the control derivatives and the method used to trim YCG offset during the spacecraft mode i s best i l lustrated by examining the relations for aileron and rudder required to trim. Using aileron-alone:

and for rudder-alone:

The critical boundary exists when the denominator goes to zero; i .e., the condition where aileron or rudder cannot produce a trim condition.

For aileron-alone, the boundary i s defined by

Using rudder-alone :

Aileron and rudder cross coupling ratios are shown in Figure 46 for the nominal entry trajectory and nominal aerodynamics. For the spacecraft control mode, the boundary for YCG trim by aileron control i s not violated for Mach numbers greater than about 1.9. Trim of a YCG offset by rudder-alone can be accomplished over the region from Mach 8 to approximately 1 . I , resulting in an overlap from Mach 8 to 1.9 for nominal

aerodynamics. To allow fo r trajectory dispersions and uncertainties in aerodynamics, t rans i t ion to the a i r c ra f t control mode; i .e., conventional aileron/rudder control, i s i n i t i a t ed a t approximately Mach 5 and is complete by Mach 1.5.

Low-speed directional s t a b i l i t y characterist ics exhibit a strong Reynolds number/angle of at tack ef fec t , Figure 47. The figure i l l u s t r a t e s the importance of full-scale Reynolds number tes t ing on high angle of at tack aerodynamics. Test data obtained from models tested a t low Reynolds numbers (5 5 x lo6 based on mean aerodynamic chord) show essential ly no change of directional s t a b i l i t y with angle of attack. The early work of Polhamus (Reference 12) and Jorgensen and Brownson (Reference 13) indicated that Reynolds number and body corner radius could have a significant ef fec t on the h i g h angle of attack characterist ics of the Orbiter. These predictions were borne out when the Orbiter model was tested a t near full-scale Reynolds number in the Ames Research Center 40x80-foot (12.2x24.4 m) wind tunnel . Referring to Figure 47, i t can be seen t h a t the high Reynolds number t e s t data shows a decrease in directional s t a b i l i t y with angle of at tack which i s in contrast to the low Reynolds number data which shows essential ly no change in s t ab i l i t y w i t h angle of attack. A t Mach numbers above 0.7, data representative of f l i g h t Reynolds numbers can be obtained in wind tunnel t e s t s a t low Reynolds numbers provided proper at tention is paid to close matching of body corner rounding on the wind tunnel models and f l igh t vehicles (Reference 13).

The early work of E.C. Polhamus (Reference 12) was used to predict the variation of directional sta- b i l i t y with angle of attack. Based on Polhamus' work, i t was determined that the low-speed/high angle of at tack directional s t a b i l i t y determined by wind tunnel t e s t s would be erroneous unless the Reynolds number were sufficiently high to permit proper simulation of the cross flow on the forward fuselage. I t was predicted tha t a close s imi lar i ty i n both magnitude and change with Reynolds number, existed between the cross flow drag coefficient for the Orbiter fuselage a t high angles of at tack (greater than 15 degrees) and a two-dimensional square cylinder a t 90 degrees angle of at tack. From t h i s s imi lar i ty , i t was con- cluded tha t most of the low-speed Orbiter t e s t data would be within the c r i t i ca l Reynolds number range, the range in which cross flow drag coefficient decreases from h i g h to low values as the Reynolds number increases from subcrit ical t o c r i t i c a l . Polhamus presented data from t e s t s made on noncircular cylinders w i t h the a i r flow directed normal to the cylinder axis. As i l l u s t r a t ed i n Figure 48 for a square-shaped cylinder w i t h rounded corners, the a i r flow will separate on the leeside a t subcri t ical Reynolds number b u t will remain attached a t supercri t ical Reynolds number, when the flow i s directed a t an angle not aligned with one of the major cross sectional axes. For the subcri t ical Reynolds number case, the r e su l t - ant body axis side force, Cy, i s positive while for the supercri t ical Reynolds number case, the s ide force i s negative. Since the center of gravity i s behind the nose, positive side force translates to positive yawing moment and negative s ide force t rans la tes to negative yawing moment. Consequently, the ef fec t of going from subcrit ical t o supercri t ical Reynolds number is t o reduce the directional s t a b i l i t y of the vehicle.

Presented i n Figure 49 i s the measured directional s t a b i l i t y variation w i t h cross flow Reynolds number from several wind tunnel t e s t s a t approximately 20 degrees angle of attack. As can be seen, there appears t o be a trend for the Orbiter directional s t a b i l i t y to decrease as the cross flow Reynolds number i s increased. The reason for the reduction of directional s t ab i l i t y with increased cross flow Reynolds number is the elimination of the flow separation a t the nose w i t h increased cross flow Reynolds number. A t angles of at tack of 15 degrees and below (Figure 47), there appears t o be no change of directional s ta- b i l i t y fo r the different t e s t s .

FLIGHT TEST RESULTS

The Approach and Landing Test (ALT) Program, Table 6, was conducted during the l a s t half of 1977 as part of the Shutt le Development Program. The Orbiter Enterprise (OV101) was launched from the Boeing 747 Shuttle Carrier Aircraft over Edwards Air Force Base, California, and glided to e i ther a landing on Rodgers Dry Lake, or on the l a s t f l i g h t , a landing on the Edwards Air Force Base runway. The program con- consisted of eight captive f l igh t s followed by three o rb i t e r f reef l ights w i t h the tailcone ins ta l led , and f inal ly two f reef l ights in which the tailcone was removed, tes t ing the orbital return configuration.

The captive f l igh t s verif ied the airworthiness of the mated configuration, accomplished Orbiter sys- tems checkout and developed the separation procedures, and verified aerodynamic forces a t separation. The separation of the Orbiter from the 747 was achieved through the aerodynamic forces on the vehicles, so one of the important objectives of the Captive Flight Program was t o verify the predicted separation forces, and adjust the Orbiter elevon sett ings fo r separation, i f required. Special load ce l l s were instal led on the Orbiter/747 s t r u t s in order to measure the separation forces t o the required accuracy.

The orbi ter l i f t and pitching moments in the presence of the 747 were the key parameters for safe separation. Figure 50 shows these coefficients as determined during the captive f l igh t s using load ce l l measurements, compared to estimates based on wind tunnel t e s t s . The f l i g h t measured coefficients were we1 1 w i t h i n the uncertainties in the prediction. Since dynamic analyses had shown acceptable separations i f the key aero coefficients were within the uncertainty band, the f l igh t measurements confirmed that separa- t ion would be acceptable, and the program proceeded to an Orbiter f reef l ight . The f i r s t separation occurred on August 12, 1977, and was as predicted.

The f i r s t three Orbiter f reef l ights were conducted with a tailcone instal led to f a i r the Orbiter 's blunt base. The tailcone provided increased glide range by increasing the Orbiter L /D from a maximum of 4.5 t o 7.5. I t also allowed increased launch a l t i tude , from 6400 m t o 7620 m. The tailcone also was intended t o reduce buffet levels a t the 747 empennage, since there was some concern tha t w i t h the Orbiter tailcone-off, the Orbiter wake may induce excessive buffeting and reduce the 747 fatigue 1 i f e . (Flight measurements of f luctuating structural loads in the 747 t a i l l a t e r re1 ieved t h i s concern.)

After th ree ta i lcone-on f r e e f l i g h t s , the t a i l c o n e was removed and the t e s t program completed w i t h the o r b i t a l con f igu ra t ion simulated. Two O r b i t e r f r e e f l i g h t s were accomplished w i t h the t a i l c o n e removed. These f l i g h t s obtained the data used t o v e r i f y the subsonic aerodynamic c h a r a c t e r i s t i c s o f the O r b i t e r i n i t s o r b i t a l r e t u r n con f igu ra t ion . Separation o f the orb i ter1747 dur ing t h e f i r s t t a i l c o n e - o f f f r e e f l i g h t i s i l l u s t r a t e d i n F igure 50. I n F igure 51, the o r b i t e r i s shown j u s t p r i o r t o touchdown on the runway a t Edwards A i r Force Base dur ing the f i n a l f r e e f l i g h t on October 26, 1977.

The s i g n i f i c a n t O r b i t e r aerodynamic c h a r a c t e r i s t i c s are compared w i t h p red ic t ions based on wind tunnel tes ts , Figures 52 through 56. Key inst rumentat ion used t o der i ve these data were an "Aerodynamic Coe f f i - c i e n t I d e n t i f i c a t i o n Package" provided by the NASAIDryden F l i g h t Research Center, cons is t ing o f th ree accelerometers and th ree r a t e gyros, and a f l i g h t t e s t noseboom which provided angles o f a t tack and side- s l i p , and p i t o t and s t a t i c pressures. A data e x t r a c t i o n pro ram was developed by the NASA t o determine the f l i g h t - d e r i v e d aerodynamic c h a r a c t e r i s t i c s (Reference 147. The f i g u r e s show t h a t t h e f l i g h t measure- ments were i n good agreement w i t h wind tunnel p red ic t ions , w i t h the lone exception o f l and ing gear drag which was overpredic ted by approximately 27 percent. Examination o f the wind tunnel data revealed t h a t the est imated Reynolds number c o r r e c t i o n t o gear drag was inadequate. Thus, the aerodynamics o b j e c t i v e o f the Approach and Landing Test Program; t o v e r i f y the low-speed aerodynamic c h a r a c t e r i s t i c s , was achieved and prov ided increased confidence f o r the nex t phase o f t h e S h u t t l e Program, the O r b i t a l F l i g h t Test.

CONCLUSIONS

Aerodynamic development o f t h e Space S h u t t l e o r b i t e r has been described. Extensive wind tunnel t e s t i n g has provided a h igh confidence 1 eve1 i n the est imated aerodynamic c h a r a c t e r i s t i c s . Results from the approach and land ing f l i g h t t e s t program v e r i f y p red ic ted aerodynamic c h a r a c t e r i s t i c s i n t h e subsonic speed regime. Accounting f o r uncer ta in t ies i n aerodynamic data w i l l a1 low incremental extension o f f l i g h t envelopes t o achieve p red ic ted operat ional c a p a b i l i t y .

REFERENCES

1. Surber, T.E. and Olsen, D.C., "Space S h u t t l e O r b i t e r Aerodynamic Development," Journal o f Spacecraft and Rockets, Vo1 . 15, No. 1, January-February 1978, pp 40-47

2. Rockwell I n t e r n a t i o n a l Space D iv is ion , "Aerodynamic Design Data Book, Volume 1 , O r b i t e r Vehiicle," November 1977, Report No. SD 72-SH-0060-1K

3. Smith, E.P., Rockwell I n t e r n a t i o n a l Space D iv is ion , "Space S h u t t l e O r b i t e r and Subsystems ,I1 June 1973, Report No. SD 72-SH-0144

4. Gamble, Joe D., Chrysler Corporat ion Space D iv is ion , "High Supersonic S t a b i l i t y and Control Character- i s t i c s o f a 0.015-Scale (Remotely Con t ro l led Elevon) Model 49-0 o f the Space S h u t t l e O r b i t e r Tested i n t h e NASAILaRC &Foot UPWT (LA63B)," May 1976, Report No. DMS DR-2279, NASA CR-144.606

NASA LaRC, "High Reynolds Number Research," October 1976, Report No. NASA CP-2009, pp 2-17

Hayes, Wallace D., and Probstein, Ronald F., "Hypersonic Flow Theory," New York and London, Academic Press, 1959, pp 333-345

Whi t f i e ld , Jack D., and G r i f f i t h , B.J.. , "Hypersonic Viscous Drag E f fec ts on B lun t Slender Cones," AIAA Journal, Vo1 . 2, No. 10, October 1964, pp 1714-1722

Bertram, M i t c h e l l H., NASA, "Hypersonic Laminar Viscous I n t e r a c t i o n E f fec ts on t h e Aerodynamics o f Two-Dimensional Wedge and Tr iangu la r Planform Wings," August 1966, Report No. NASA TN D-3523

Rockwell I n t e r n a t i o n a l Space D iv is ion , "Aerodynamic Design Substant ia t ion Report, Volume 1, O r b i t e r Vehicle," February 1978, Report No. SD 74-SH-0206-1 K

Stein, Robert E., Jr. , Lockheed E lec t ron ics Inc., "Pro ject 331 3, Mu1 t i p l e Regression Analysis Program f o r t h e Aerodynamic C o e f f i c i e n t Analysis System (NASA Contract NA9-122000)," January 1973

Rausch, J. R., General Dynamics Convair D iv i s ion , "Space Shu t t le O r b i t e r Rear Mounted Reaction Control System J e t I n t e r a c t i o n Study," May 1977, Report No. CASD-NSC-77-003

Pol hamus, E. C. , "Effect o f Flow Incidence and Reynolds Number on Low Speed Aerodynamic Charac te r i s t i cs of Several Nonci rcu lar Cyl inders w i t h App l i ca t ion t o D i rec t iona l S t a b i l i t y and Spinning," NASA Technical Report R-29, 1959

8

Jorgenson, Leland H. and Brownson, Jack J., " E f f e c t o f Reynolds Number and Body Corner Radius on Aerodynamic Charac te r i s t i cs f o r Space Shuttle-Type Vehicle a t Subsonic Mach Numbers," NASA TN D-6615, January 1972

14. Romere, P.O., E ichb la t t , D.L., underwood, J.M., and Howes, D.B., "The Space S h u t t l e O r b i t e r Approach and Landing Tests-A Cor re la t ion o f F l i g h t and Predic ted Performance Data," AIAA Paper 78-793, AIAA Tenth Aerodynamic Test ing Conference, San Diego, CAY A p r i l 19-21, 1978

ACKNOWLEDGEMENTS

The aerodynamic design ac t iv i t i e s described in t h i s report are being accomplished by the Space Division, Rockwell International, under NASA contract NAS9-14000. Overall technical management of orbi ter develop- ment i s accomplished by the Space Shuttle Program Office a t the NASA Johnson Space Center (JSC). Aero- dynamic development i s directed by JSC through the Space Shuttle Flight Performance Technical Management area, Mr. Bass Redd, Technical Manager, supported by Mr. James C. Young, Aerodynamic Subsystem Manager. In addition t o JSC, the other NASA Centers (Ames Research Center, Langley Research Center, and Marshall Space Flight Center) have provided extensive support t o the Shuttle development i n accomplishment of the Wind Tunnel Program and i n analysis of t e s t results . In addition, the Langley Research Center has provided additional program support through the i r independent investigations of nonl inear aerodynamics.

A t Rockwell, s ignificant contributions t o th i s paper have been made by Messrs. D.C. Olsen, H.S. Dresser, W.R. Russell and L.M. Gaines.

Table 1 BASELINE REFERENCE MISSIONS

ORBIT LAUNCH INCL. ALTfTUDE VJRATIO!

MISSION SITE OBJECTIVE (DEG) (10 n) (DAYS)

1 K C PAYLOAD DELIVERY 28.5 277.8 1 7

Table 4 SPACE SHUTTLE WIND TUNNEL MODEL SUMMARY AUGUST 1972 TO FIRST ORBITAL FLIGHT

PAYLOAD ( l o 3 ~ g )

ASCENT DESCENT

29.48 14.51

AERODYNAMICS

AERODYNAMIC HEATING

STRUCNRAL DYNAMICS

TOTAL

SPACELAB

Tabl e 2 AERODYNAMIC DESIGN REQUIREMENTS

NUMBER OF MODELS

HYPERSONIC 25 DEG TO 5 0 DEG TRPJ6ONIC 0 DEG TO 1 5 DEG SUBSONIC -5 DEG TO 2 0 DEG

Tabl e 5 SIGN1 FICANT AERODYNAMIC PARAMETERS

ORBITER VEHICLE

11

22

12

45

MATED LAUNCH VEHICLE

23

14

6

43

CARRIER AIRCRAFT

4

0

2

6

L I D

REAL GAS EFFECTS NOT

TOTAL

38

36

20

94

AEROOYWIC PARAMETERS

CENTER OF GRAVITY RANGE

MINIMUM T W E L

DESIGN RANGE 2% BODY LENGTH

0.65 LB - 0.675 LB

FLIGHT REGl lE

. DESIGN HINGE W E N T CONOI- . TRANSONIC MIND TUNNEL 1 IF!: TO I TIONS . DEFINES CONTROL SURFACE STALL DATA ACCURACIES

CONDITIONS LANDING PERFORMRNCE

PAYLOAD

LANDING WEIGHT (MITH PAYLOAD)

MINIMUM DESIGN TOUCHDOWI~ SPEED, VD

YHY PAPAMETERS ARE SIGNIFICANT

EUCH 5 TO 6 . AILERON USED FOR BOTH ROLL a . m t i T w L SURFACE INTER- YAY TRIM A W E NdCH S BEFORE ACTIOH RUDDER BEUmE EFFECTIVE

YAY JET FUEL USAGE ( W E TO CG I cn6' / 1 OFFSET) 1 I

AERO CONCERN I N DEFINITION OF PAPAMETERS

LONGITUDINAL STABILITY

MINIMUM HYPERSONIC STATIC MARGIN

MINIMUM SUBSONIC STATIC MARGIN (RFT CENTER OF GRAVITY)

, C , Cng . C HIGH SUPER- . BENEEN EUCH 1.5 TO 5. . CONTROL SURFACE INTER- 6 6 , C 5 H WER S USED FOR W. YAW ACTIONS 1 6 ROLL TRIM. AILERON 03- . RUDDER EFFECTIVEHESS AT

SUPEWTRAN- ORDINdTES TURN HIGH a, HACH SONIC NdUI . YAU JET NEEDED UNTIL RUDDER . AEROEWTIC EFFECTS 2.5 TO 1.0 I S EFFECTIVE . TRANSONIC WIND TUNNEL . DEFINES SYITCH-OVER POINT DATA ACWRAClES

POSITIVE

-2% LB (-5.45% MAC)

LIFTIDRAG EaDULATION

PEAK SUBSONIC VALUE (GEAR-UP. 655 = 0 )

PEAK SUBSONIC VALUE (GEAR-UP. 655 = 8 5 DEG NOT LESS THAN 4.4

NOT LESS THAN 2.5 Table 6 APPROACH & LANDING TEST PROGRAM SUMMARY

Table 3 SPACE SHUTTLE WIND TUNNEL TEST HOURS SUMMARY AUGUST 1972 TO FIRST ORBITAL FLIGHT

CAPTIVE FLIGHT

747 6 ORBITER, TAILWNE-ON, ORBITER UNHANNEO L UNPOUEREO

FLUTTER CLEQANCE, PERFOWCE, 6 STA- BILITY L mNmoL VERIFICATION

ORBITER FUNCTIONAL CHECK, SEPAlTlON LORDS, L PROCEDURES VERIFICATION

L4RRlER ALONE

AERODYNAMICS 14,700 8.100 3,400 I AERODYNAMIC I 4,500 1 6,400 1 0 HEATING

NO. OF FLIGHTS

4

7

I TAlLCOhE-ON CRM FAHILIARIIATION, SYSTW CHECK.

STABILITY 6 CONTROL. L PERFORMNEE

TAILCONE-OFF VERIFICATION

VERIFICATION OF ORBITER APPROACH L LANO- ING CAPABILITY

ORBITER VEHICLE

STRUCTURAL I OYNAMICS 1 lsooO 1 1'600 ( 400

CONFIGUPATION

MODIFIED 747

MODIFIED 747 WITH ORBITER ATTACH STRUCTS 6 TIP FIMS

CAPTIVE IllERT

4 FERRY WIIFIGURATION. TAIL- FERRY QU~LIFICATION L PERFORMPNCE VERI- CONE-OH FICATION

OBJECTIVE

FUNCTIONAL CHECK, FLUTTER, L STALL CHECKS

FLUTTER CLEAPANCE, PERFORMANCE, b STA- BILITY 1 CONTROL VERIFICATlON

LAUNCH VEHICLE

I TOTAL 1 2 0 , 2 0 0 1 1 6 , 1 0 0 1 3 , 8 0 0 1 4 0 , 1 0 0 1

SOLID ROCKET BWSTER (SRB) EXTERNAL 3.70 DlA

CARRIER AIRCRAFT

' g& vmr.sms. ( l a z ) 249.909 28.192

bSPECT RATIO 2.265 1.615 SHEEP (DEG)

LEADING EDGE 45 45 GLOVE 81

n.A.c. (.I DlHEMAL (DEG)

12.060 3.043 3.5

TOTAL HOURS

CONTROL SURFACE AREA 6 IUXINJU DEFLECTION ARM (m') DEFLECTION (DEG)

ELEWN (WE SIOE) 19.509 -35 TO +20 RUWER 10.233 r22.8 SPEED BRAE 10.233 0 TO 87.2 BODY F U P 12.541 -11.7 m r22.5

SRB THRUST ATTACH

I ORBITER

TANKIORBITER AFT ATTACH

I \

F i g . 1 SPACE SHUTTLE VEHICLE Fig. 2 ORBITER VEHICLE

F i g . 3 THERMAL PROTECTION SUBSYSTEM F i g . 4 SHUTTLE MISSION PROFILE

F i g .

TIME FRon EMTRY (SECONDS)

5 ORBITER ENTRY TRAJECTORY

CENTER OF

GRAVITY LOCATION

W E L I N E UHOIHG YEIGm I14.Sl5 Kg PAYLMOJ

WlIWJI IAYOIHG I .- .A

I (No PAYLOAD) ,i*4&HzPEi;G SPEED

; 5.4 W E C TAILWINO]

I

, , , , i , , 60 7 0 80 9 0 1 W 110

ORBITER LANDING YEIGHl (1WO Kg)

F i g . 6 LANDING

rCONDITIONSAI) CONTROL LIMITS

WIT M Y AT SEA LEVEL 14,515 Kg PAYLOAD 5.1 d S E C TAlLYlNO

34 - ORBITER YET GRWVEO

MIY KUW ih OISPLRSIONI

U - (W YINOI L DISPERSION

70 80 90 100 110 120 T W U M VELOCITY (WSEC TRUE GROUND SPEED1

REQUIREMENTS

0.57 FWD CONTROL LIMIT

0.58

0.61

0.63

0.65

0.67

MACH NUMBER

F i g . 7 ORBITER TRIM L I M I T S

Up r Surlare: 0.1397f 0.0508 T P ~ G ~ ~ Edge Radius: 0.1524 ( M a i i a u m J

Lower WlnglExcepl E levonl

ALL D I M E N S I O N S IN CENTIMETERS Lowor Fuseloge Y/LS 0.3

F i g . 8 ORBITER MOLD LINE/STEP CRITERIA

SYSTEMS REQUIREMENT REVIEW

ORBITAL FLlGHT POR

OR0 101 ASSEMOLY 6 ROLL-OUT

FIRST CAPTIVE FLiGHT

APPROACH L LANOIllG TLST

ORB 102 ASSEMBLI L ROLL-OUT

FIRST MANNED ORBITAL FLIGHT

F i g . 9 SPACE SHUTTLE PROGRAM MILESTONES

I = ' AEOC A I m

2 los; I

L - - - 2

WDEL

ARC 9x7 0.05 AEOC A 0.02

TRRJECTORY

AEDC B

10' MOEL FACILITY SUUI

WOfL I SIM FACILITY SCALE

o 0.5 I 1.5 2 2.5 2 4 6 8 1 0 10 20 a I M NWER I C H RUlER WH NUMBER

F i g . 1 0 ORBITER FLIGHT REYNOLDS NUMBER SIMULATION

F i g . 11 ORBITER AERODYNAMIC WIND TUNNEL PROGRAM F i g . 1 2 OV102 MODEL FIDELITY

0 a-w I 1.. . . .1 1 . 6.0

F i g . 13 L I F T COEFFICIENT TOLERANCES

0 0.2 0.4 0.6 0.8 0.9

C N w t ~ ~ ~ ~ ~ ~ c i

F i g . 14 NORMAL FORCE COEFFICIENT COMPARISON OF WIND TUNNEL TO FLIGHT DATA

F i g . 15 ORBITER EVOLUTION SUMMARY

,_( MRGIN~ND 0.5% L STATIC CG

0 . 3 W Ylblt BLENDED DELTA 0 . 4 2 5 y /

WUBLE DELTA I ' I- .REDUCE GLOVE LEADINGEDGE

RADIUS TO 15.24 sn HORlUL TO GLOVE LEADING EDGE

.HOVE GLOVE-TO-WIN6 INIERSECTION I

FS 8 0 0 FS IWO TO 0.425 SEMI-SPAN Y l T H S W L FAIRING AT InTERSECTlMl

. N I S I 6 UUIBER FM( IWROVEO BASIC LOID

F i g . 16 BLENDEDIDOUBLE-DELTA WING MODI FICATION

"NO CG

F i g . 17 BLENDEOIDOUBLE-DELTA WING MODI FICATION EFFECT ON LONGITUDINAL CHARACTERISTICS

PDR VEHICLE CUR VEHICLE PDR VEHICLE CDR VEHlCLT

. I W 1 NOSE USUlED N . M L L N O Y UDlUI 6 FAIUD

1AIIY WINKHA ADILOW NOY naon~s ~UNSI~WN a IIKX wnuw 6 #€DUCTS SWFAQ liM?

F i g . 18 REDUCED NOSE RADIUS. AND REFAIRED NOSE SECTION

PDR VEHICLE

i i W ?OW1

CDR VEHICLE

F i g . 2 0 UPPER BODY LINES

F i g . 2 2 RCS AND CONTROL SURFACE UTIL IZATION DURING ENTRY

RCS EFFECTIVENESS COMPONENTS

JET THRUST PLUME IMPINGEMENT FLOW F IELD INTERACTIONS

F

F i g . 2 3 REACTION CONTROL JET INTERACTIONS

. A w l o m FArnNG USUlS w LV€W C U M 6 INCUAYO IWAUILDS

F i g . 19 LOWER WINGIBODY FAIRING

ANGLE OF ATTACK

- - - - - 0 1 2 3 5 6 7 8

REUTlYE VELOCITY. 10W WSEC

F i g . 21 TYPICAL ORBITER ENTRY PROFILE

NOZZLE TEST SYM NUMBER NLDBER FACILITY

OA-82.M-22 LaRC CFHT OA-82,w2 OA-82.MAZ2

MA-22 MA-22 OA-169 AEDC B OA-169

0.04

ii I-

E =0.02

0 POINTS = 150 STO. OEY. = 0.002368

0 0.02 0.04 0.06 0.08 0.10 0.12 mmENTUM RATIO $j/C

NO7ZLE TEST SYM NUEBER NUlgER FACILITY

OA-82.MA-22 LaRC CFHT MA-22

-- OA-169 AEDC B OA-169 OA-169

0 0.02 0.04 0.06

MASS FLOY RATIO hj/mm

F i g . 2 4 REACTION CONTROL JET CORRELATION

a = 20' 6 = O0 6E = 68F = OD 9, = 47.9 N/m2

CONTROL 3-JET IMPINGEMENT INTERACTION NET AXIS WMENT (N-rn) MOMENT (N-m) MOMENT (N-m) MOMENT (N-m)

ROLL t36.590 -8:030 -27,040 t1.520

PITCH -126,600 t8.690 t43,775 -74,135

YAW -38,300 -7,860 -4,860 -51,020

VACUUN THRUST FOR ONE JET (N)

Nj = 3570 Aj = -756 Y j = 1298

F i g . 25 REACTION CONTROL JET MOMENTS

SnW. NT P. FKlLIN or-n r n r v l v s

E-B 0: Y

t; $ 0.010 2 SYH FACILITY 'A = S "2 .. > MSFC - 14"TWT

ARC - 3.5'HWT

1 AEDC - FWT

0.W1

4 6 8 10 12 14 16 18 20 22 24 26 28 30 HRCH NUWER

F i g . 26 VARIATION OF VISCOUS PARAMETER ALONG NOMINAL ENTRY TRAJECTORY

F i g . 27 EFFECTS OF VISCOUS INTERACTION ON NORMAL AND AXIAL FORCE

0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08

VISCWS P M T E R vL

F i g . 28 EFFECTS OF VISCOUS INTERACTION PITCHING MOMENT

TPS DESIGN TRAJECTORY 14414.1

WITHOUT VISCWS INTERACTION AND LOW DENSITY CONSIDERATIONS

\------------

ACTION REGIME

0 0 100 200

ALTITUDE - 10' METERS

F i g . 29 VISCOUS INTERACTION EFFECT ON L I FT-TO-DRAG RATIO

Fig. 3 0 VISCOUS INTERACTION EFFECT ON CG TRIM CAPABILITY

SYN FACILITY

0 ARC 11x11

Fig. 32 LOW-SPEED LONGITUDINAL CHARACTERISTICS

PITCHING WMENT C PITCHING MOMENT C % . 6 5 ~ , % . 6 5 ~ ,

Fig. 31 LONGITUDINAL CHARACTERISTICS SUMMARY

SHOCK INWCED

Fig. 33 LEESIDE FLOW PATTERNS

ATTACHED FLW SHOCK INWCED LEADING EDGE SEPARATION SEPARATION

H - 6.0

- ANGLE OF ATTACK (DEG)

Fig. 3 4 LEESIDE FLOW BOUNDARIES

I I 1 I

' I ' I b. g K 5 % T A U I E D SHOCK DRAUIED

MUI 8.0

\L‘olO\, S W IUM FACILITY

C.P. L W I N G E m

M H T c w m

AEDC A 6, * -35- lUCH 5.0

0 4 8 1 2 16 20 24 28 32 ANGLE OF ATTACK (DE6)

F i g . 35 LEESI'DE FLOW SEPARATION EFFECT ON PITCH STABILITY F i g . 36 HYPERSONIC CONTROL EFFECTIVENESS

"L I I I 0.5d I

0.40 0.d. O d l 0.72 016

CENTER OF GRAVITY. XCG/LB

F i g . 37 TRIM CAPABILITY

FORWARD CG

per deg

Cl, -0.001 - per dcg

1 I , * a I I I , , 1

0.2 0.4 0.6 0.8 1.0 2 4 6 8 1 0 20 MACH NUMBER

F i g . 38 ELEVON DEFLECTION SCHEDULE F i g . 39 LATERAL-DIRECTIONAL STABILITY DERIVATIVES

0.W2- - LW)IIIl!N o SUIEWLE F O M M D CENTER OF GWLVlTY . 6. - 0 OEG

C

F i g . 4 0 AILERON EFFECTIVENESS

. NOMINAL SOIEWLE 6 - O O E G

1

F i g . 41 RUDDER EFFECTIVENESS

Fig. 42 EFFECT OF ELEVON DEFLECTION ON CIRECTIONAL STABILITY

SYM TEST NO. El

M * 0.6 SVM TEST NO. o * !iO M - 3.0

AC

(PER OEG)

-30 -20 -10 0 t10 6.- OEG

AC::. t' /-"' + (PER OEG)

-0.002

Fig. 43 EFFECT OF ELEVON DEFLECTION ON ROLL DUE TO SIDESLIP

-

, ACn6 0 (PER MG)

-0.001

-0. 002

- M = 3.0

a - 15'

- , -I. I I

-

-

SYM 1 TEST NO.

M = 0.6 + 1 OA-145 0.005- -

a = 5O

0.004 -

0.003 - C

(PER DEG)

0 s ' ' ' ' ' ' -30 -20 -10 0 t10

C I I I

- AERO MIA BOOK

t' 4

6, - DEG 6,- MG

I 1 I

- , I 1

Fig. 44 EFFECT OF ELEVON DEFLECTION ON AILERON ROLL DERIVATIVE

-40 030 -20 -10 0 t10 t20 -30 -20 -10 0 +lo +20 6,- OEG 6, - OEG

M = 3.0

Cn 6a N.OOl~,aT cn;yll

(PER DEG) (PER DEG)'

-0.001 -0.001 I l l . l . L ' I I

-30 -20 -10 0 +lo -30 -20 -10 0 t10 6,- OEG 6, -- OEG

Fig. 45 EFFECT OF ELEVON DEFLECTION ON AILERON YAW DERIVATIVE

Fig. 46 CROSS COUPLING RATIOS

'I MI2 NWER - 0.5. q - 9580 Nlm'

RIGHT I U S W N T

0 R I M 4

A RIM 5

I UNCERTAIW

Fig. 52 COMPARISON OF FLIGHT VS. PREDICTED DATA FOR ORBITER MATED TO 747 CARRIER

I :' - - - TOLERANCE

, C3

3 g 4 -

3 C: 3 ' F k -1 2 -

VARIATION

1 I I

FREE FLIGHT 4

A FREE RIGHT 5

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

TRIMD LIFT COEFFICIENT CL TRIH

Fig. 53 COMPARISON OF FLIGHT VS. PREDICTED L/D

DYHAnIC PRESSURE - 9580 N h ' 0.7 -

0.6 - FREE FLIMT 4

U

& 0.5 - A FREE RIGHT 5

W

8 0.4 -

-.- VARIATION

0 2 4 6 8 10 12 14

ANGLE OF ATTACK (KG)

/--

PREDICTED - - - -_

, - R A FREE FLIGHT 5 ---- -/-

TOLERANCE

- - - VARIATION

0 I I I I I I

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

TRImD LIFT COEPFICIEM CL TRIM

Fig. 55 COMPARISON OF FLIGHT VS. PREDICTED TRIMMED ELEVON DEFLECTION

Fig. 54 COMPARISON OF FLIGHT VS. PREDICTED LIFT COEFFICIENT

FREE RIWT 4

A M E FL1m 5 .A& noB ACD . 0.0215

1

EXTENDING

I I I 1 -1 -2 0 2 4 6 8 10 12 14

TIE (SEWNOS)

Fig. 56 COMPARISON OF FLIGHT VS. PREDICTED DRAG