System Technology Analysis of Aeroassisted Orbital Transfer Vehicles: Moderate Lift/Drag (0.75-1.5)

DOCUMENT NO. 85SDS2184 NASA CONTRACT NO. NAS8-35096

PHASE I FINAL REPORT

VOLUME IB-PART I STUDY RESULTS

SYSTEM TECHNOLOGY ANALYSIS OF AEROASSISTED ORBITAL TRANSFER VEHICLES: MODERATE LIFT/DRAG (0.75-1.5)

AUGUST 1985

(bASA-CI_-1791q1) 5XS_[_M T];CEI_CLCGI ANALYSIS N87-26CfQb ¢.F AE_OASSIS_D O_BI_IL TRA_S}E_ VEHICLES: _C/}£RATE LI.F_/I;_JG (0.'JS-1.5),VCLrdBE IB, FIBT 1, S_UDY 5Ec[_L_I5 _inal _Fczt (General Unclas £lectric Co.) 228 _ Avail: _II_ HC G3/|6 00855_5

SUBMITTED TO

GEORGE C. MARSHALL SPACE FLIGHT CENTER NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MARSHALL SPACE FLIGHT CENTER , ALABAMA 35812

RE-ENTRY SYSTEMS OPERATIONS 3198 Chestnut St.. PhiladelDhia. PA 19101

GENERAL @ ELECTRIC FORWARD

This final report of the "System Technology Analysis of Aeroassisted Orbital Transfer Vehiclps: Moderate Lift/Drag (0.75-1.5)" was prepared by the General Electric Company, Space Systems Division for the National Aeronautics and Space Administration's George C. Marshall Space Flight Center (MSFC) in accordance with Contract NAS8-35096. The General Electric Company, Space Systems Division was supported by the Grumman Aerospace Corporation as a subcontracto_ during the conduct of this study. This study was conducted under the direction of the NASA Study Manager, Mr. Robert E. Austin, during the period from OctobeL 1982 through June 1985.

The first phase of this program focused on a ground based AOTV and was completed in September 1983. The second phase was directed towards a space based AOTV and the cryofueled propulsion subsystem-configuration interactions and was completed in March of 1985. The second phase was jointly sponsored by NASA-MSFC and thp NASA Lewis Research Center (LeRC). Dr. Larry Cooper was the LeRC study manager.

This final report is organized into the following three documents:

volume IA Executive Summary - Parts I & II Volume IB Study Results - Parts I & II

Volume II Supporting Research and Technology Report

Volume III Cost and Work Breakdown Structure/ Dictionary

Part I of these volumes covers Phase 1 results, while Part II covers Phase 2 results. The following personnel were major contributors to this study in the areas shown. Study Manager and Principle Investigator D.E. Florence (215-823-3129) study Manager for Grumman Subcontract G. Fischer (516-575-2361) Concept/Configuration Design R. Hassett G. Fischer K. Sneddon Propulsion Subsystem G. Fischer K. Sneddon Man Rating . G. Fischer Crew Capsule G. Fischer Astrodynamics/Flight Mechanics W. Letts R. Klund Aerodynamics J. Berman J. Gardner C. Harris

Aerothermodynamics R. Brewer T. LaMonica L. Arrington D. Nestler

Thermal Protection Materials Dr. R. Tanzilli

Cost Analysis G. Fischer

ii TABLE OF CONTENTS

INTRODUCTION

SYSTEM ANALYSIS

2.1 Flight Mechanics (or Mission Analyses)

2.1.1 Mission Model 2.1.2 _V Budgets 2.1.3 Aerodynamic Plane Change 2.1.4 Typical Trajectories and Performance

2.2 Aerothermodynamics

2.3 Aerodynamic Configuration Development

2.4 Configuration/Concept Development

2.4.1 Mass Estimates 2.4.2 Initial Configuration Screening to Meet Center-of-Mass Requirements 2.4.3 Unmanned/Manned Vehicles Schedule Implications 2.4.4 Manned Cabin Definition 2.4.5 Small Manned AOTV 2.4.6 Perigee Kick AOTV

2.5 Payload Delivery Capability

3.0 SYSTEM/SUBSYSTEM TRADES

3.1 Man-Rating

3.2 Propulsion

3.3 Thermal Protection Subsystem (TPS)

3.4 Structure

3.5 Avionics

3.6 Flight Controls

3.7 Aerospace Support Equipment

3.8 Electrical Power Subsystems

4.0 REFERENCES

iv PRECEDING PAGE BLANK NOT FiLME.D VOLUME I - PART B

1.0 INTRODUCTION

Significant performance benefits can be realized via aerodynamic braking and/or aerodynamic maneuvering on return from higher altitude orbits to low Earth orbit, Reference 1-5. This approach substantially reduces the mission propellant requirements by using the aerodynamic drag, D, to brake the vehicle to near circular velocity and the aerodynamic lift, L, to null out accumulated errors as well as change the orbital inclination to that required for rendezous with the Space Shuttle Orbiter. A study has been completed where broad concept evaluations were performed and the technology requirements and sensitivites for aeroassisted OTV's over a range of vehicle hypersonic L/D from 0.75 to 1.5 were systematically identified and assessed. The aeroassisted OTV is capable of evolving from an initial delivery only system to one eventually capable of supporting manned roundtrip missions to geosynchronous orbit. Concept screenings has been conducted on numerous configurations spanning the L/D - 0.75 to 1.5 range, and several with attractive features have been identified.

Initial payload capability has been evaluated for a baseline of _elivery to GEO, six hour polar, and Molniya (12 hours x 63.4 orbits with return and recovery of the AOTV at LEO. Evolutionary payload requirements that have been assessed include a GEO servicing mission (6K up and 2K return) and a manned GEO mission (14K roundtrip). 2.0 SYSTEMANALYSIS 2.1 Flight Mechanics 2.1.i General Mission Model The generalized mission model addressed is summarized in Figure 2.1-1. The initial ground based AOTV's will be deployed from a 150 nmi circular orbit, launched from ETR at an orbital inclination of 28.5 _. Eventually, _he space based AOTV's will be in a 200 nmi circular orbit at 28.5 inclination. Launch vehicles considered include the standard STS, an improved STS, the aft cargo compartment (ACC) and the shuttle derived cargo vehicle. Operating scenarios were established for the several reference missions and _v budgets determined for use in the performance compuations. Effective use of aeromaneuver for return from Molniya is not possible due to large _ V required for apsis rotation of the Molniya orbit. Initial payload capbility has been evaluated for a baseline of _elivery to GEO, six hour polar, and Molniya (12 hours x 63.4 orbits with return and recovery of the AOTV at LEO. Evolutionary payload requirements that have been assessed include a GEO servicing mission (6K up and 2K return) and a manned GEO mission (14K roundtrip). 8

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The characteristic velocity for the baseline missions were determined parametrically with Shuttle parking orbit inclination and transfer method, i.e., two-impulse, three impulse, or transfer through infinity. It was assumed that all of the inclination change for the outbound leg was performed propulsively. The inbound leg was treated parametrically with propulsive inclination change and apses rotation.

The impulsive transfer from some initial parking orbit to a desired final orbit can be performed in several ways; a one, two, or three-impulse transfer, and a transfer through infinity. One-impulse transfers are not common since the two orbits must intersect or be tangent. The more common type of transfer is the two-impulse, of which the most familiar (for coplanar transfers) is the Hohmann transfer. Three impulse, time-open transfers are rare but have been shown to be optimal for transfers to highly ellliptical orbits where large inclination changes are required (16). The transfer through infinity assumes that the initial impulse places the vehicle on an asymptotic escape orbit from which a small impulse at a large radius (approaching infinity) can be made to complete the transfer. The second impulse is small due to the small orbital velocity at a large radius. While constraining the transfer to finite radii the optimality of either two or three-impulse transfers are dictated by required plane change and the initial and final orbit apogee and perigee radii.

The equatorial, circular geosynchronous mission is straight forward in that the logical choice for the shuttle parking orbit is 28.5 degrees so as to minimize the required plane change. It can be shown from Gobetz and Doll (19) that the transfer can be made optimally using a two-impulse transfer. The first impulse of 7985 ft/sec inserts the vehicle into a transfer orbit with perigee at 150 nm (Shuttle parking orbit altitude) and apogee at 19,365 nm (circular geosynchronous altitude). The second impulse of 6009 ft/sec, approximately 5.1 hours later, circularizes the vehicle at GEO and changes the orbital inclination from 28.5 degrees to equatorial. The return leg consists of a transfer to an orbit with the perigee within the atmosphere (60 nm altitude) and requires an impulse from 5400 ft/sec if no propulsive inclination change is needed to 6420 ft/sec if all inclination change is performed propulsively. Upon exit from the atmosphere following aerodynamic braking and maneuvering an additional impulse of approximately 200 ft/sec is required to circularize in low earth orbit. Another 300 ft/sec is budgeted for LEO phasing maneuvers. The above budget is summarized in Figure 2.1-2. i I , •

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-a The six-hour polar mission, due to the large inclination change, requires a considerable characteristic velocity. The outbound transfer from LEO made with either two or three impulse maneuvers, depending on necessary plane change, is shown in Figure 2.1-3. Basically for Shuttle inclinations greater than 40 the transfer can be o_timally with only two impulses, for inclinations less than 4_ the three-impulse transfer is optimal. For transfer from a 28.5 shuttle orbit the outbound delta V is approximately 17,300 ft/sec. The return from the six-hour orbit to entry is shown in Figure 2.1-4 as a function of plane change performed prop_isively. As much as 13,000 ft/sec impulse may be required if 60 v of plane change is needed. The total outbound and inbound delta V is shown in Figure 2.1-5 as a function of both inbound plane change (that not accomplished aerodynamically) an_30 shuttleShuttle orbit andinclination.return to Asa _9 planebe seen (60aotransferplane changefrom a would require 30,700 ft/sec total delta V. In addition to these impulses another 300 ft/sec for phasing and 200 ft/sec for LEO circularization following atmospheric exit should be budgeted.

The Molniya mission provided additional complexity in that the transfer is no longer circular to circular but circular to elliptical which requires placing (or rotating) the line of apses to the desired position. The Molniya orbit is defined to have a period of 12 hours, inclination of 63.4 and perigee altitude of 400 nm. The apogee is usually constrained to be located over the northern hemisphere implying that the argument of perigee will lie between 180 and 270 v. In performing a non- coplanar circular-to-ellipt_cal two-impulse transfer the transfe angle should not exceed 180 for time-open transfers. The transfer eliipse is usally entered just befor_ its p_rigee, and perigee passage on this ellipse occurs from 0 v to 90 before perigee of the terminal ellipse. In addition the final orbit is always entered near a node. A schematic of a typical transfer from LEO to Molniya is shown in Figure 2.1-6. As would be expected the delta V increases with inclination change. Also since the orbital velocity at the ascending node of the transfer and final orbit increases as the argument of perigee increases (from 180 ° to 270 ° the delta V will also increase. These trends are illustrated in Figure 2.1-7 which shows the outbound delta V as a functio9 of shuttleoorbit inclination for two arguments of perigee; 220 (apogee 40 from ascending node) and 270 (apogee 90 from ascending node). The solid lines indicate a two-impulse transfer while the dashed lines indicate three-impulse maneuvers. As can be seen a delta V ranging from 8400 ft/sec to 15,300 ft/sec may be required for the outbound leg. The return or inbound leg may again require some or all of the inclination change and some rotation of the line of apses (so that atmospheric entry can be made near a node - thereby maximizing FIGURE 2.1-3 DELTA V REQUIREMENTS FOR ASCENT TRANSFER FROM LOW EARTH ORBIT

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v pla_e change via aerodynamic lift). A return from Moln_ya with a 220- argument of perigee would require approximately 35 of apses rotation to put entry at or near a node. The inbound delta V as a function of inclination change and apses rotation is shown in Figure 2.1-8 and indicates that between 8500 and 9500 ft/sec may be required with a 35 apses rotation.

2.1.3 Aerodynamic Plane Change

Earlier studies (16) reported results for plane change assuming that the flight is performed following entry at the overshoot bound (minimum entry path angle to ensure aerodynamic capture) and using a constant drag control law (17) used extensively in aerocapture studies. However it is demonstrated below that operation at the overshoot bound and use of the constant drag control law is not necessarily compatible with maximizing aerodynamic turning for plane change.

The hypersonic L/D varies significantly with altitude, due to the viscous drag effects, but has been treated as a constant here since it has been demonstrated (6,22) that inclusion of the high altitude effects has a trivial effect upon the plane change obtained. The AOTV mission ends at a Mach number of about 25; therefore, the usual low hypersonic and supersonic variation of the aerodynamic characteristics has no effect here either.

A series of computations were performed for a representative AOTV assumed to weigh 10,000 pounds, with a lift to drag ratio (L/D) of 1.0 and a W/C_A of 68 psf. Two methods were used to determine the maximum p_ane change; first, entry is made at a path angle greater than or equal to the overshoot bound. The lift vector is modulated so as to maintain a reference drag that has been pre-selected to correspond to that deceleration sensed at the minimum altitude. The vehicle is commanded to rotate the lift vector full up at the proper time so that atmospheric exit is made at the optimum velocity (optimum in terms of minimal delta velocity for orbit circularization). In an alternate approach, the vehicle flys a constant bank angle throughout the flight and entry is made at the path angle which will result in an atmospheric exit at the optimum velocity. In addition to plane change reference values of hypersonic convective heat transfer (normalized to a one foot nose radius) is given for both methods.

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o 0 10 20 30 40 50 60 InclinationChange - deg 11 The results for flight using the constant reference drag control law are illustrated in Figure 2.1-9. As can be seen a 33% increase in plane change capability oxer operation at the overshoot bound can _e realized (from 8.5 _ at the oxershoot bound to a maximum of 11.5 at an entry path angle of 4.8 _ and reference drag of 1.0 G's). The stagnation point heat flux increases from 170 BTU/sq.ft.sec to 205 BTU/sq.ft.sec for a one foot nose radius. Figure 2.1-10 shows results for flight at a constant lift vector bank aggle. It can beoseen that the plane change is maximized at a 90 bank angle (90 corresponds to all lift vectored normal to the flight path plane and 0 indicates full lift up). In essence when flying at a constant 90 _ bank angle the vehicle is operating similar to a full aerobraking device, depleting the proper amount of velocity before centrifugual acceleration pulls the vehicle out of the atmosphere, while using all _vgilable lift to turn. This maximum plane change of 14.6 is approximately 70% greater than the plane change at the overshoot bound using the reference drag control law. The normalized stagnation point heat _lux is approximately 274 BTU/sq.ft.sec for the constant 90 _ bank angle maneuver.

2.1.4 Typical Trajectories and Performance

Time histories of stagnation point heat flux, free stream Reynolds number, and dynamic pressure are shown in Figures 2.1-11, 12 and 13, respectively, for the overshoot bound/ reference drag flight, the maximum plane change/reference dr_g flight, and the maximum plane change/constant bank angle (90 _) flight trajectories. The primary point to be derived from these figures is the fact that although the stagnation point heat flux is higher for the maximum plane change trajectories the soak or heating time tends to be significantly shorter. Therefore, these lower integrated heat loads although yielding higher skin temperatures may require less thermal protection. The respective velocity, altitude, 6nd Mach number histories are illustrated in Figures 2.1-14, 15 and 16.

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16 Performance studies have been conducted for return of mid L/D vehicles from GEO, 5 x GEO, and 6-hour polar circular orbits employing steering laws that include constant deceleration cruise at the overshoot and undershoot bounds, and constant bank angle cruise. Orbital plane change obtained is summarized in Figure 2.1-17, where it is shown that plane change capability increases with hypersonic L/D and entry velocity (maximum for the 5 x GEO return) for a specific steering law. The 90 bank angle provides the maximum plane change.

Use of the various steering laws results in different minimum altitudes and thus different maximum heating rates, Figures 2.1-18 and 19. It can be noted that maximum heat transfer rate increases with vehicle ballistic coefficient, W/C_A, with increasing entry velocity (5 x GEO results in maximum entry velocity) and with decreasing minimum flight altitudes (constant 90 bank angle results in minimum flight altitudes).

Minimum flight altitudes for these vehicles extend down to 160 kft for return from 5 x GEO of a 24K ibs, L/D = 1.5 vehicle. At altitudes below 190 kft, local boundary layer transition to turbulent flow may occur with a resulting increase in convective heat transfer by a factor of two to five. This must be avoided in order to minimize peak temperatures on the aft frustum area. A possible simple strategy involves employing an out-of-plane propulsive burn to obtain part or all of the desired plane change during the deorbit burn. This provides a significant increase in minimum flight altitude, laminar flow over the aft frustum, and reduced local heat transfer rates, Figure 2.1-18. At 5 x GEO altitude there is a very small propellant weight penalty for this burn. This technique allows flying a reference drag deceleration trajectory at the overshoot boundary.

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Numerous steering law evaluations have been conducted (5, 22, 29, 30, 31, 32 and 33) to determine the magnitude of LEO circularization burn _V uncertainties resulting from an off-nominal atmosphere, errors in entry interface state conditions (VE, _E), and uncertainties in AOTV aerodynamics.

The insensitivity of an L/D = 1.5 AOTV to variations from the nominal in the atmopsphere density or to errors in the apriori estimate of the drag coefficient have been evaluated by personnel from NASA JSC and are illustrated in Figure 2.1-20. Note that the mid L/D is relatively insensitive to atmospheric and drag coefficient uncertainties.

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Several hypersonic flow regimes are traversed as the AOTV enters the atmosphere and executes the pull-up maneuver. Regimes encountered begin with free molecular flow and proceed through the transitional and viscous merging slip flows, finally entering full continuum flow during the constant "g" or minimum altitude portion of the mission and then proceeding back through the rarefield flow regime to free molecular as the exit maneuver is performed. Free molecular and continuum drag and heat-transfer characteristics are well defined from both _heory and ground and flight test data. Characteristics in the transition and slip flow regimes are less well defined and predictions are based on a shock Reynolds number semi-empirical bridging technique developed by Gilbert and Goldberg (25). This bridging relationship represents a correlation of the numerical solutions of a theoretical model of the hypersonic viscous shock layer of air in chemical equilibrium over a range of shock Reynolds numbers of 50-10 and is illustrated in Figure 2.2-1. The range of shock Reynolds numbers computed at peak heating for the various AOTV trajectories is illustrated in Figure 2.2-1 and indicates that _,,= v=,_cle fLu_u,, is uF_**_ in full continuum L=gime during this time. For purposes of this evaluation, only continuum heating will be considered.

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24 Numerous vehicle configuration and entry trajectory parameters effect the magnitude of the hypersonic convective and radiative heat transfer rates experienced by the AOTV. These parameters include entry velocity (mission dependent), entry path angle (determined apriori depending on steering law selected), vehicle configuration and size, ballistic coefficient (W/CDA), angle of attack, steering law and number of atmospheric passes. The combined impact of W/C,A, entry trajectory/steering law and mission have been presented in Figure 2.1-19 for a single pass mission. In this study vehicles with significantly different configurations and size have been evaluated. The nominal cases evaluated have been flown at an angle of attack consistent with maximum L/D (to maximize paylsad delivered)'. Typic_l values of angle of attack range from 25 for L/D - 0.75 to 15 v for L/D - 1.5, Figure _.3-II.

The convective hypersonic heat transfer computed during this study was based on equilibrium flow and a fully catalytic thermal protection material surface. Space shuttle flight experience as well as numerous investigators (26, 28) have indicated that the shuttle is experiencing convective heat flux 20% less than the fully catalytic values, but more than the 50% -^;"-_ that ...... i_ be expected with a fully ,n,-_atalytic surface. Computations for a range of AOTV's has indicated (20, 21) up to a 67% reduction potential for a fully non-catalytic material. Using the shuttle experience of a partially non-catalytic surface, for the AOTV flight conditions, results in an expected peak surface tempsrature reduction of up to 200_F on the aft frustum and up to 400 F on the nose.

Estimates have been made of equilibrium hot gas radiation using the results of Page(13)for the range of initial vehicles selected, Table 2.2-1. It can be seen that the hot gas radiation for the nose area is approximately 10% of the convective heat transfer for all cases except the manned vehicle (heavy AOTV) return from 5 x GEO where it increases to about 20%. Aft on the vehicle, the hot gas radiation drops significantly and is expected to be an even small fraction of the convective heat transfer. Hence, for the purposes of this study, hot gas radiation will be neglected, since it is a second order effect compared to all other parameters.

The coupling involved between steering law employed, aerodynamic plane change obtained, resulting hypersonic convective heat load and maximum heat transfer rate, heating time and thermal protection system weight required is illustrated in Table2._-_or an L/D = 1.04 and W/CDA = 68 psf Note that as the plane change is increase from the overshoot bound value, the heat load and heating time decrease significantly while the maximum heat transfer rate increases. Thus, higher allowable local TPS surface temperatures, permit a lighter thermal protection subsystem.

25 Table 2.2-I. EQUILIBRIUM HOT GAS RADIATION IS SMALL REALTIVE TO CONVECTIVE HEAT TRANSFER

RN= 1 FT o o L/D MISSION qRsTAG ,, qSTAG,_ (BTU/FT L SEC) (BTU/FT L SEC)

0.75 5XGEO 22 294

1.04 5XGEO 31 362

1.5 5XGEO 60 511

1.5 5XGEO 155 776 MANNEDRETURN

26 z_

O z_ _ C_ -I- -1-

Im LIJ

0 i,i Lt_(I)

rY J__Jn

LI- < _

_m c_ C_J

J,i

-r-

Z z _ c_ w=

_mt l,i

i, 0 Im t_ CJ I,i I, kL i,i c_ ! C_J c_ Lil _J

t_ Im z_

(11

z_ _ _o _

27 2.3 Aerodynamic Configuration Development

The AOTV configuration must provide a high packaging efficiency for the propellant tanks and various subsystems as well as meet the external constraints of the launch vehicle. The length and weight constraints to be used in this study for the various contemporary and advanced launch vehicles have been specified by NASA and are summarized in Table 2.3-1. The principal aerodynamic configuration drivers and assumptions identified at the initiation of this study are summarized in Table 2.3-2.

28 ==1 x,.

I-- o .-i- o 0 13,1

o U.I

uJ_ I=- ._.. Z mz 0 O_

a.I- 0 _ Z X

_o> -i- I_ "'0 I-.i ll.:

-r- z -'r" O0 I-- Z 0 o o Z I.U _J .-I I- z I,U :E U.I I-- ,-I c-o LU U ..I u !- Q. =l= :E ul (lJ I.u i- _w e,=,. 0 m--I .,I- 0 ro /.I.I m t-- 0 U C_ 0 I.U z U.I e,, w:> > 0 0 I-0 --I, k--O (/) I-- I.- I.l. .< mU 0

29 0

I,-- Z 0

< _1

I-.- ,_1

Vl u.I "0 g_ Z O' 0 irl I.- 0 >.. < ,,_1 r,n U a3 0 0 t_ tn r,, .J I.u IJJ ,..I

m Z t- u. I- LU I.IJ ..J Z t- o U 1.1.. f_ I- Z 0 0 I.U 0 0 0 0 d 0 I d I- ,_J Z 0 < °. _J I.U X X < U w u. e_ < Z uJ _J Z Z i-- LU < >- Z Z _J n- m i I- tn O 0 z 0 0 kg Z 0 < < _J U .J e_ ..J Z Z _J < < 0 I e,, O3 I/1 >. ILl >- Z C_ ..J r"- r-'- t-" I .43 0 Z _0 _J I- I'-- I- I-- _J z

30 The hypersonic L/D varies significantly with altitude, due to the viscous drag effects, (6,23) but has been treated as a constant here since it has been demonstrated that inclusion of the high altitude effects has a 3rd order effect on the plane change obtained and payload delivered (6,22). Some insight into these results is obtained by noting that the flight regime where 90% of the velocity loss is experienced (20) is the continuum flow regime.

To provide initial direction to the mid L/D AOTV packaging studies, it was necessary to have an axial center-of-mass location requirement for both launch (propellant tanks loaded, AOTV in the shuttle orbiter) and entry (propellant tanks empty and in some cases staged).

xcm requirement at launch is clearly defined in "Shuttle Orbiter/Cargo Standard Interfaces" ICD2-19001, JSC07700, Vol. 14, Attachment I, Rev. G (24) for the STD 65,000 lb. STS. The entry Xcm requirement for both the AMOOS type configurations and the higher L/D biconics that are 60 ft. long is in the vicinity of 55% of the vehicle length (aft of the nose) to prov _ .... _- angle consistent with (L/D) max, Table 2.3-3. This value was used in the initial configuration screening process.

31 I,N c_ > I U') 14") I,/'I e_ C3 X _D Z > Z Lit 0. I.IJ t) L/. n, I X U

m Z _- z 0 _I LU i _" UJ < UJ O_ c_ D U < 0 LL .=- i._ o =2. i_ _i co 7 •-• c_ ,-. • ,-• ,- ,_ 0 U U_ < C_ _u 0 X LL < < 0 z Z >- Iz W c_ Lu 0 I- LU z ._] C_ ILl ILl < U D C LU < ,,, C_ _ • =,,=, .U I-I.- z w < i z V-- < w<

_ 0 "J,_ I t_ _ w < d < 0.1 m m w o i--. I1 W ILl 0 0 0 w w z 0 0 0 > o_ O < < < 0 I-- I- I- I- i i t :_ C. C. 0 > < IJJ W ILl i D O O , C, UJ W W L._ _. _ (..9 -_- -r z 1,1 w z U --- 14. v ,/ v _ < < _ z C,) LJ U w w w _ O 0 0 0 0 w ,,, 0 w I1 C.) ...j ...j ..j m r,D _.D U3

32 To conduct the initial 3-DOF flight mechanics evaluations, where in atmosphere flight control steering laws were evaluated, primary operating altitudes, hypersonic flow regimes, and reference heat transfer determined, it was necessary to specify some typical aerodynamic characteristics. A survey of previous mid L/D AOTV type vehicles was conducted, Table 2.3-4. For the initial evaluations an AOTV return weight of 10,000 ibs was assumed. Using the aerodynamic characteristics of AMOOS 5B, operating at (L/D)max, results in a W/CLA of 68 psf.

33 -i-

I._.._JQ._ z

t._ b-.4o --_ _J l-- LL _.J

1.1.

..I co

.J 4¢ ,,it o o _o u s ..I w_ II

n, Lu u.i z o o .._ t_ 0 v !- w z

_w t.O 0 > L_ z 0 0 C_ m 0 I-- < I'-'- < < O 0 X I < (D UJ W n- < LI. z I .....,. Z W 0 O 0 _J

34 The mid L/D configuration evolution was aided by the large aerodynamic coefficient analytical and experimental data bank that already existed for conic and biconic bodies wish various control surfaces and aft frustum angles down to 4 . Additional characteristics are available from the AMOOS studies, Reference 6 for cylindrical aft bodies and an aft frustum angle of 0.5 _. To supplement these available results, additisnal ne_ computations were performed for aft frustum angles of i and 2 employing HABP, Reference 8, various nose and vehicle length combinations, and various nose bend angles. This chronological sequence of events is outlined in Figure 2.3-1.

The initial configuration class defined for performing the additional computations is illustrated in Figure 2.3-2. The results of these computations are summarized in Figures 2.3-3, 4, 5, and 6 for a vehicle with a full nose bend.

The dramatic effect that the nose length has on hypersonic L/D is illustrated in Figure 2.3-3; the change of nose radius from one to two feet is shown to have a negligible effect on maximum L/D in Figure 2.3-4. The aft movement of center of pressure location --_..... angle uf attack and increased nose length is illustrated in Figure 2.3-5; the effect of increased volume, non-circular cross section nose on hypersonic L/D is illustrated in Figure 2.3-6.

35 I

I, BI _ _ _ _ I tm, I.._1 Zl i, w o_m o o Z -IW 0____, O m I U')I I-- zl

:::::3 Ol .... _o_ I-- F,, F.-. _J <_ t'r" w O -- Z 11_" _ooZ x g I.iJ W C.9 M D._I 14. 1:3 • ILl'" ° , / O i O I-- u <_ r'r" ,/ _1 _' G3 (...9 ii i + z W O1--1 '"-, Or'_ --W i, O g")l JELl) -- I"-- IJIJ Lf_z J-- Z '"_1 (D <[ -"_' ' L.O • 0 -- O_W i.i.,, i,.i.. -- O zr,'l Q ,7,U_ 0C3 O1-1 '_- _c01m m E_ o:_, o 2->-I .'-'_-_,._IL_ " '_ l.t.l(_lnl._J +4 _ _N\ I.I. ×£z 0 U31_ ZU3 "x

Z l, I llll_l.O >-. zz, _ I r_.l_ Z owl q'_-I i x}--On," rr_. _ 123 °_I. I I'-"rd")L_')": O r'r LLI <[: E3 I- z z 0 \ w rd') _J Z 0 i,i D <.9 I-- z __jW _1 U3 :3 _O zZ_

(.3(.o xz _Jz 7-'" n" % >-,.7_"_/11 _ I, z" • •

I.£.I"I i +,+ + ,+, '

36 l--

u_ r i

0 ul

...I

LLI

Z 0 m h-

i LL Z 0 o o m

Z >-

0 n'- u.l

m b

I u. l'" Ll. ,C,

z 0 o I'-- z c_ Jl II z z

37 Figure 2.3-3. EFFECT OF NOSE LENGTH ON HYPERSONIC LIFT/DRAG RATIO

2.0 m

1.9

- / %_ LVE H = 60 FT 1.8 -#/ • R N = 1.5 FT 0 _ 1.7 f "\

_1.6

I.- / \ ,\ Li. t ¢ _m .J U zt.q 0 ul ,v _.I.3 "1/ "\ 3s l PACKACING >- -r VOLUME INCREASES !.2

1.1 15 VISCOUS EFFECTS NOT INCLUDED 1.0 I I I I I ~ io_ 0 5 I0 15 20 25 ANGLE OF ATTACK, _(o)

Figure 2.3-4. EFFECT OF ANGLE OF ATTACK AND NOSE RADIUS ON HYPERSONIC LIFT/DRAG RATIO

2.0

1.9

1.8 .1 LVE H = 60 FT

1.7 : = 20 FT I- NOSE <

U 1.6 ,i( Q 1.5 l,- u.

I.q U z 0 1.3

a. >, z 1.2 _i (FT)

1.1

1.0 I ! I I I 5 10 15 20 25

ANCLE OF ATTACK. ,_(o)

33 Figure 2.3-5. EFFECT OF ANGLE OF ATTACK AND NOSE LENGTH ON CENTER OF PRESSURE LOCATION

(FT)

..,._ 35 0.5

;> ,J o. U L x - O.q z O

U O .J ua 0.3 ,v

u1 uJ e_ a. u. O 0.2 ._ RN = 1"5 FT iv. ud I-' Z Ltl U 0.1 , I I I I I 5 10 15 20 25

ANGLE OF ATTACK, _(o)

Figure 2.3-6. EFFECT OF NOSE CROSS SECTION ON HYPERSONIC LIFT/DRAG RATIO

LVE H = 60 FT

CNOSE = 20 FT NOSE CROSS SECTION R N = 2FT 1.5 - 0

LL D ,.I / (,.) Z 00.5 u_ ac uJ )- -r /

0 I | I I I 0 5 10 15 20 25

ANGLE OF ATTACK, cz(°)

CONCLUSION: NON CIRCULAR CROSS SECTION NOSE OFFERS INCREASED PACKACINC VOLUME AT SMALL LOSS OF L/D

39 For a single stage vehicle, propulsion stage packaging trends were evaluated to determine vehicle center-of-mass possibilities for combinations of total vehicle length, LV, and nose length, Ln, Figure 2.3-7. Additional aerodynamic computations were performed for shorter vehicles and for vehicles with less than a full nose bend, Figures 2.3-8. These results, in combination with the parametric center-of-pressure locations, Figure 2.3-9, were used to define three configurations, Figure 2.3-10, that span the range of L/D from 0.75 to 1.5 for further evaluation.

4O -C

Z 0

0 oz

Z o 0

< Z

0 >

0 •

-0

•T- _ % _'0

41 q

XVIAI (Oll)

42 0,I ©

c_ •"_ 0 _ .._ © C_ o [-

: _: : ! _0 0 : : • : : / _ 0 0 Z _ . , _ G . _:_ ¢_

t_ 0 0 © -_ 0 0 i : : i ! © i ! ! ! o _ < i 0 C_ _ m 0 ©

o j • .,_ 0 90 g'O I,0 _0 _0 ©

I © 0 © > Z \ :

c_ °,,,_

r,.)v iiiii iiiiiiiiiiiiiiiQ; ©

,-..1 _.. 0 © ...... _....._..oi..... 0 Z •. g._ Z © 0

0 ©

% gO 0

43 %)

ORIGINAE PAGE IS %) Lt_ OF POOR QUALITy

%) .,=_ II o

O_ II II tnZ_

o X

%) .,=.4

C90

CD ...... l ...... • !

! r_ .,_ ¢,; @ ,< 0 t._ k9 8 _D l! t4. O -_

• 0,,_ • p,,,l : c-v O tO _b "o II Il _} .,=_ i, -..<

0 !| 0 Ib -.w. --Q

I °_ t_ I // / 0 / \, /

44 Specific aerodynamic characteristics of these three configurations have been generated and are summarized in Figures 2.3-11, 12 and 13. Note that _he angles of attack required to obtain (L/D) max range from 15 to 25 _, Figure 2.3-11. The effect of decreased nose bend on hypersonic L/D and static pitch stability, _ Cm/ _Cn is illustrated in Figure 2.3-14. Figu:e 2.3-11 Variation of L/D with Angle of Attack for AOTV Interim Configurations

: I,-- 45. I N -- 7.5 J() ..---- ._ i : i, GO tN :'.() l.. . 1. ,rj() IN

NEWTON I AN AERODYNAMICS o RN = 2 0 FT o_ ...../..i ..... \_ ..... !\ ...... : ¢_ ii" ..-....?kt"---.i."\.' o ! ._J o0 )-I " / _ ' _L-_

i • t t • / /./':/

V : : : : i

o.o to.o 28.0 38.0 40.0 _8.o 800 70.0 80.0 Angle of AtLack cx (Deg)

Figure 2.3-12 Variation of Center of Pressure with Angle of Attack for AOTV Interim Configurations

I- O

0 ...... !i...... ! ! : : _ : ! J_-" "---" :

', _. -/$" ..__ tO ...... : -,-,"_"....- /

r/j rtj

S., ...... / i._.;/...... _...... i...... _...... i......

o ./<" :: i i ,-,ok, z-..,-,,,-,-:. ;;/ ...... _...... _...... :...... r.---_:_ ..... -q...... : : _ I --,I"_,IN : 7F_ | •// i I I[::: _,0 ]N :: |0 . I

O NEWTONIAN AERODYNAMICS

R N = 2. 0 FT

o.o I 1.o _o.o ,fit) (1 40.0 50.0 GO O 70 () RO 0 An_le of AI.L;_ck c((l)eg) 46 O O O

J : I Z

_0 o 0 o ° i,,-,-I .< Z _ o

)7 Z II t.O ,' / z °p,--I / I, I11111 _ o ° f--,,I / 7 O ¢g .a,..a ...... _ ...... t.: ...... l_...... i:i I" Z o ¢.) \ " ° t,-,-I O

...... ° ...... _ ......

o 0

-0 >...... : ...... "x ...... ':,: ...... __._'_ "_i _"'_-_: ...... E.- < O "7 J .<

I,.,,,.I ...... : ...... !...... _-- _. _'_::'_. "%--\ .... X ...... ¢-- O o

_ __. _ _N',\

O : _ _ _ : • -\\,.\

oel ...... _ ...... _ ...... \._ ...... 0 "q,\ ,%

0 0 0

U3 q_a!3T_Jgo D 3,jt.q %

47 0

°_--I

) ' " Z

rj II I1',11' Ii oc_ o _'/_/ "" _o _,..__.iZZ .. ,...._i,..._ I .,_ C_ . _ ]. o oo:_l < ? ;/ r,O r..O,',_"q '-' _ /._ :/ , I z _ " ij X II II, II; o /:_ :' _,_I _ ii I o , I_ _: /! /i • c ...... , , r _ ...... i....:...... li ...... }":...... _ • !_ _ /i /_ -_

o I_ ...... : ...... i: ...... :?/: ...... i: ...... / i: ...... or..6 <_

,r-4

: : i -_; i c_

• (,_ 0

i ,// _ Z ...... _-...... _...... -/ i_...... i>_/...... :...... o _: / i "/ i: I' : 0 0 t,i,-t Z

/ i i : ; • i o 0 .,, - _- ...... _ ...... 4 ,/ • ?J

0 cD 0

B'I 9"I #'I 8"I I 9"0 b 0

ORIGINAL PAGE I_ <):: P,<),'-)R QUALITY 48, Estimates have been made of the aft frustum angle and vehicle length effect on (L/D)max and Xcp/Lv using the RVCOR code. The RVCOR code is a high speed engineering design computer code based on algorithms developed from exact flow field computations, and flight and ground test data. These results can be employed to generate the incremental effect on (L/D)max of increased aft frustum angle and nose length, Figure 2.3-15, 16 and 17.

49 FIGURE 2.3-15

RVCOR ''Estimates ' ' of AOTV Configuration Trends

R H = 7.5' ; X F = 40' _ _ _

' c21

: ...... i...... i......

48.0 SO.O 5g.O 00.@

I,,V -- FT

FIGURE 2.3-16 RVCOR ''Estimates'' of AOTV Configuration Trends

RB _ 7.5' " R N 2' " X F =40'

I _IJ>_ \ • \ i

i _-l..... i ...... '-_ -- -_...... _2°

rO,O

L v - FT

5O "7 d

:D CD

51 2.4 Configuration/Concept Development

2.4.1 Mass Estimates

To start the AMOTV mission analysis it is necessary to establish a set of vehicle reference mass properties. This was done by reviewing the OTV mass properties developed by the various contractors who have evaluated the APOTV and AOTV missions (3, 4, 6, 9). Since the vehicles varied in size, weight of propellant, configuration and subsystem requirements, the _=_=_=**_= w_ight_ uan be considered only in trends.

For parts that are directly affected by the vehicle configuration, such as shell length or diameter and tank volume, a unit weight per variable was selected. For subsystems and parts, like avionics and engine, a unit weight is used that is assumed constant for any AOTV vehicle. The structure is assumed for initial sizing to be not critical for the aeromaneuver loading since it now appears that the g level for the atmosphere deceleration will not exceed 1 to 1.5. Therefore, the all propulsive vehicle structure was reviewed as well as the AOTV for unit weight. It was assumed that the vehicle is roughly cylindrical, about 15 feet in diameter.The cryogenic tank material usually selected is 2219-TB7 aluminum. The final tank shape used affects the efficiency of the structure and hence its unit weight, i.e., the toroidal fuel tank is less structurally efficient than a spherical one.

A summary of the mass properties used in preceeding OTV studies is given in Table 2.4-1 and 2. Table 2.4-3 contains the unit AOTV masses employed during this study for configuration and mission evaluation.

The evaluation described above has been made for a basic 65,000 ibs shuttle payload with an AOTV vehicle diameter of 15 feet. For increased shuttle payload masses the following assumptions are made:

o The propulsion related components and subsystem weights are based on total propellant mass.

o The EPS and avionics subsystem masses are assumed constant.

o The structure unit mass increases as the square root of the propellant mass ratio for a range of two-to- one as shown in the MSFC AMOOS study (9).

o The TPS mass will increase by 10% for each doubling of the re-entry weight.

52 {/1

QJ G) c- C)

m---. 000 d

II II

• o . . . C) 00000 r-_

I_LI: _0000 t_ _00_0

Q; QJ

P" S.- _.. QJ Q; 00000 P > •

('_ _i (.C) l I;1 (;1

.S.- r,,. f,,..

_ (_ c:E

('M p-- UUU U(JU U U (-_ c_ U U t.l- LI- L,_ U U t.[.. 1,_. I.l- (,/') t._. _,__ 0

N (._ ,"'_ _'_ Z_..) r-_ _-_ I--- 0 =:_ _j 0

_,0 (...) "r- I I ._1 __ Lz') ::Dr", Z _-_

I--- _._ _ Z

53 ,-- O {.} °_..

__ S- (_ oJO E

o_ (-

u') c- U')

'X_ LC)'_,-- (.) r'_ C,.) CP • ,-'- O •_-- _1_ II •,'- r,-.O _ (D e- -'_ QJ CO" r-. •_ O ,,, .._ ...-.. "_ II II _ S.. u"l C_ I oc_ ed • -r

°_-- I.O -r

or)

_--_IO r,,.C_j _i._ ¢'_,.--,--O I(',')I'_.

rv, _ m

(.}

(D C_J O,J

OOOOOOO P. r_ u_ _D(_ C) OOr-.r-.

o _-- _ "_ O 4-) .>- o=

(I; _- (" S.. rt}

or} c.) od I O u_ (" r_ oJ O _-

N "-_ ::3: S.. •s,-- c._ (- c_r', O-¢: 5..._"- -_-- U') (" c_ F-- _._- (D ._-_ u_ O _._ v ",--C_J v -I-_ _" S- _3 CZ (D_ Z _ _ O _ _ _ -_ _Z • mm (..)-r-r_" I_ c_O

r_. -ID_ o ,--- Z c_ 0 0 _" _j_ II :_)

_- Z _-- ___ D h_J r-_ rv" O {._ CD mO

L._ rY" _.1 __1 "_ Ls.l

54 Table 2.4-3. UNIT AOTV MASSES EMPLOYED

SUBSYSTEM UNIT MASS

STRUCTURE SHELL 25#/ft of length A combined 1.04 psf of vehicle surface, SUPPORT 20#/ft of length which includes a 10% eo,_#,_._ used after I st qtr _/I/ FUEL TANK 60#/k#H 2 OXlD. TANK lO#/k#02 FLAPS 49#/k#AM

TPS TANK INSUL. 8#k# prop SHELL 28#/ft of length

PROPULSION ENGINE 433# TVS, PLUMBING 12#/k# prop

ACS 38#/k#AM

EPS 600#

AVIONICS 607#

INFLT. LOSSES 8.4#/k# prop

RES & RESIDS. 0.02# prop

CONTINGENCY 10% is used after the first quarter

55 For use in performance calculations during the first and second quarters, mass estimates are made for several vehicles, Tables 2.4-4, 5 and 6. The structural shell and frame supports and the external thermal protection system were estimated based on AOTV length. During the remainder of the study, these weights were estimated based on total AOTV surface area, Table 2.4-7 As the total AOTV surface area decreased with increasing aft frustum angle and longer noses, this approach provided a more representative estimate.

56 ORIGINAI_ P_/G_ IS OF POOR QUALIT_

Table 2.4-4 ,NITIAL8ASEL'NEAOTVMASSPRO_RTIES Table 2.4-5 INITIAL BASELINE AOTV MASS PROPERTIES GEOD_LIV_RY,SKSTS Lv=60ft GEoS_RVICI.G_ .P 2. _AC.• 65K STS LV=60ft

SUBSYSTEM MASS ILBS) SUBSYSTEM MASS |LBSI

STRUCTURE 3682 STRUCTURE 37S6 SHELL 1500 SIIELL Ill0

SUPPORT 1200 SUPPORT I #0o

FU[I TANK 18q FUEL TANK 39(_

OXIDIZER TANK 380 OXIDIZER TANK )_i

FLAPS 21| FLAPS 26_

TPS 2035 TPS

TANK INSULATION 355 TANK INSULATION $_g

SHELL INSULATION I ilO SHELL INSULATION 1El0

PROPULSION 1118 PROPULSION 1217

ENGINE 133 ENGINE ll33

TVS. PLUMBING ?55 TVS. PI.UMBI 1_ 151

ACS 374; ACS qSi

FP_-, 6OO E_ COO

AVIONICS S07 AVIONICS fdP7

DRY WEIGHT II151 DRY WEIGHT Ills

Table 2.4-7. BASELINE AHOTV I._SS I'NOI'ENIII.S Tabl e 2.4-6 ,N,TiALB*SSL,NEAOTVMASSPROPERTIES GEO I_K UP & IqK BACK -IINIK STS GE0 DELIVERY - 65K SIS - Lv - 3_ FI

MASS ILBSI SUBSYSTEM

STRUCTURE 5952 SUBSYSTEM MASS Ell|S)

StlELL 2308 SIRUCTURE SUPPOR T 1016 FUEL TANK 6_q SHELL & FRAMES 13!;5 OXIDIZER TANK &211 FLAPS 3W; FLAPS $50 FUEL TANK. 3"_1| OXIDIZER TANK 39b

TPS 22£2 THEI_IAL PROTECTION TANK INSULATION 582 SHELL iNSULATION 1680 TANK INSULATION 3hl EXTERNAL TPS I}44 PROPULSION _r071 ENGINE 133 PROPULSION

TVS, PLUMBING 1238 ENGINE 4/h PLUMBING, TVS 54? ACS 9SO ATTITUDE CONTROL 305

F PS 600 ELECTRICAL pOWER 6(,()

AVIONICS f_7 AVIONICS _6611

TOTAL DRY HASS (LBS) 6330* DRY WEIGHT IZOi2

* INCI.UDES IO_ CONTIMGEMCY

57 2.4.2 Initial Configuration Screening to Meet Center-of-Mass Requirements

The aerodynamic configuration selected must, in addition to meeting the external dimensional constraints of the launch vehicle, provide packaging room for the propellant tanks and other subsystems so that the launch configuration with tanks full meets the launch vehicle center-of-mass requirement, Figure 2.4-1_.and the entry configuration with tanks empty meets the center-of-mass requirement to trim the vehicle at the desired angle of attack during the aeromaneuver, Figure 2.4-i_. The desired angle of attack is obtained by placing the entry center-of-mass at the AOTV center-of-pressure location for that angle-of-attack. The selected angle of attack for the baseline "o_c!es wi 11 _ _ _ .... _ • i_ "- . ' _ ...... maximum plane change capability for the vehicle.

58 0 l, o,,,=I,k, ._lid -- ! r,rrr 0 (D _-- Z 0 w I- ,,/ _

Zwr_ e_< _ /1 LU _'0 LU_

_ l 1,uu. / U

._. i\ I I\ -I

_"///_:_X .LN:qAI3UIrlO:1EI NOI.LVO0"I

0 X Z A I.i_1 0

u'l (j-)

u,_ 0

L,13

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_2 Two single stage AOTV configurations were evaluated for a GEO delivery mission from a 65K pound orbiter.

Configuration A of Figure 2.4-1Bis 45 feet long and configuration B is 60 feet long. Other dimensions are given in the notes below the table. Three arrangements are shown of configuration A and two of B.

Evaluation parameters selected are:

XCM/Lv - Center-of-mass (CM) location is critical for the re-entry mission phase to obtain the desired angle-of- attack. CM forward is better.

Available Payload Ba[ Length - Although required payload dimensio-ns have not bee-n specified for this study, it is assumed that a larger payload bay is desireable.

Growth to Return Payload - The critical item for this parameter is the effect on the center-of-mass (CM) location for the AOTV during re-entry, i.e., the AOTV configuration, A-l, returning without payload would have satisfactory CM, but with a returning payload in the aft location the CM would be too far aft for re-entry flight. The A-2 shape with the payload forward is conducive to correct CM location.

Package in Orbiter for CM Requirement - For this analysis the allow-able center-_f---mass range for the orbiter with 65K pounds in the bay is compared to the CM of the AOTV. Configuration A-2 is the only one that can be placed in the orbiter bay nose forward. The others require a nose aft position in the orbiter or rearrange the fuel and oxidizer tanks as mentioned.

Nose Clamshell Door - The complication in structure and me--_anism of a c-_shell door nose on the AOTV to provide for engine firing or payload removal results in a mass penalty.

Side P/L Door - The side payload door structure and mec_anl-sm would be an improvement over the nose clamshell, but is still less desirable than the base end insertion and removal.

Manned Capsule can be Added - Here it is assumed that the manned capsule w--0_id--be added to the payload bay. In the A-I and B-I configuration cases the CM location would be too far aft with a returning manned capsule in the payload location.

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61 This analysis was made to evaluate internal packaging arrangements for a 60 foot long AOTV with an L/D m 1 for a nominal round trip payload of 14,000 pounds to GEO, Figure 2.4-2.

A propellant weight of 71,000 pounds was used. Tank placement an___onfiguration were varied to conserve space in each design. A _ 2°ellipse was used for tank ends, except for the conical/sphere tanks in configurations 1 and 3. Fuel and oxidizer tank locations were varied to obtained a minimum XCM/Lv.

Cu_,f-- ig urations 1-3 used a .o.....=._...... ;_._.. _ °..4.°_.._ ..... Configurations 4 and 5 place a forward firing engine in the nose. The latter arrangement is more space efficient but requires an articulated nose that is open for firing and closed for the remainder of the mission.

For a preliminary design, XCM/Lv limit of 0.6 was assumed. Configurations 1 and 5 just met this limit of XCM/Lv = 0.6; with number 5's returning payload mass in the center of its bay. Configuration 2 is the clear winner for this requirement with a XCM/Lv - .456. This is done by placing the return payload in the nose. During re-entry flight the propellant mass has been reduced to that required to circularize and rendezvous with the orbiter and the reserves and residuals.

Another mass property criterion for the AOTV is based on the capability of the orbiter to handle the center-of-mass of the 100K pound AOTV in its bay. It was assumed that the allowed CM span of 379 to 538 inches aft in the bay for a 65K pound payload would apply to the 100K pound capacity orbiter.

For example, Configuration 1 has a center-of-mass at its 100K pound weight at 513 inches from the nose. This dimension is within the bay allowable span of 379 and 538. Hence the AOTV can be placed nose forward in the orbiter. Configurations 3-5 have CM's that require nose-aft placement in the orbiter bay.

Payload bay length available for the configurations is listed. The payload bay length of 8 feet for configuration 1 is the shortest. Shape 2 has the longest length of 28 feet but the diameter is reduced by its nose location. Configuration 5 with a payload bay length of 25 feet has the largest volume.

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63 Satellite designs are driven by significant cost drivers and market opportunities, Figure 2.4-3. One of these is the cost of transport from Earth to LEO. Shuttle launch charges drive satellite designers toward short, 15' diameter satellites. European marketing opportunities limit some satellites to 12' diameters (for Ariane compatibility). Both 12' and 125' satellite diameters exceed the intern_l payload bay capacity of a very long (60') biconic AOTV with a 2 half cone angle on the aft frustum. Consequently, for nearly all biconic vehicles, some provision will have to be made to attach payloads to the exterior of the AOTV, since _ost payloads will not fit inside. Obviously, half cone angles <2 should be evaluated.

64 Figure 2.4-3 SOME AOTV REQUIREMENTS -- PAYLOADS

FULL SIZE PAYLOADS

• SATELLITE DESIGNERS MINIMIZE LAUNCH COSTS - FULL 15 FT DIAMETER MINIMIZES LAUNCH LENGTH

• SOME MISSIONS REQUIRE LONG (> 30 FT) SATELLITES

• INTERNAL PAYLOAD CAPABILITY OF 60 F_" LONG AOTV * - MAX DIA 11 TO 13 FT - MAX LENGTH 12 TO 21 FT

CONCLUSION

AOTV MUST HAVE CAPABILITY OF SUPPORTING PAYLOADS OUTSIDE OF AEROSHELL

Figure 2-4.4 SATELLITE RETRIEVAL

• NO SCIENTIFIC OR COMMERCIAL REQUIREMENT TO RETURN A SATELLITE TO LEO FROM HEO

• MILITARY REQUIREMENTS UNKNOWN

CONCLUSIONS

1 PAYLOAD SPACE WITHIN AEROSHELL MUST ONLY ACCOMMODATE

A) SAT SERVICING EQUIPMENT WHICH IS RETURNED TO LEO

B) MANNED CAPSULE FOR SAT SERVICING

2 PAYLOAD SPACE LARGER THAN NEEDED BY IA & 1B) IS EXPENSIVE (SHORT VEHICLES ARE CHEAPER)

65 The issue has been addressed of whether a need exists for a payload bay within an AOTV, Figure 2.4-4. Since external payload provisions must be provided, payloads on the way out (LEO to GEO) do not require an internal payload bay. An internal payload bay is required of any payload which must survive the severe heating developed during atmospheric entry near the end of an AOTV mission.

The cost of returning a payload from a High Energy Orbiter (like GEO) is quite substantial. Principly, this cost is the transportation cost of delivering sufficient propellant (to LEO and then to GEO) to visit and de-orbit the retrieved satellite. If the AOTV mission on a particular flight is only satellite retrieval, the total cost of _etrieval approximates the cost of 1 STS launch (-_$84M) for long, large diameter AOTVs which are capable of retrieval. This $84M far exceeds the cost of most commercial satellites to date. A commercial need for a refurbished satellite would be better met by building a new satellite than by expending $84M for retrieval. This cost burden is also considered excessive for any scientified satellites which may be in a HE0.

Much lower retrieval costs are possible if the AOTV performs a satellite delivery mission in the early part of a retrieval flight. The propellant necessary for retrieval (e.g., 2400 ib to return a 6000 ib satellite) occupies space (and weight) that another paying customer (satellite) could use. Since a long, large diameter AOTV that is capable of retrieval leaves only a small amount of payload bay length for payloads, the 2400 ib of propellant occupies a place of one out of 2 or 3 payloads. Consequently, the mission cost of retrieval approximates the los income from one of two, or one of three, payloads which would have shared the total launch cost. This implies a minimum cost of retrieval of $84M .-+-"3 = $28M, when the retrieval mission is combined with a satellite delivery mission. This cost is also judged excessive for a commercial venture.

On those rare oGcasions when satellite retrieval is required, the mission can be flown much less expensively by using a short AOTV (without an internal payload bay) in an all propulsive mode. The additional propellant capacity needed for this mission can be stored in external tanks which can be re-used on other missions.

The only payload that must be retrieved, and survive the atmospheric maneuvering phase of a mission, is a crew capsule. However, the requirements of manned missions may drive a biconic AOTV so uniquely that a manned mission capable vehicle may be unsuited for the competitive mission of delivery of commercial satellites. Consequently, our AOTV designs for the satellite delivery missions will not contain internal payload bays.

66 Biconic AOTVs fly at fixed angles of attack in pitch. To maximize vehicle maneuverability, it is desirable to fly at the angle of attack which maximizes the ratio of Lift/Drag (L/D). To obtain this, the vehicle center of mass (CM) must coincide with the vehicle center of pressure (Cp).

Some amount of control of the Cp location is possible by using flaps to extend the vehicle on the windward side. However, CM variations can be very substantial. If a single vehicle is to perform all AOTV missions (a "universal AOTV"), the entry weight could vary from i0,000 ib to 24,000 lb. To enable such a substantial weight change without altering the AOTV's aerodynamic characteristics, it is recommended that the payload CM be located at the AOTV CM, which is at the same location (along the vehicle length) as the AOTV's Cp, Figure 2.4-5.

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6_ Some characteristics of a possible universal biconic AOTV are outlined in Figures 2.4-6, 7 and 8. The payload bays on both the single engine and two engine versions are adequate to contain a Bare Bones crew capsule. External payload attachment is possible on the aft end "docking ring". Internal payload capability is also available. The engines fire in the forward direction, after a deployable nose segment has been rotated out of the way.

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/OPEN POSITION OF AOTV NOSE 71 Some of the advantages of short, ground based AOTVs are outlined in Figure 2.4-9. For example, a 3 foot length reduction will lower Shuttle launch charges by $300M in five years.

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73 Even with space basing of OTVs, short AOTVs still offer economic advantages when they are used in a ground based mode, Figure 2.4-10.

A single space base at 28 ° inclination implies that high inclination mission will be ground based. Also, prior to total storage of a space based AOTV at the Space Station, a number of ground launched AOTV flights will occur while the technology of operating in a space based mode is being developed and verified. As many as 50 ground based flights may be required while the Space Station learns how to perform payload manifesting and transfer, propellant manifesting, storage and transfer, and AOTV inspection.

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75 2.4.3 Unmanned/Manned Vehicles Schedule Implications

Past OTV studies have planned on a 4 to 5 year period of unmanned AOTV flights before starting operation with a manned AOTV, Figure 2.4-11. Present expectations are for a manned space station to be operational several years before the first new reuseable OTV becomes operational. With a significant manned presence in space, manned OTV missions may be desired immediately, or shortly after, a new reuseable OTV becomes operational. Consequently, the initial AOTV may be a manned (or man rated) vehicle, Figure 2.4-12.

76 Figure 2.4-11 SOME AOTV REQUIREMENTS: SCHEDULE IMPLICATIONS

PAST STUDIES

BOEING t}• ABOTV IOC 1986-87 _ 5 YEARS OF

1980 t• MANNED ABOTV IOC 1991-921 UNMANNED FLIGHTS

EVOLUTIONARY VEHICLE DEV'T /

DYNAMICS • MANNED OTV lOG 1992 UNMANNED OENERAL1980 l"• 0TVAOTV OC'"IOC? (1992) I FLIGHTS YEA'SO,

Fi .qure 2. '_-12PRESENT EXPECTATIONS :

• SPACE STATION IOC 1990- 91

- PEOPLE CONTINUOUSLY IN SPACE DOING WORK

• UNMANNED AOTV IOC 1992 - 93 I _ TIME TO PROVE • MANNED AOTV IOC 1993- 9_? _ SAFETY OF AOTV

- MISSIONS ARE UNCERTAIN, TIMING IS UNCERTAIN

• SERVICE COMMERCIAL AND/OR GOVERNMENT SATELLITES

• SERVICE NATIONAL ASSETS (SBR, LARGE PLATFORMS, ETC.)

• CONSTRUCT/ASSEMBLE NATIONAL ASSETS

CONCLUSIONS:

1. INITIAL AOTV MAY BE MAN RATED AND INITIALLY FLOWN UNMANNED

2. DEFINE ONE AOTV FOR ALL MISSIONS

3. EXAMINE EVOLUTIONARY EXPANSION OF AOTV CAPABILITY WITH DIFFERENT VEHICLES

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QRUMMAN DERIVATION OF BARE BONES CREW CAPSULE WEIGHT FROM LM ASCENT STAGE

LM 12 ASCENT STAGE BARE BONES CAPSULE AOTV AVIONICS

(LB) (LB) (LB) STRUCT F, MECHN 1406 LESS TANKS ETC 1100 STAB 6 CONT 189 20t IN CAPSULE 30 151 NAV 8 GUIDANCE 191 IIO1&IN CAPSULE 76 115 CREW PROVISIONS 175 LESS LUNAR EQUIP 'Iq6 ECLS 297 297

ILLUMINATION (INSTRUMENTATION) 132 132 CO EPS 737 qo'i, IN CAPSULE 295 _q2 PROPULSION q69 REACTION CONTROL 2q2 COMMUNICATION 11q 50't, IN CAPSULE 57 57

CONT. f, DISPLAYS 232 232 EXPLOSIVES 29

MANUFACTURERS VARIATION [-60) D qt 53 LB 2373 LB 765 (REF)

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The baseline state-of-the-art thermal protection subsystem is of the Reusable Surface Insulation type (RSI and FRCI) bonded to a Strain Isolation Pad (SIP) where required with a silicone based adhesive. The baseline TPS thickness (weight) requirements have been determined for a design criteria of I) initial temperature of 100 v and 2) maximum bond line temperature of 350 ° . Thickness requirements were determined, Figure 3.3-1 through use of the Reaction Kinetics Ablation Program (REKAP) code (37). Thermal conductivity as a function of both temperature and pressure was employed. Thermal soak-out of the AOTV TPS occurs during the lift out of the atmosphere at near vacuum pressures in contrast to the Space Shuttle Orbiter where soak-out occurs at near one atmosphere pressure.

The total weight of the thermal protection subsystem can be reduced significantly through three separate means: i) thermaloconditioning of the TPS prior to entry by cold-soaking it to -100 , 2) reduction in the weight of the protective c_ating, 3) permitting the structure/bond line to soak out to 600 v.

It has been estimated that thermal conditioning prior to entry could reduce the RSI weight requirement by as much as 23%, reducing the protective coating by 50% could result in a reduction of the total TPS weight by 9% and increasing to 600 ° the allowable maximum structure bond line temperature could . result in a TPS weight reduction of 37%.

Estimates have been made of the maximum surface temperatures expected on the nose and aft frustum areas of this class of vehicles, Figure 3.3-2 and 3. The convective heating rates are for equilibrium flow and a fully catalytic surface. The non catalytic nature of the TPS coating and the non-equilibrium nature of th_ flow are expected to produce maximum temperatures _00-200 _ lower than those quoted for the aft frustum and up to 400 lower for the nose. Estimates of eqilibrium hot gas radiation, using Pages results (13) demonstrated it to be a second order effect, so it has been neglected in this study..

Current state-of-the-art, normal growth and accelerated growth increases in peak allowable surface temperatures have been specified by Goldstein and Curry (34 and 35), Figure 3.3-4. These have been compared to the predicted maximum surface temperatures for these AOTVs.

It can be observed that most of the aft frustum maximum surface temperatures fall within the current capability of FRCI,

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201 with the exception of L/D = 1.5 vehicles returning from 5 x GEO, where the 2500 v temperature capability is exceeded due to probable turbulent flow if single pass capture is used. On the nose, due to the relatively small nose radius, maximum surface temperatures are expected to exceed even the 4000 accelerated growth TPS capability for most of the missions involving one pass capture.

Four different approaches to rescue the nose temperature have been investigated and include: i) alternate steering laws, 2) higher angles of attack, 3) multiple rather than single atmopsheric passes, and 4) a larger nose radius. Evaluations to date have considered entry at (L/D) maximum only for the purpose of producing maximum aerodynamic plane change. The convective heat flux reductions possible by opeorating toward the one pass capture bound rather than going for maximum plane change, are illustrated in Figure 3.2-2.

Significant reductions in peak nose surface tempeorature can also be obtained by flying at an angle of attack greater than that for (L/D) max. and accepting less plane change. By flying at this larger angle of attack, larger lift and drag coefficients are obtained and thus the vehicle can operate at a higher minimum altitude while reducing the peak heat transfer rate experienced. The net result is a decrease in maximum nose temperature and a small increase in maximum windward surface temperature due to the higher angle of attack.

It has been demonstrated in the AMOSS Studies (6) that use of multi pass rather than single pass capture significantly reduces the maximum heat transfer rates experienced and hence maximum surface temperatures, while the total plane change obtained aerodynamically remains the same as for a single pass. These results can be combined to determine those combinations of steering law and number of passes to limit the nose cap to temperatures acceptable fo'r reuse. These results are summarized for return from 5X GEO and 6 hr polar in Figures 3.3-5 and 6 where it can be seen that for a given allowable temperature limit, the magnitude of aerodynamic plane change achievable (and hence, delivered payload.) increases with number of a_mospheric passes permitted. An L/D = 1.5 AOTV, flying at a 15 angle of attack, will require 4 atmospheric passes to a_hieve nearly all of its maximum aerodyanamic plane change (--_26 v) capability if nose temperature is limited to 4000 ° .

The effect of increasing the nose radius has also been investigated. For a given steering law and angle of attack, the required nose radius to reduce the maximum nose temperatures to those required for reusable passive TPS for a one pass capture have been estimated. Increasing the nose radius also reduces the vehicle L/D due to increased drag, thus allowing the vehicle to operate a a higher altitude, again reducing the peak heat transfer rate experienced, but reducing the magnitude of the plane change obtained. For example, if two passes are required

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203 in order to meet temperature limits for the reference vehicle with a nose radius of 2 feet, then by doubling the nose radius to about 4 ft., the number of passes may be reduced to one.

If the combination of aerodyanamic plane change and number of atmopsheric passes required is unacceptable, then alternate approaches of either transpiration cooling, heat pipe, or use of a refurbishable thermal protection material will be required. Transpiration cooling using either a liquid or gas injectant has been flight qualified on small nose radius (<2 inches) vehicles for some time. Development effort is current_underway on much larger systems. Heat pipe designs for noses _nd leading edges have been developed and ground test in previous studies (38, 39, 40).

For a ground based AOTV, an alternate approach to a transpiration or heat pipe cooled nose or a multipass reusable RCC nose is the use of a refurbishable state-of-the-art low density ablator. The potential for downstream TPS surface contamination from ablation products would have to be considered.

204 3.4 Structure Subsystem The AOTV baseline state-of-the-art structural subsystem consists of the following elements: i) a graphite epoxy structure, 2) aluminum cryogenic fuel and oxidizer tanks, and 3) split body flaps for use in center-of-mass offset management. Several recent trade studies and surveys have been conducted to determine the value of structural material advances and design/property improvements for advanced space vehicles. An organized activity is underway, spearheaded by the "AIAA Composite Structure Subcommittee" composed of members from NASA, Government agencies, and manufacturing firms. Their recent survey revealed composite structure design _llowables were obtained from a wide variety of test methods and many different failure criteria were employed. The committee is striving for:

l) Unified testing to determine allowables - consensus is to use uniaxial data and analysis to determine composite properties.

2) Unified failure criteria for design of composite structure.

We recommend that AOTV beocme an advocate for this organized activity. It is projected that a i0 to 30% weight reduction of AOTV structural weight is possible with an improved definition of design properties.

Numerous advanced structural materials are under investigation/development and include graphite composites, metal matrix composites - and doped'and undoped aluminum lithium alloys. It is projected that additional structural weight reductions are possible by use of metal matrix materials reinforced with graphite whiskers (14). The higher temperature limits (e.g., 600_F-1000°F) of these materials will provide additional reduction in TPS weight.

205 3.5 Avionics Subsystem The reference state-of-the-art avionics subsystem weight of 600 pounds is driven primarily by degree of autonomy and redundancy. The future OTV technology study, (15), projected 50% weight reduction due to normal growth technology employing laser gyros and data bus. Our evaluation of an expandable fly-by-wire delivery stage design suggests up to 70% weight reduction is possible if avionics autonomy could be sacrificed. Numerous hardware and software advanced development activities are underway, e.g., spacecraft flight computers are currently under development that can process 450,000 instructions per second with up to one million words memory (16 bit). Hence, significant reductions in avionics subsystem weight appear possible.

3.6 Flight Control Subsystem

Atmospheric roll control using a hybrid combination of Reaction Control System (RCS) and aero surfaces for trim control or additional control authority is expected to provide the most attractive Flight Control actuation approach. Attitude sensing and acceleration information will be provided by inertial sensors which may also serve as instrumentation sensors or in some cases, provide data during other phases of the mission.

These conclusions are based on previous studies (36) which evaluated the mid L/D ae

206 The information from the inertial measuring sensors will provide inputs for the flight path/attitude algorithm which will provide vehicle position, velocity, attitude and attitude rate which are used for atmospheric flight path control and analytic drag guidance°

Flight path control is provided by the attitude control system through the command of either gas jets or aerodynamic control surfaces for roll modulation of the lift vector. The RCS is required for exoatmospheric attitude contol and is therefore part of the AOTV, but attempts to adapt for atmospheric roll control, using two level thrustor or pulse width modulation have resulted in heavier systems without comparable payoff in accuracies. Aerosurfaces can provide the roll control for additional weight penalty, but also augment pitch and provide damping.

Additional benefits resulting from the combined reaction and aero torque approach include: (I) the potential to design with some margin for redundant contingency control modes, (2) potential to extend control authority to the pitch and yaw axes to provide damping, or if found necessary, full 3-axis control and (3) potential for tighter attitude controls during atmospheric operation and entry/exit transition regimes.

The roll control subsystem providesrequired bank-reverse-bank roll command logic and torques to modulate lift in the plane to maintain Dref, while minimizing the out-of-plane components of lift error, to minimize exit state errors.

Bank angle commands provide the desired in-plane lift with modulation for the in-plane magnitude resulting in out-of-plane components which are nulled by commanding roll attitude sweeps of the fixed trim in both plus and minus directions.

Figure 3.6-1 shows a block diagram of a possible hybrid flight control using combined RCS and Split Windward Flap (SWF). The in-plane lift is modulated for altitude adjustments to maintain constant Dref. Primary lift magnitude is derived from basic aeroconfiguration with a percentage, or /k-lift, lift control authority available from the flaps.

Roll control must cause the vehicle to follow commands from the guidance law in a stable manner. Studies have shown the strong sensitivity of control sizing to upsetting roll disturbance torques which arise primarily from mass property asymetries resulting from loads, equipment or manuacturing tolerances. Out-of-plane angle of attack can result from asymmetries which contribute to error build-ups in out-of-plane lift.

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208 In addition to cg offsets, variations in the aerodynamic center-of-pressure (CP), from design, can affect the aero performance and resulting control authority. Combination of CG and CP variations and the resulting static margin changes also impact control authority.

Advantages of a hybrid roll control approach are: i) improved accuracy in attitude control 2) blending during approach or entry dynamics 3) blending during exit or departure dynamics, and 4) augmented lift control authority.

Transition scheduling from full RCS roll control in the rarefied regime to predominant SWF atmospheric roll control, i.e., blending, will be controlled in the digital computer.

209 3.7 Aerospace Support Equipment [ is illustrated in Figure 3.7-i. i The preferred general arrangement lfor transmitting loads between the Orbiter payload bay and AOTVY Vertical (ZZo) loads are removed at 4 locations on the Orbiter longerons. Lateral (+Yo) loads are removed at an orbiter keel fitting at the bottom o7 the bay. Fore (-Xo) and After (+Xo) loads are removed at the same 2 aft longeron location_that carry vertical loads.

_N_X; Mmnl_vmmnF _n _n_c_ iK a two phase operation. First, the two +Yo longeron fittings and the keel fitting release the AOTV. The AOTV is then rotated by the orbiter RMS about the xo axis on the remaining -Yo longeron fittings until it is outside the payload bay (about 90 ° of rotation). In this position, the AOTV can be checked out. For manned AOTVs like Figure 2.4-21, half the aeroshell can be opened to expose the crew capsule. The crew capsule can be rotated out of the AOTV to permit attachment of a "transfer tunnel" from the orbiter cabin. This allows shirtsleeve transfer of the crew to the AOTV. The second phase of deployment occurs after the AOTV is closed and prepared for flight. At that time, the two -Yo longeron attachments release the AOTV and springs at the attachments propel the AOTV away from the Orbiter. Another mode of deployment could use the orbiter RMS for a separation impulse.

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211 A ROM preliminary design of the Airborne Support Equipment (ASE) structure, which attachs the AOTV to the orbiter, indicates that a very lightweight structure is possible, Figure 3.7-2. The baseline structural element is a hollow steel tube with a diameter/thickness ratio of 20. Applying the utlimate load factors shown on the facing page to a 61,300 ib AOTV (H-IM when fully loaded), and using modest strength steel (FTu = 180,000 psi),the four longeron tubes and one keel tube weigh about ii0 lb. Adding weight for mechanisms to permit deployment motions and contro!_ the total structural system weight for holding (and releasing) a large AOTV within the Orbiter is 270 pounds.

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213 A review of orbiter longeron bridge fittings (from JSC07700) shows 14 different sizes, with an average weight of 141 pounds. Since this AOTV installation has the aft fittings near the maximum density region of the AOTV (the LOX tank), the aft fittings are highly loaded. Since the aft are highly loaded, the forward longeron fittings are lightly loaded. Since the number of highly loaded fittings is equal to the number of lightly loaded fittings, an average fitting weight can be assumed for all. Four average longeron bridge fittings weigh 564 pounds.

The 12 different keel fittings of JSC 07700 have an average weight of 201 pounds, and a standard deviation of 53 pounds. Since the keel fitting is very highly loaded in this AOTV application (H-IM), a keel fitting weight of 254 pounds was chosen. This is very close to the heaviest keel fitting in the list.

After adjusting the total of all longeron bridge and keel fittings weights for the 675 ib allowed all payloads (at no charge), the net bridge fitting weight penalty charged to the ASE for support of a 61,000 ib vehicle is 143 pounds, Figure 3.7-3.

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215 A weight estimate was prepared for the hardware elements which are necessary to dump 45,000 ib of LOX/LH propellants, from the AOTV to discharge ports on the orbiter, within 5 minutes, Figure 3.7-4. The major weight advantage of this system, when compared with currently planned, systems, is the single large (5' diameter) tank for storage of high pressure helium gas. The current state-of-the-art filament wound tank saves many hundreds of pounds over multiple small bottles.

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217 Subsystem elements of the Airborne Support Equipment for two programs are compared in Figure 3.7-5. Centaur G' (a preliminary definition from the Centaur Project Office) and a large biconic AOTV (typically, H-IM).

The first four lines of the table have been discussed previously (for AOTV). Residuals are virtually the same for both vehicles. A large weight difference occurs in Avionics. Centaur G' has extensive instrumentation to monitor the pressure stabilized bulkhead, as we!l as a number of batteries. The AOTV is intended to be nearly autonomous, with a significant on-board health monitoring system. Consequently, some of the Centaur ASE Avionics are on board the AOTV as part of the vehicles Avionics weight. The i00 ib allocated to AOTV ASE Avionics was estimated sa satisfactory for an autonomous vehicle, the last entry in the _OTV table is for an orbiter stored collapsable tunnel to permit shirtsleeve transfer of personnel from the orbiter cabin to the crew capsule of a manned AOTV.

The substantially lighter weight of the AOTV ASE (-6500 ib) allows more propellant to be loaded in ground based AOTVs. This allows higher payload weights delivered to GEO.

218 i

..I w z z

_C w

uJ Z Z 0 mr uJ A

219 3.8 Electrical Power Subsystem

The AOTV baseline state-of-the-art electrical power subsystem consists of the following elements: i) two 3.5 KW STS orbiter fuel cells with reactants, 2) back-up batteries, 3) tanks and plumbing, 4) power conditioning equipment, electrical harness and equipment supports.

The reference state-of=the-art electrical power subsystem weight is 600 pounds. It is projected that use of non-metallics like graphite in the power section to replace magnesium and nickel and continuing improvement in battery technology and other _=_...... ^_ __,,=...._u_y_,,,._ll_ ..... result in 20 to 38% wei ght saving. Currently the NASA sponsored OAST "NASA Regenerative Fuel Cell Program" at JSC and Lewis has this task in their program, but it is unfunded. It is recommended that AOTV become an advocate for task funding.

220 4.0 REFERENCES

1. Austin, R.E., Cruz, M.I., and French, J.R., "System Design Concepts and Requirements for Aeroassisted Orbital Transfer Vehicles", AIAA Paper 82-1379, AIAA 9th Atmospheric Flight Mechanics Conference, San Diego, CA, August 1982.

• Walberg, G.D., "A Review of Aeroassisted Orbit Transfer", AIAA Paper No. 82-1378, San Diego, CA, August 9-11, 1982.

. The Boeing Company, "Orbital Transfer Vehicle Concept Definition Study" Report NO. D180-26090, Volumes i-6, 1980.

• General Dynamics - Convair Division, "Orbital Transfer Vehicle (OTV) Concept Definition Study", Report No. GDC- ASP-80-012, Volumes 1-6, 1981.

. Letts, W.K. and Pelekanos, A., "Aeroassisted Orbital Transfer Mission Evaluation", AIAA Paper 82-1380, AIAA 9th Atmospheric Flight Mechanics Conference, San Diego, CA, August 1982.

, Andrews, C.D., "Feasibility and Tradeoff Study of an Aero- maneuvering Orbit-to-Orbit Shuttle (AMOOS)", LMSC-HREC TR D306600, June 1973.

° Boyland, R.E., Sherman, S.W., and Morfin, H.W., "Manned Geosynchronous Mission Requirements and Systems Analysis Study", NASA CR-160429, Grumman Aerospace Corporation, Bethpage, NY, November.1979.

. Gentry, A.E., Smith, D.N., and Oliver, W.R., "The Mark IV Supersonic_Hypersonic Arbitrary-Body Program", Volumes I, II, and III, AFFDL-TR-73-159, November 1973.

. Program Development, NASA Marshall Space Flight Center, "Orbit Transfer Systems with Emphasis on Shuttle Applica- tions - 1986-1991" NASA TM X-73394, 1977

10. Aerojet Liquid Rocket Company, "Orbit Transfer Rocket Engine Technology Program", Program Review NAS3-23170, 23 February 1983.

11. Pratt and Whitney Aircraft, "Orbit Transfer Rocket Engine Technology Program - Final Review", conducted for NASA- Lewis Reserach Center under Contract NAS3-231721, February 1983.

221 12. Rocketdyne, "Orbit Transfer Rocket Engine Technology Program", NAS3-23172, Final Review, February 22-23, 1983.

13. Page, W.A., Compton, D.L., Borucki, W.J., Ciffone, D.L., and Cooper, D.M., "Radiative Transport in Inviscid Non- adiabatic Stagnation Region Shock Layers", AIAA Paper 68- 784, Los Angeles, CA, 1968; also AIAA Selected Reprints/ Volume VII - Radiative Gas Dynamics, edited by R.J. Goulard, June 1969, pp. 51-65; also AIAA Progress in Astronautics and Aeronautics: Thermal Design Principles of Spacecraft and Entry Bodies, Volume 21, edited by J.T. Bevans, Academic Press, New York, 1969, pp. 75-114.

14. Trent, DoJ., Testarmata, J.P. and B[ooks, J.L., "Advanced Materials for STS Applications", Rockwell Int. Doc #STS 83-1046, presented at NASA/LaRC Symposium on Advances in TPS and Structures for Future Space Transportation, Dec. 1983.

15. Davis, E.E., "Future Orbital Transfer Vehicle Technology Study", NASA CR3535, May 1982.

16. W.R. Letts and A. Pelekanos, "Aeroassisted Orbital Transfer Mission Evaluation", Paper No. 82-1380, AIAA 9th Atmospheric Flight Mechanics Conference, San Diego, Cal., August 9-11, 1982.

17. M.I. Cruz, "The Aerocapture Vehicle Mission Design Concept", Paper No. 79-0893, NASA/AIAA Conference on Advanced Technology for Future Space Systems, Langley Research Center, Hampton, VA., May 8, 1979.

18. J.T. Betts, "Optimal Three-Burn Transfer", pp. 861-864, Vol. 15, No. 6, June 1977 AIAA Journal.

19. F.W. Gobetz and J.R. Doll, "A Survey of Impulsive Trajectories", pp. 801-834, Vol. 7, No. 5, May 1969 AIAA Journal.

20. Scott, C.D., et al, l"The Aerothermodynamics and Thermal Protection Ssytem Challenge for an Aerobraking Orbital Transfer Vehicle", presented at the NASA Symposium on Recent Advances in TPS and Structures for Future Space Transportation Systems, Dec. 13-15, 1983.

21. Shinn, John, J.J. Jones, "Chemical Nonequibrium Effects on Flowfields for Aeroassist Orbital Transfer Vehicles", AIAA Paper #83-0214, presented at AIAA 21st Aerospace Sciences Meeting, Jan. 10-13, 1983, Reno, Nevada.

222 M

22_ Talay, T.A., N.H. White, and J.C. Naftel, "Impact of Atmospheric Uncertainties and Viscous Interaction Effects on the Performance of Aero-Assisted Orbital Transfer Vehicles", AIAA Paper #84-0408, presented at AIAA 22nd Aerospace Sciences Meeting, Jan. 1984, Reno, Nevada.

23. wilhite, A.W., J.P. Arrington, and R.S. McCandless, "Performance Aerodynamics of Aero-Assisted Orbital Transfer Vehicles", AIAA Paper #84-0406, presented at AIAA 22nd Aerospace Sciences Meeting, Jan. 1984, Reno, Nevada.

24. NASA Space Shuttle System Payload Accommodations, JSC 07700, Vol. XIV, Rev. G, Sept. 1978.

25. Gilbert , L . , and Goldberg, L., "A Reynolds Number Scaling Theory for Hypersonic Ablation", AIAA Paper 67-155, Jan. 1967.

26. Scott, C.D., "Effects of Nonequilbrium and Catalysis on Shuttle Heat Transfer", AIAA Paper #83-1485, presented at AIAA 18th Thermophysics Conference, June 1-3, 1983, Montreal, Canada.

27. Moss, J.N., and G.A. Bird, "Direct Simulation of Transitional Flow for Hypersonic Re-entry Conditions", AIAA Paper #84-0223, presented at AIAA 22nd Aerospace Sciences Meeting, January 9-12, 1984, Reno, Nevada.

28. Zoby, E.V., et al, "Temperature Dependent Reaction-Rate Expression for Oxygen Recombination at Shuttle Entry Conditions", AIAA Paper #84-0224, presented at AIAA 22nd Aerospace Sciences Meeting, January 9-23, 1984, Reno, Nev.

29. Powell, R.W., et al., "Performance Evaluation of the Atmospheric Phase of Aeromaneuvering Orbit Tranfer Vehicles", AIAA Paper #84-0405, AIAA 22nd Aerospace Sciences Meeting, January 9-12, 1984, Reno, Nevada.

30. Cruz, M.I., "Trajectory Optimization and Closed Loop Guidance of Aeroassisted Orbital Transfer", AIAA Paper #83-413, presented at AAS/AIAA Astrodynamics Specialist Conference, Lake Placid, New York, August 22-25, 1983.

Hill, 0., NASA JSC, personal communication, April 1983.

Gamble, M.J. and L. Skalecki, NASA JSC, personal communication, April 1983.

33. Meese, K.D. and N.X. Vinh, "Minimum-Fuel Aeroassisted Coplanar Orbit Transfer Using Lift Modulation", AIAA Paper #83-2994, presented at the AIAA Atmospheric Flight Mechanics Conference, August 15, 1983, Gatlinburg, Tenn.

223 34. Goldstein, H., NASA ARC, personal communication, 17 Jan. 1983.

Curry, D., NASA JSC, personal communication, 2 June 1983.

Armento, R., "Mars Aerocapture Vehicle Definition Study Final Report", GE Report 79SDR2258, 17 August 1979.

37. Cline, P.B., and F.B. Schultz, "Investigation of the Effect of Material Properties on Composite Ablative Material Behavior, NASA CR-72142, April 1967.

38. Niblock, G.A., Reeder, J.C., and Huneidi, F., Four Space Shuttle Wing Leading Edge Concepts. J. Spacecraft and Rockets, Vol 41 No 5 May 1974, pp. 3i4-_ZU

39. Camarda, Charles J.: Analysis and Radiant Heating Tests of a Heat-Pipe-Cooled Leading Edge. NASA TN D-8468, 1977.

40. Camarda, C.J., "Aerothermal Tests of a Heat-Pipe-Cooled Leading Edge at Mach 7", NASA Technical Paper 1320, Nov. 1978.

224