N O T I C E This Document Has Been Reproduced From

N O T I C E

THIS DOCUMENT HAS BEEN REPRODUCED FROM MICROFICHE. ALTHOUGH IT IS RECOGNIZED THAT CERTAIN PORTIONS ARE ILLEGIBLE, IT IS BEING RELEASED IN THE INTEREST OF MAKING AVAILABLE AS MUCH INFORMATION AS POSSIBLE i^ G DC -A S P-80-010 CONTRACT NAS8-33527

N81-11101 (NASA-C13-101594) LOW THRUST VEHICLE CONCEPT l STUDY Final Report (General Dynamics Corp.) 291 p HC A13/MF A01 CSCL 22B f Unclas G3/18 29132

LOW THRUST VEHICLE CONCEPT STUDY

GENERAL DYNAMICS Conveii Division Prepared under Contract NAS8-33527, Task 7

Prepared by Advanced Space Programs GENERAL DYNAMICS CONVAIR DIVISION San Diego, California FOREWORD

This report aocunents the results of Contract NAS8-33527, Task 7 — "Loa, Thrust Vehicle Concept Study". This study was conduced over a 9-month period from September 1979 to May 1980. The NASA/MSFC Program Manager was D. R. Saxton. The General Dynamics Program Manager was W.J. Ketchum.

In addition to the program managers, many General Dynamics Convair person- nel contributed to the study. The key individuals and their contributions are:

P. J. Griff1th Computer/Simulation R. Ma.ffucci Propellant Acquisition S. Mald Avionics/Power F. Merion Thermodynamics C. D. Pengelley Structural Dynamics R.C. Risley Costs, Schedules E. I. Seiden Weights L. E. Siden Conceptual Designs P. Slysh Structures/Synthesis

All data in this report are presented in the English System.

PRECEDING PACE SI ANK NOT FILMED

iii TABLE OF CONTFNTS

Section Page

1 INTRODUCTION ...... 1-1 1.1 Objectives ...... 1-1 1.2 Study Approach ...... 1-1

2 MISSION/PAYLOAD DEFINITION ...... 2-1 2.1 Potential Missions/Payloads for Low-thrust Propulsion . . . .2-1 2.2 OTV Requirements ...... 2-2 2.3 Selected Payload Characteristics...... 2-2 2.3.1 Space Based Radar . . . . .2-4 2.3.2 Geoplatform ...... 2- 10 2.3.3 Power Array , , . , , ...... 2-11 2.3.4 External forces and torques at LEO...... 02-12

3 CANDIDATE LOW-THRUST PROPULSION SYSTEM CONCEPTS, . , . 3-1 3.1 OTV Characteristics. , , , , , . , , , , . , 03-1 3.2 Propulsion Characteristics , , . , . . . . . , .3-1 3.2.1 Thrust Transient ...... 3-1 3.2.2 Distributed Thrust ...... 3-4 3.2.3 Thrust Vector Coni rol. , , ...... , .3-5 3.2.4 Exhaust Plume Interaction , , . , . . . , . , . 3-6 3.2.5 Engines ...... 3-6

4 PERFORMANCE ANALYSIS ...... 4-1 4.1 Computer Synthesis/Optimization ...... 4-2 4.2 Payload Configurations ...... 4-3 4.3 Payload- OTV-Shuttle Length Fit...... o4-4 4.4 OTV Characteristics...... 4-4 4.4.1 Velocity Requirements vs. Initial Thrust-to -Wei rht . .4-5 4.4.2 Isp vs. Thrust...... 4-6 4.4.3 OTV Length vs. Propellant Weight ...... 4-6 4.4.4 Mass Fraction vs. Propellant Weight ...... :4-6 4.4.5 Mission Losses ...... 4-6 4.4.6 Summary of OTC' Parameters ...... 4-8 4.5 Payload Characteristics...... 4-8 4.5.1 Space Based Radar - Tetrahedral Truss Arm (SBR-A) .4-8 4.5.2 Space Based Radar - Tetrahedral Truss Ring (SBR-R) .4-9 4.5.3 Geoplatform ...... 4-10 4.6 Shuttle Characteristics . , ...... 4-10

PRECEDING PAGE BLANK NOT FILMF-^ v TABLE OF CONTENTS (Contd)

Section Page

4 4.7 OPTOTV Computer Program ...... 4-10 (Contd) 4.7.1 Missions ...... 4-11 4.7.2 Growth Capability ...... 4-11 4.8 Results ...... 4-12 4.8. 1 Space Based Radar-A Analysis Results ...... 4-14 4.8.2 Space Based Radar- R Analysis Results ...... 4-24 4.8.3 Geoplatform Analysis Results ...... 4-30 4.8.4 Summary Analysis Results • • • • • • • . • • 4-36

5 BASELINE VEHICLE CONCEPT IDENTIFICATION AND DEFINITION . 5-1 5.1 Baseline Configuration, Description, and Weight ...... ^-1 5.2 Subsystems * ...... 5-6 5.2.1 Torus L02 Tank ...... 5-6 5.2.2 Propellant Acquisition ...... 5-12 5.2.3 Insulation ...... 5-29 5.2.4 Pressurization, Tank Pressure Control, and Abort Dump ...... 5-11 5.2.5 Engine Feed Ducts ...... 5-51 5.2.6 Fill and Drain ...... 5-51 5.2.7 Propellant Utilization . _ . . . . 5-51 5.2.8 Auxiliary Propulsion/Attitude Control ...... 5-52 5.2. 9 Avionics/Power ...... 5-53 5.3 Installation in Shuttle ...... 5-59 5.3.1 Adapter Structure ...... 5-59 5.3.2 Helium Storage ...... 5-59 5.3.3 Plumbing ...... 5-59 5.3.4 Deployment ...... 5-61

6 PROPULSION/SUBSYSTEM TECHNOLOGY REQUIREMENTS . . . . 6-1 6.1 Torus Tank ...... 6-1 6.2 Low Thrust Engine ...... 6-4

7 COSTS AND SCHEDULE ESTIMATES ...... 7-1 7.1 Cost Methodology ...... 7-1 7.2 Ground rules /Assumptions ...... 7-2 7.3 Costs for the Low Thrust OTV ...... 7-8 7.4 Program Schedule/ Funding Requirements ...... 7-9

8 REFERENCES ...... 8-1

^R TABLE OF CONTENTS (Contd)

Appendix Page

1 DEFINITIONS OF ALGORITHMS — SBR-A & Geoplatform Analyses . . Al-1 2 DE FINITIONS OF ALGORITHMS — SBR- R Analysis ...... A2-1 3 OPTOTV COMPUTER SYMBOLS DEFINITIONS...... A3-1 4 OPTOTV PROGRAM LISTING (BASIC) ...... A4-1 5 OPTOTV COMPUTER SYMBOLS AND LINE NUMBER CROSS REFERENCES ...... 45-1 6 SUMMARY OF OPTOTV PROGRAM INPUT-OUTPUT PARAMETER CATEGORIES ...... A6-1 7 OPTOTV COMPUTER PROGRAM FLOW DIAGRAM ...... A7-1 8 SOLAR ELECTRIC PROPULSION SYSTEM ...... A8-1 9 SPACE BASED RADAR-A ANALYSIS ...... A9-1 10 SPACE BASED RADAR-R ANALYSIS ...... A10-1 11 GEOPI.ATFORM ANALYSIS ...... A11-1 12 DISTRIBUTED THRITST ANALYSIS ...... Al2-1 13 COSTS ...... A13-1

vii r. i 3

LIST OF FIGURES

Figure Page

2-1 Payload allocation ...... 2-2 2-2 Missions/payloads ...... 2-3 2-3 Space based phased-array radar concept...... 2-4 2-4 Tetrahedral truss arm space-based radar ...... 2-5 2-5 SBR and OTV in orbiter ...... 2-6 2-6 Tetrahedral truss arm deployment sequence ...... 2-7 2-7 GDC tetrahedral truss demonstration...... 2-8 2-8 Tetrahedral truss space based radar ...... 2-9 2-9 Tetrahedral ring deployment sequence ...... 2-9 2-10 Geostationary platform ...... 2-10 2-11 SEPS array ...... 2-11 2-12 SEPS power array deployment . . . o • • • • • • • 2-12 2-13 Advanced 100 kW power array . . . o ...... 2-12 3-1 Orbit transfer vehicles propulsion systems• ...... •3-1 3-2 Candidate OTV concepts • • • • • . • • 3-2 3-3 OTV options • . . . . • . • . . . .3-2 3-4 OTV length and mass fraction ...... • .3-3 3-5 Thrust transient interaction • ...... •3-3 3-6 Minimum dynamic response * ...... 3-4 3-7 Distributed thrust - 2 thrusters <_ 0.2 D apart ...... 3-5 3-8 LSS dynamic response to thrust vector control system ...... 3-7 3-9 Exhaust plume interaction ...... 3-8 3-10 Engine options ...... 3-8 3-11 Low thrust engine performance ...... 3-9 4-1 Performance analysis ...... 4-1 4-2 Computer program overview...... 4-2 4-3 Payload-OTV-Shuttle length fit ...... 4-4 4-4 Low thrust AV requirements - LEO to GEO...... 4-5 4-5 Isp vs. thrust (low thrust engine performance)...... 4-6 4-6 OTV length vs. propellant weight ...... 4-6 4-7 OTV mass fraction vs. propellant weight ...... 4-7 4-8 SBR-A stowed geometries of bay and trusses ...... 4-9 4-9 Stowed geometry of SBR-R ...... 4-9 4-10 SBR-A engine thrust, propellant weight, velocity increment, and specific impulse vs. TW...... 4-15 4-11 SBR-A OTV loaded weight and payload capability and weight vs. TW .4-15 4-12 SBR-A. 1 payload, 1 Shuttle. WS = 60, 000 lb, diameter vs. TW . 04-16 4-13 Effect of engine thrust & number of burns on size of SBR-A . . . .4-17 4-14 SBR-A. Mission dependent losses, burn time, and mission time vs. TW ...... 4-18 viii

A A IF-

LIST OF FIGURES (Contd)

Figure Page

4-15 Effect of dynamic factor (KO) on size of SBR-A ...... 4-18 4-16 Effect of tip weights on size of SBR-A ...... 4-!9 4-17 Effect of lens density (WL) on size of SBR-A ...... 4-19 4-18 Effect of truss material on size of SBR-A ...... 4-20 4-39 SBR-A. (I,C) length of stowed payload and (LP) length of OTV vs. TW, compared to cargo bay upper length limits ...... 4-20 4-20 SBR-A. (PC) critical longeron buckling load and (P) induced longeron load vs. TW ...... 4-21 4-21 SBR-A. (T1) primary strut wall thickness and (D1) primary strut tube diameter vs. TW ...... 4-21 4-22 SBR-A. (IS) specific impulse held constant at 450 sec and N = 9, 5, 2. Diameter vs. TLV. . 4-22 4-23 Effect of Shuttle capability (WS) on size of SBR-A . . .4-22 4-24 Effect of constant acceleration (variable thrust) on size of SBR-A .4-23 4-25 Effect of reduced engine performance on size of SBR-A . . . . .4-23 4-26 Effect of number of Shuttles or size of SBR-A...... 4-24 4-27 SBR-R. Engine thrust, propellant weight, velocity increment, and specific impulse vs. TW . • ...... 4-25 4-28 SRB-R. OTV loaded weight and payload capability and weight vs. TW ...... 04-25 4-29 Effect of engine thrust and number of burns on size of SBR-R . . .4-26 4-30 Effect of lens density (WL) on size of SBR-R ...... 4-27 4-31 SBR-R. (WN) power spider weight = 10 and 15 lb. Diameter vs. TW . 4-27 4-32 SBR-R. (TL) lens thickness 0.086 and 0. 125 inch. Diameter vs. TW. 4-28 4-33 SBR-R. (ZH) hub weight fraction 0.47 and 0.65 inch. Diameter vs. TW ...... 4-28 4-34 SBR-R. (A) truss face width = 150, 200, 300, and 400 inches. Diameter vs. TW ...... 4-29 4-35 Effect of truss material on size of SBR-R , . . . . , .4-29 4-36 SBR-R. K0 = 2 and 1...... , , 4-30 4-37 Geoplatform. Engine thrust, propellant weight, velc city increment, and specific impulse vs. TR: ...... 4-31 4-38 Geoplatform. OTV loaded weight and payload capability and weight vs. TW ...... 4-31 4-39 Effect of engine thrust and number of burns on size of geoplatform . .4-32 4-40 Effect of tip weights (WT) on size of geoplatform . . , . . , , , 4-33 4-41 Geoplatform. (ZH) hub weight fraction = 0.47 and 0.65 inch, WT = 1200 lb. Diameter vs. TW ...... 4-33 4-42 Effect of truss material on size of geoplatform ...... 4-34 4-43 Geoplatform. K0 = 2 and 1 ...... 4-35

ix LIST OF FIGURES (Contd)

Figure Page

5-1 Baseline design definition. .5-1 5-2 Baseline low thrust OTV ...... 5-? 5-3 Low thrust OTV (model) . . . . .5-3 5-4 Low thrust OTV dimensions . . . . .5-4 5-5 Torus LO2 tank design...... 5-7 5-6 Torus tank considerations - I. Geometry .5-8 5-7 Torus tank considerations - II. Membrane behavior. .5-8 5-8 Torus tank considerations -III. Membrane support options .5-9 5-j Torus tank considerations - W. Membrane forming and joining . . .5-9 5-10 Torus tank considerations - V. Structural concept options . . . .5-10 5-11 LH2 trapped vapor head vs. acceleration for various screens. . . .5-15 5-12 LO2 trapped vapor head vs. acceleration for various screens. . . .5-15 5-13 LO2 tank propellant acquisition device ...... 5-16 5-14 LO2 acquisition with thrust misalignment ...... 5-16 5-15 Torus acquisition device design ...... 5-17 5-16 Screen surface retention pressure in 13 psi LO2 ...... 5-20 5-17 LO2 screen flow pressure loss, m - 1.863 lb/sec ...... 5-21 5-18 LO2 tank acQiisition system ring manifold pressure loss, th = 1.863 lb/sec ...... 5-22 5-19 Branch cha.ni.el configuration ...... 5-23 5-20 Total LO2 acquisition device system pressure drop for flow into one branch channel, m = 1.863 lb/sec ...... 5-23 5-21 LH2 tank acquisition device ...... 5-25 5-22 LH2 acquisition with thrust misalignment ...... 5-27 5-23 Screen surface retention pressure in 14 psi LH2 ...... 5-27 5-24 Total LH2 acquisition device system pressure drop for flow into one branch channel, m = 0.311 ?b/sec ...... 5-28 5-25 Mission profile selected for subsystems definition ...... 5-29 5-26 OTV tanks average boiloff flux during ascent ...... 5-31 5-27 Space boiloff rates for insulated LH2 and LO2 tanks ...... 5-31 5-28 MLI insulation system weight ...... 5-32 5-29 Vent option vs. no-vent option during LTPS/LSS checkout period . .5-33 5-30 40-hr checkout (payload penalty) ...... 05-33 5-31 MLI f.nfluence upon L02 tank vapor residuals ...... 5-34 5-32 Payload optimization of LO2 tank MLI system ...... 5-34 5-33 Torus insulation design ...... 5-35 5-34 Torus insulation blanker (8) ...... 5-36 5-35 Torus insulation purge...... 3-37 5-36 Purge system enclosure ...... 5-38 5-37 MLI influence upon LH2 tank vapor residuals ...... 5-40 5-38 Payload optimization of LH2 tank MLI system ...... 5-40

x LIST OF FIGURES (Contd)

Figure Page

5-39 Propellant tank pressurization system ...... 5-41 5-40 Zero-g thermodynamic vent system ...... 5-44 5-41 Propellant tank pressure histories for eight-burn OTV mission . . .5-45 5-42 Helium supply system for abort dump ...... 5-48 5-43 OTV mission parameters influence upon L02 tank ...... 5-49 5-44 OTV mission parameters influence upon LH2 tank ...... 5-50 5-45 Data buses and advanced redundancy technigugs ...... 5-54 5-46 Enhanced reliability for OTV ...... 5-55 5-47 Power/mass efficiency operating regimes for var,'.ous power sources .5-56 5-48 Low thrust OTV interfaces ...... 5-60 5-49 Deployment ...... 5-61 5-50 Shuttle c. g. • ...... 5-62 6-1 Torus tank tests (Ref. AIAA 70-1325) ...... 6-2 6-2 Aluminum torus tank ...... 6-3 7-1 Summary work breakdown structure ...... 7-2 7-2 DDT&E WBS ...... 7-3 7-3 First unit WPS ...... 7-4 7-4 Production phase WBS ...... 7-5 7-5 Operations phase WBS ...... 7-6 7-6 Low thrust OTV program schedule ...... 7-9 7-7 Low thrust OTV annual funding requirements . _ ...... 7-10

xi LIST OF TABLES

Table Page

2-1 Potential missions/payloads for low-thrust propulsion • • • • • . 2-1 2-2 Design and operational characteristics of selected payloads • • • • 2-3 3-1 Low thrust engine technology • • • • • • • • • • • • • . 3-9 3-2 Fixed thrust vs. throttling • • • • . . • . • • • • • . .3-10 3-3 Mission time (0.05 g max) • . . . . . • . . . • . • . •3•-10 4-1 Mission dependent losses ...... 4-7 4-2 Weight summaries for baseline configurations at optimum thrust-to-weight ...... 4-36 5-1 Weight summary low thrust OTV . . . . .5-4 5-2 Detailed dry weight breakdown ...... 5-5 5-3 Detailed weight breakdown of propellants and fluids . . . . .5-6 5-4 Screen types . . . . . , ...... 5-19 5-5 Assumptions and requirements for insulation optimization study. . .5-30 5-6 Helium pressurant mass quantities for low thrust OTV mission . . . 5-43 5-7 OTV abort dump pressurization system requirements . . . 5-46 5-8 OTV mission transfer orbit parameters . , . . . . .5-50 5-9 Low thrust OTV mission profile...... 5-52 5-10 Low thrust OTV avionics weight and power summary ...... 5-57 7-1 Low thrust OTV development and theoretical first unit costs . . . .7-7 7-2 Production and operation costs ...... , ...... 7-8 7-3 Low thrust OTV funding requirements . , . , . . . . , . .7-10

)di SUMMARY

Large Space Systems (LSS) such as the Geostationary Communications Platform (GP) and Space Based Radar (SBR) are planned for the late 1980's and the 1990 1 x. These are "next generation" spacecraft as large as 600 feet in size and up to 25, 000 pounds in weight. Forty-seven such missions have been forecast (1987-2000) in the OTV Concept Definition Study mission model, reference 1.

It will be advantageous to deploy and check out these expensive spacecraft in Low Earth Orbit (LEO) while still attached to the Orbiter, so any problems can be fixed, ever, by EVA, if necessary. The Space Shuttle will offer this opportunity. Once deployed and functioning, low acceleration during transfer to higher orbits (GEO) would minimize stresses on the structure, allowing larger -.ze and/or lower :.eight spacecraft.

This report documerts results of a "Low Thrust Vehicle Concept Study" conducted over a 9-month period, September 1979-May 1980, to investi6*1te and define new low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low-acceleration orbit transfer of large space systems. This study for NASA/MSFC was conducted in coordination with low-thrust engine/propulsion studies/technology efforts at NASA/LeRC and NASA/MSFC. The results of this systems/concept study are intended to help guide the propulsion technology effort already underway and to provide additional data to compare new, low-thrust chemical propulsion systems with other propulsion approaches.

Study results indicate that it is cost-effective and least risk to combine the OTV and stowed spacecraft in a single 65K Shuttle. Inspection of the mission model shows that there are 25 such missions, starting in 1987. Multiple Shuttles (LSS in one, OTV in another) result in only a 20% increase in LSS (SBR) diameter over single Shuttle launches.

Synthesis and optimization of the LSS characteristics and OTV capability resulted in determination of the optimum thrust-to-weight ratio (T/W) and thrust level.

For the Space Based Radar with radial truss arms (center thrust application), the optimum T/W (maximum) is 0. 1, giving a thrust of 2, 000 1b.

For the annular truss (edge-on thrust application) the structure is not as sensitive.

For the geoplatform, optimum T/': is 0. 15 (3, 000 lb thrust).

The effects of LSS structure material, weight distribution, and unit area density were also evaluated. In general, results in the 1-3 K thrust range were relatively insensitive. xiii A constant-thrust, 9-burn trajectory gives better performance (and is less sensitive than constant acceleration-variable thrust, 2 burn) and eliminates increased engine complexity (multiple low-thrust levels). The overall impact of these 2 modes of operation has yet to be evaluated; however, analysis of OTV ii. sulation and pressurization requirements has determined that propellant tank vapor residualsiprassures are little affected by the number of burns.or engine thrust level. Increased mission duration (80 versus 60 hours total time including checkout, deployment, transfer) makes little difference.

Engine thrust transient results in a dynamic factor of approximately 2. T`,.is can be reduced by using a slow, or a stepped thrust transient, but either complicates the engine, and results in little improvement (3%) in the LSS size.

Distributed thrust, in addition to complicating the design of the OTV and LSS, increases dynamic loading on the structure due to the difficulty in exact phasing of multiple thrusters.

To maximize the Orbiter payload bay volume available for the large space structure, a torus L02 tank is used to achieve minimum OTV length. For the 65K Shuttle, the OTV is —18 feet long, having a propellant loading of 38, 000 lb and a burnout weight of 6,000 lb. Considering the Orbiter support equipment and rotation, over 35 feet is allowed for payload length. The c. g. of the OTV, payload, and support equipment fall within the Shuttle constraints. All propellanis are dumped overboard in the event of an abort.

The technology of torus tanks was investigated and a unique acquisition device was conceived that minimizes residuals no matter what the thrust offset. Only one propellant outlet is required, and no separate sumps are needed. Thrust transient analysis indicated that no negative accelerations occur on the propellant during engine operations.

Several types of engines were considered; a new low-fixed thrust pump-fed engine and a low-thrust (pumped idle) mode of the OTV engine. Using 1500-1b thrust at 455 sec Isp and a 9-burn trajectory, a payload mass of —16, 000 lb can be delivered to GEO.

CONCLUSIONS

Based on the groundrules of this study, an optimized low-thrust OTV configured specifically for orbit transfer of large space systems has been defined. The following conclusions are made:

xiv • Engine for an optimized low thrust stage — Very low thrust (< 1 K) not required. — 1-3 K thrust range appears optimum. — Thrust transient not a concern. — Throttling probably not worthwhile. — Multiple thrusters complicate OTV/LSS design and aggravate LSS loads.

• Optimum vehicle for low acceleration missions — Single Shuttle launch (LSS and expendable OTV) most cost-effective and least risk. — Multiple Shuttles incr-_aase LSS (SBR) diameter 20%. — Short OTV (which maximizes space available for packaged LSS) favors use of torus tank. — Propellant tank pressures/vapor residuals little affected by engine thrust level or number of burns.

RECOMMENDATIONS

• Further study — Revise results as new mission and spacecraft data become available (especially as the Geoplatform design evolves). Reevaluate study results as LeRC low-thrust engine studies produce design concepts and cost data. Coordinate with OTV study to compare total mission model requirements and costs of a dedicated low thrust system vs. a conventional OTV operating in the pumped-idle mode. Purther evaluate benefits of deploying LSS at LEO vs. GE O. -- Evaluate how Centaur (with idle mode) could satisfy initial requirements.

• Technology development Hardware R&D is necessary for the engines and vehicle subsystems (low- thrust engine options, torus tank, acquisition, insulation) in order to utilize the dedicated low-thrust system.

i xv 1 INTRODUCTION

Many of the large space systems that have been identified as candidates for transportation to geosynchronous orbit are deployed at low earth orbit and are very sensitive to and have minimum capability to withstand acceleration loads during transfer to higher orbits. Because the real requirements and constraints concerning accelerations are currently little known, the effects on propulsion of a wide range of accelerations were investigated. This study investigated and defined low-thrust chemical propulsion systems configured specifically for low-acceleration orbit transfer of large space systems.

1. 1 OBJECTIVES

The specific objectives of this study were:

a. Characterize missions which require or benefit from low-thrust orbital transfer.

b. Identify, define, evaluate, and compare candidate low-thrust liquid propulsion orbital transfer stage/vehicle concepts.

c. Investigate payload/vehicle interactions and design implications.

d. Determine propulsion/system characteristics having the greatest influence upon system suitability/capability.

e. Identify and describe propulsion technology requirements.

1.2 STUDY APPROACH

This study was conducted in six major parts over a 9-month period.

During the study, missions and payloads were analyzed to select representative large apace systems and their design and operational characteristics which are impacted by the orbit transfer system.

Candidate low-thrust propulsion system concepts and their characteristics were generated to meet the large space systems mission requirements.

The perfor-nance capability of the candidate low-thrust propulsion system concepts were analyzed over a wide range of maximum payload acceleration levels.

1-1 Interactions of low-thrust propulsion concepts and mission operations with the candidate payload configurations were analyzed to determine and quantify mutual design implications.

The propulsion system characteristics which are the design drivees and have the greatest influence upon the system suitability/capability were identified.

A baseline OTV concept was identified and defined to the conceptual level.

Engine/propulsion and unique subsystem requirements were identified for technology/ component planning.

Based on the conceptual-level definition of the propulsion system elements, cost estimates were developed, and milestones and major events were identified for development and production.

1-2

2 MISSION /PAYLOAD DEFINITION

Missions and payloads were analyzed to select representative large space systems and their design and operational characteristics which are impacted by the orbit transfer system.

2.1 POTENTIAL MISSIONS/PAY LOADS FOR LOW-THRUST PROPULSION

Mission pLowning (NASA and DOD) tafor:nation (reference 1) was used to identify potential low thrust missions, payload characteristics, transportation needs, and schedule requirements. Table 2-1 summarizes these data.

Table 2-1. Potential missions/payloads for low-thrust propulsion.

NUMBER IOC

GEO-platform Demo - 12, 500 lb x 25 ft 1 1981 GEO-platform - 15, 000 lb x 25 ft 12 1992

Space Based Radar Polar - 10, 000 lb x 25-35 ft 3 1988 GEO - 15, 000 - 25, 000 lb x 60 ft 2 1991 Nominal Model DOD Class 2 - 12, 000 lb x 20 ft 4 1990

DOD Class 3 - 25, 000 lb x 25 ft 8 1992 j

Pers Comm - 54, 000 lb (3 parts) each - 12 1993 18, 000 lb x 60 ft -17 -ray telescope/Gravity wave 1997 interferometer (space fab) 'Alwx'Model Solar Power Demo (space fab) 1995

2-1

2.2 OTV REQUIREMENTS

From these data, the range of requirements imposed on the OTV was determined. Figure 2-1 shows that for payload IOC's in the first 5 years of LSS operations (1987- 1992), single Shuttle launches are sufficient.

Starting in 1991, longer and heavier payloads require multiple Shuttle operations

25 • CL3 ('92) • SBR (191)GEO

MULTIPLE SHUTTLES LEO-GEO IN ONE, OTV MOTHER) PAYLOAD 20 (22 MISSIONS) 0000 LB) FL SINGLE SHUTTLE 165K1 • PE RS COM 1'931 (PL + OTV) 126 MISSIONS)

15 ------0GP('92)----'^ • S6 ('91) GEO 1

CL2 ('901 • 0 GP DEM ('871 POLAR POLAR 1 SBR ('881 SBR ('881 1 10 0 10 20 30 40 50 60 PAYLOAD LENGTH (FT)

* 54.OW L8 (3 PARTS) Figure 2-1. Payload allocation.

2.3 SELEC"_ :,) PAYLOAD CHARACTERISTICS

The Geoplatform Communication Antenna System and the Space-Based Radar Antennas (shown in Figure 2-2; are the leading near term missions. These were selected for analysis. Table 2-2 shows that the mission drivers are: 1987 IOC, 35 ft payload, 15.000 lb payload, and geosynchronous mission.

A solar power array was initially considered but was determined to be an unlikely candidate for low thrust chemical propulsion. (See Section 2.3.3.)

2-2

SPACE BASED RADAR GEOPLATFORM

0 N'l ^a

Figure 2-2. Missions /payloads.

Table 2-2. Design and operational characteristics of selected payloads.

SBR G P POLAR GEO EXPER OPR DESIGN CHARACTERISTICS WEIGHT (LB) 10,000 15,000- 12,500 15,000 (NOM) 25,000 STOWED LENGTH (FT) 25-35 60 25 25

OPERATIONAL CHARACTERISTICS

MISSION 5600 N. MI. GEO GEO GEO PO LA R IOC 1988 1991 19 B7 1992

AIRCRAFT ADVANCED .ADVANCED SHIP, GROUND COMMUNICATION COMMUNICATION FUNCTION SAME VEHICLE SKIN AND EARTH AND EARTH TRACKING OBSERVATION OBSERVATIONS

LIFE 10 YR 10 YR 5 YR 16 YR (NOM) SERVICING NO NO PEST EVERY 1-1/2 YR CD IMPACT 'D BY OTV REF: NASA/MSFC 29 FE13 1980 SELECTED MISSIONS ARE THE GEOPL.ATFORM .AN[) SPACE 13ASED RADAR. DRIVING REQUIREMENTS ARE:1987 IOC; 25-35 FT PAYLOAD LENGTH. 15,000 LB PAYLOAD WEIGHT TO GEOSYNCHRONOUS ORBIT.

0111, VIAL PAGE IS c)—.q OF POOR QUALITY — 2.3.1 SPACE BASED RADAR. The phased-array Space Based Radar concept is shown in Figure 2-3. Two LSS versions being considered by GDC for the SBR are the tetrahedral truss (arm) and the tetrahedral truss (ring).

Feed

Feedj I ^• RF feed radiates phase command mast \ & radar signal to top of lens Radiation from feed Solar array Lens • Lens-mounted electronics phase shifts, amplifies & transmits radar signal

Phased array output radiation ^. Multiple target detection/tracking via narrow (- 0.08 °) electronically To target scanned (t 60 °) beam

Figure 2-3. Space based phased-array radar concept.

a. Mission goals 1. Preclude need for expensive upkeep of dew line and AWACS flights. 2. Provide earlier advance warning.

b. Background 1. Ten years of feasibility studies of ocean surveillance sensors. 2. "On-orbit assembly" studies for SAMSO in 1978 (reference 2). 3. DARPA technology underway, including new GDC lens study. 4. Recent NASA/MSFC RFP for flight experiment of large deployable antenna.

c. Concepts 1. Polar Orbit (a) Approximately 200 ft diameter gives good resolution (b) 6 to 12 spacecraft give coverage

2-4

W (c) IOC could be as early as 1988 (d) Each spacecraft weighs 10, 000 pounds and requires about 25-35 ft stowed length.

2. GEO orbit (a) 300 to 600 ft diameter needed for resolution. (b) 1 or 2 spacecraft required (c) IOC probably would follow polar-orbit concept (d) Each spacecraft weighs 15,000-25, 000 pounds and requires about 60 ft stowed length

2.3.1.1 Tetrahedral Truss (Arm). The tetrahedral truss (arm) concept is shown in Figure 2-4. Packaging of the SBR and OTV in the orbiter is shown in Figure 2-5.

EIGHTS, LB 3TAL PAYLOAD 15,930 rRUCTURE 2,069 ODES (JOINTS) 4,590 ENS (ANTENNA) 4,245 US (CORE) 5,021 IAMETER, FEET 420

ATV

HUB - `1 SOLAR POWER ARRAYS `MAST LENS ARRAY i

PHIS CONCEPT WAS DEFINED IN DETAIL BY GDC DURING ODA STUDIES FOR THE AIR FORCE

Figure 2-4. Tetrahedral truss arm space-based radar.

2-5

R ^r_J N'^] .,SBR_.,

Figure 2-5. SBR and OTV in orbiter.

Six radial expandable truss structures support an active lens array having hexagonal flat pattern. The hub to which these truss structures is attached also mounts an antenna-feed support boom. Packages such as solar arrays, attitude controllers, expendables, and receiver/transmitter equipment are located at either the truss ends or at the hub. The OTV is also attached to the hub and its propulsion thrust vector is normal to the plane of the lens array.

Deployment of the tetrahedral truss (arm) concept is shown in Figure 2-6. The OTV is located at the center of the six-arm truss and thrust is applied normal to the truss and array.

2-6

0 ^N

V 14^

Figure 2-6. Tetrahedral truss arm deployment sequence.

The tetrahedral truss under development by GDC for LSS applications ( Figure 2-7) has been selected as the basic structural element in the space- based radar. A proto- type deployable truss has been manufactured of graphite/epoxy (GY70/X-30) tubular members with aluminum end fittings, hinges, and hubs.

A principal feature of the beam lies in its ability to accomplish controlled, bay- by-bay deployment.

2-7 A.

.r wM'M NZ

A B , .. •^. fir/ ^wr..

. s h6. C lrdf-N^ ..

Fi hiure 2-7. GDC tetrahedral truss demonstration.

2.3. 1. -° g).Tetrahedral TrussTruss (Rin This concept (annular phased arra y antenna), shown in Figure 2-ti, has planar (edge) thrust application. The design flexibility of the basic PFTA (parabolic expandable tetrahedral antenna) has been exploited to form a ring structure.

The annular expandable truss supports a lens array. The hub, in this case, is located on the array periphery, and the OTZ' thrust vector acts on the hub along; a radial in the plane of the array. The feed support boom is assumed to be retracted during; LEO-to-GFO transfer and, as such, adds only to the hub weight.

The deployment sequence is shown in Figure 2-9.

^- d TRUSS

WEIGHTS. A oTV TOTAL PAYLOAD 15,204 ^^\ICI/,, STRUCTURE 1,082 NODES (JOINTS 2AW LENS (ANTENNA) 6,291 HUB (CORE) 4,861 MAST -/rte\ \ DIAMETER. FEET 216 98

BAY

CTHE TETRAHEDRAL TRUSS IS ALSO USED TO FORM ANNULAR RING FRAME Figure 2-8. Tetrahedral truss space based radar.

• IU-310VE M(IIMI.E O S'1'()WF:1) CUNFIGIIIWrION • itOTATE tli PEit STAGE

1+ -i*

IOI ATTACH MU1IULF, '1'0 lll'111•:11 S'1'AG4:

01 • Itl•:AIuV F: MuUU IJ.' S'1'IkJN(:IIACh • S'INAV IN ()ItII1TLII ^`^) IIGI'LI^1' S'1'tUll"1'1111 F: Figure 2-9. Tetrahedral ring deployment sequence. 2-9 2.3.2 GEOPLATFORMs One version of the Geoplatform Communication Antenna concept being considered is shown in Figure 2-10.

WEIGHTS, L8 / TOTAL PAYLOAD 16,798 STRUCTURE 816 NODES (JOINTS) 1,811 HUB (CORE) 6,972 TRUSS END TOTAL 7,200 DIAMETER, FEET 220

;1 { P't

THE GEOPLATFORM IS AN EVOLVING CONCEPT. THIS EARLY DESIGNOPTION ALSO USES THE TETRAHEDRAL TRUSS.

Figure 2-10. Geostationary platform.

The Geoplatform truss arrangement is similar to that of the SBR tetrahedral truss (arm) but it does not include a lens array and associated feed. Individual add-on packages, such as solar arrays, antennas, attitude controllers, expendables, and receiver/transmitter equipment are, as in the case of the SBR, simulated by equivalent masses located at either the truss ends or at the hub. The OTV thrust vector is coaxial with the hub and normal to the plane of the radial trusses.

a. Mission Goals 1. Maximize efficient use of available frequency spet-trurn through frequency reuse and other iced technologies. 2. Reduce congestion Ati the geosynchronous orbital arc. 3. Reduce costs by subsystem sharing and "economy of scale". 4. Use primarily for communications (commercial, :NASA, and DOD) but also offer tenancy and support for experiments, etc.

b. Background. NASA/NISFC Phase A conceptual definition continuing by GDC with COMSAT, coordinated with commercial interests.

-10 c. Concepts 1. Range from very large, docked modules to a group of platforms "flying in formation". 2. Range in weight from 12, 500 to 37, 000 pounds requiring 25 to 60 feet stowed length. 3. Early experimental platform planned for 1987; operational units by 1992.

2.3.3 POWER ARRAY. Since the SBR is moderately flexible and the GP is relatively stiff, we initially considered solar arrays to evaluate very flexible structures. How- ever, investigation revealed that the current SEPS arrays ( Figures 2-11 and 2-12) are designed to be retracted on orbit in cast of solar flares and, therefore, are not really required to be transferred in a deployed condition. (For both the SBR and GP, the solar arrays are not deployed until geostationary orbit is achieved.) Advanced (hardened) solar arrays could be designed to be transferred deployed since they may not have to be retracted on orbit. (Rigid array concepts have been evaluated under Contract NAS8-33442 and some structural concepts are similar to the SBR, such that similar results can be exrreeted.) However, in the future, advanced solar arrays (Figure 2-13) will likeb be self-powered (Ion or MPD engines) and, therefore, are not likely candidates for chemical propulsion transfer. As a result, no further consideration was given to solar arrays.

Figure 2-11. SEPS array. I^

2-11 ORMi.N; ',1. PAGE' tS UP' P"w ill QUAITI'l- FULL EXTENSION INS*

STOWED/CAGED J

READY TO EXTEND EXTENDING

Figure 2-12. SEPS power array deployment.

496.2 FT

246.0 FT

J L L (_ 1 lN( L^ (M 27.3 FT t1l 1Ja I7Z— T a l__ 1.0 FT 167.6 IN. TYP

Figure 2-13. Advanced 100 kW power array.

2.3.4 EXTERNAL FORCES AND TORQUES AT LEO_. The drag force and torques due to air, solar pressure, and gravity gradient at 200 n. mi. were determined. Using the SBR (arm) concept as an example, the drag (0.72 lb) and torque (13.2 ft-1b) are very small compared to the main engine thrust (>1000 lb), and the external forces vanish during the first burn, even on a 9-burn trajectory.

2-12 5 ," 3 CANDIDATE LOW-THRUST PROPULSION SYSTEM CONCEPTS

Analyses were conducted for expendable vs. reusable, single stage vs. 2-stage, single vs. multiple Shuttle launches, and 65K vs. 100K shuttles. (See Figures 3-1, 3-2, and 3-3.) The most cost-effective option is the single Shuttle-expendable OTV.

As reported in Section 2, the first 5 yewrs of LSS operations do not require long (60 ft) payloads.

The single (65K) shuttle-expendable OTV option was, therefore, selected for primary study.

3.1 OTV CHARACTERISTICS

To obtain the shortest possible stage to allow maximum payload length, a torus L02 tank configuration is superior to all others (conventional suspended tanks, nested tanks). For 40, 000 lb propellant at MR = 6, a savings of 9 feet in length is realized.

Although the torus tank itself is heavier, the shorter structural shell and support systems com- Figure 3-1. Orbit transfer pensate, resulting in nearly equal weight. vehicle. The OTV Concept Definition Study (Contract NAS8-33533) has generated detail data on many configurations, permitting determination of OTV parametric relationships as shown in Figure 3-4.

Note that these represent basic weights including all vehicle subsystems. The effects of mission duration, engine thrust, number of engine burns, etc., are separately evaluated.

3.2 PROPULSION CHARACTERISTICS

3.2.1 THRUST TRANSIENT. Engine start and cutoff thrust transient (Figure 3-5) induces oscillation of the LSS, which causes changes in acceleration of the OTV and the propellant mass, both of which are of interest when considering LSS structural loading and OTV propellant acquisition.

3-1

GEU j SINGLE STAGE OTV 140.8860IW M.F. J } L02IL112 PAYLOAD 1 #t EXPENDABLE (REUSABLE NO PL RETURN) 14000 AV UP OR DOWN

NGLE S►IUTTL A !C IU A L^^ IoTV 6 PL ON OTHER FOR OTV "K I* 66K SHUTTLES---1 ^— IIIUKSHUTTLES 511U1 TlE SHUT

tldk PL 34' # IYKI # IL

' t4IK PL #tZ"I"L IIUKI Ww U1V # llbK) 6UK UTV I ___ — I

6JK (4 1 a' al, I

l 1

Figure 3 Candidate OTV concepts.

Jfiu IsP bU \y UU MF 14000 ^1`5 kACH WAY (All PHOPI)tSIVL1 Jb Jyb cost (SM) J1 SHUTILL \v 1 lb LXPOTV JU - )IRAI \5 J RkUSL U1 V

l^ Jb lKX1U F^sF .

\ Ste`` J11 61kX1 -

PAYLVAU I IUIIU l81 1b xY` NL1 \ SIIU`

Ib Slltll1^ L^ I yll; Ll^'^.^ 11 5^lulil t Ill Ill wit

a 6b 10t) bb 100 SHUI 11 L CAPA0M I Y IIL" t U) SIM 11 L CAPAUII 11 \ JIM) L U)

SINGLL SHUTTLE/SINGLE STAGE EXPENDABLE 01V ISMUSI COST - EFFECIIVL AND LEAST HISK. HEOUMLS SIIO14 1 O TV 10HUS I AN K. 1 Figure 3-3. OT`' options. 3-2 r

LENGTH. FT 60 0 MR^6 50 0.90 CONVENTIONAL 40 MASS FRACTION O CENTAUR NESTED 1761N. DIA 0.80 L02/LH2 O TORUS 20 MRx6 m CONVENTIONAL TOROID

0L 0.700 0 50 100 150 50 100 PROPELLANT WEIGHT, 1000 LB PROPELLANT WEIGHT, 1000 LB

TORUS TANK GIVES SHORTEST OTV, WITH LITTLE (<3%) WEIGHT PENALTY — SELECTED FOR BASELINE.] Figure 3-4. OTV length and mass fraction.

STRUCTURE ACCELERATION

• ...,. ^^^^. ^ Ayy CLGf^NI IVIY

ENGINE THRUST

t OBJECTIVES • STRUCTURE INTEGRITY • PROPELLANT ACQUISITION • THRUST VECTOR CONTROL c.p.

(GIMBAL POINT t Figure 3-5. Thrust transient interaction. 3-3

Figure 3-6 shows results of an investigation of the dynamic response of the OTV (coupled with the payload) as a function of thrust rise time. It can be seen that no negative acceleration occurs.

SYSIEM; MPL UNIFONMLYDISTNIBUIE0 PAY LOAD ACCEL. OF MWMa',2^ / 1 1 Me (MIN) D.6 ^' +— —1 -- i ACCEL. OF C.O — i MOTV I r . MPHOP MWMa ^ ^ / MOt1El. D4 76L— -- 75L - i M. MWMe - 4 Mb Mb MASSLESS BEAM HMI 42 ` X" (MA X Mb/Ma i bl M. - U b MPl # 1iAUTV * 1i1PROP MIJMa - . - Mb - U 2b MPL 0 ^ 1 1 1 0 0.2 0.4 0.8 0.8 1.0 1.2 1.4 1.5 THRUST RISE TIME/ NATURAL PERIOD OF SYSTEM Figure 3-6. Aiinimum dynamic response.

Similarly, the maximum d ynamic acceleration was determined. In general, a dynamic factor of ^2 resulted.

Consideration has been given to stretching out the engine thrust transient to result in factors nearer to 1. Since engine thrust rise times are typically less than 1/2 second and the structures' natural periods are 2 seconds and greater, this would require engine design changes. However, as discussed in Section 4, the effect on the structure is negligible whether the factor is 1 or 2 and, therefore, there is no need to slow down the thrust transient.

3.2.2 DISTRIBUTED THRUST. Application of distributed thrust on the payload structure should ideally result in better load distribution mid. therefore, lower weight, and could offer the possibility of a common main propulsion and attitude control system. However, the imposed design complications (plumbing, etc., for a deployable system) and the dynamic characteristics incurred by expected variations in thrust transient phasing are such that this would be difficult to achieve.

3-4 Ideally, the static bending moments and the dynamic response would be reduced If the thrust rise and cutoff of each engine were phased exactly together. However, dynamic response could be increased since exact phasing is unlikely with multiple, distributed thrusters. (Even with quite closely related thrusters, e. g. , Atlas, Centaur, significant lateral dynamic loads are encountered.)

Figure 3-7 shows the results of an analysis for two thrusters separated by 20% (of the structure diameter). The figure shows that small differences in thrust transient cause significant increase in dynamic factors. Refer to Appendix 12 for details.

2.2 TIP 2.1 T- SEC --^ (THRUST TRANSIENT J

2.0

1.9 DYNAMIC .7 FACTOR 1.8

OTV 1.7 /

1.6 .3 .5

s 1.5 .7

1.4 L 0 10 .2u I, TIME LAG (SEC)

PROP - 0 STRUCTURE MAXIMUM THRUST - 1000 LB. OTV - 6000 11 - 0.5 HZ DIA - 400 FT PL - 16000 f2 - 0.75 HZ

Figure 3-7. Distributed thrust - 2 thrusters < 0.2 D apart.

3.2.3 THRUST VECTOR CONTROL. The OTV will have a thrust vector control system using vehicle rate and attitude to provide commands to gimbal the engine. The control systems also have the capability of shaping the thrust vector commands as a function of time and frequency. The following assumptions were made for this analysis.

a. An autopilot using attitude and rate without filtering is assumed.

b. The lowest two elastic modes for the payload/upper stage combination were calculated by an approximate analytic technique. The frequencies are 0.2 and 0.65 Hz.

3-5 c. The control frequency was set at 0.011 Hz.

d. The elastic modes were assumed to have zero damping, this being a worst case. (All damping is provided by the attitude control system.)

Since the elastic modal frequencies are well above the control frequency, coupling is weak and elas'{eity has little effect on the rigid body response.

The analysis showed that the rigid body and elastic modes are stable when an initial attitude error of 5° was used. The rate traces (Figure 3-8) show that although the elastic modes are excited by the large attitude step, they are damping out. Analysis, therefore, shows that current upper stage control systems are adequate.

The c. g. for the baseline OTV is always forward of the gimbal point due to the attached payload. (After payload insertion, the OTV ACS provides disposal AV.)

3.2.4 EXHAUST PLUME INTERACTION. The question of OTV engine exhaust plume impingement on the LSS was analyzed and was determined to present no problem if high expansion ratio nozzles are used. (See Figure 3-9.) For a hydrogen-oxygen engine with an e = 400, the Prandtl/Meyer turning angle for Mach = -is 82.9 degrees. Even with the engine gimbaled 10 degrees, impingement would not occur until a radial distance of 385 feet was reached, assuming a 20-ft long OTV. This exceeds any of the payloads currently being evaluated.

3.2. 5 ENGINES. Low thrust engine performance data were obtained from the LeRC low thrust engine studies, from MSFC OTV engine (pumped idle mode) studies, and from Pratt & Whitney for the RL10, IM, each of which shows a falloff in Isp at lower thrust levels. The performance for either the new low thrust engine or the advanced OTV engine is equivalent, while the RL10, 11B rims about 30 seconds less.

Comparison of low thrust engine characteristics is shown in Figures 3-10 and 3-11, and in Table 3-1 (reference sources are indicated on Figure 3-11).

The factors considered in evaluating fixed thrust vs. throttling (discussed in Section 4) are shown in Tables 3-2 and 3-3.

3-6

0 M C , e • F

_^ C O ^ G ^ < ^ C S, C

C O f^ Z U C^ < < U 8A

U .r

R+ V °^ t~ x 8C it v

o ° o 0 S-ot x o e o C -ut x (Li) (z) 3aost ((rfw lU 3aox

N 0.

Q] C O i < V '^ Y U i T {{Ai —rz = Q • GCS d x x se r g=y a16 , w a i z p G C O F Z F= — JO us it i « u —ci3 r^ Z a a^ `o c. s U U U Fx Etc '. I ri i I

e w

^ y r O - N i ' ' : 1 IT C-Ct K ^Gt x 1531 lt) 3QOK LLD (C) 3QOK (anu (t) 3aorc

3-7 -10° 100 GIMBAL M -m

IW MOM OO GIMBAL 60400 r- 1.22 THERE IS NO PLUME INTERACTION WITH PAYLOAD IF HIGH EXPANSION RATIO NOZZLES ARE USED

Figure 3-9. Exhaust plume interaction.

OTV ENGINE LOW THRUST ENGINE PUMPED IDLE MODE

4%'

NEW NEW NEW +KIT* RL10.118 THRUST, LB 1500 1500 1500 3500 Isp, SEC 455 455 470 435

*CHAMBER/NOZZLE (SMALLER THROAT, COUNTERFLOW NOZZLE)

Figure 3-10. Engine options.

3-8

L02/L"2 MR-6 480 KITTEO 0 CHAMBER/NOZZLE ^^ 1 r J/ ADVANCED I _^ OTV ENGINE I 480 PUMPEOIOLE 1 ^^ (P&W) Q /^ ® ADVANCED X DEL OTV ENGINE ISP PUMPED IDLE 1110 14 (SEC) i• 440— Q NEW LOW T14HUST J 14 (L•RCI O RL10118 O

1 MCR-79467 (MARTIN) (rot. 3) 420 O6 2 R060-123 (ROCK[TOYNC) (h1.4) RLIOA-3-3 (MR- 61 3 RR-126116 (PRATT & WH11'NKY) (rot, 6) 4 /11.12263 (PRATT & WHITNILY) (rot. 6) -20SEC 6 Of 106664A (PRATT i WNITN[Y) (IF CUT OFF TO 60"LI

0 1000 2000 3000 4000 THRUST, LBF 0 PL 0 isP - 63 LBALC

Figure 3-11. Low thrust engine performance.

Table 3-1. Low thrust engine technology.

NEW LOW THRUST PUMPED IDLE (OTV ENGINE) L TECHNOLOGY - - SMALL PUMPS, - PERFORMANCE AND CONCERNS COOLING, AND STADILITY AT 100 PERFORMANCE THRUST

3-9

.. Table 3-2. Fixed thrust vs. throttling.

FIXED LOW THRUST THROTTLING (VARIABLE T/W) (FIXED T/W)

DESIGN -- SIMPLER — MORE COMPLEX PAYLOAD (LSS SIZE) — MORE -- LESS (MORE SENSITIVE) NUMBER OF BURNS — MORE (8)* — FEWER (2-4) MISSION DURATION -- 3-1/4 DAYS* — 2-1/2 DAYS

* NO CONCERN FOR PAYLOADS, OTV, OR OPERATIONS

• MULTIPLE BURNS ARE SOA (e.g., 7 BURN CENTAUR) • INFLUENCE OF ENGINE THRUST LEVEL OR NUMBER OF BURNS HAS MINIMAL EFFECT ON OTV TANK PRESSURES OR VAPOR RESIDUALS • TDRSS WILL PROVIDE COMMUNICATION/TRACKING • ELECTRONICS BEING DESIGNED FOR 5 YR IN 5600-N. Ml. ORBIT

Table 3-3. Mission time (0.05g max).

TWO-BURN CONSTANT NINE-BURN ACCELERATION CONSTANT THRUST

LAUNCH/CHECKOUT IN LEO 40 HR 40 HR

ENGINE BURN TIME 2.5 HR 5 HR

LEO-GEO TRANSFER 8 IIR 24 HR

DISPOSAL ORBIT PLACEMENT 12 IIR 12 HR 62.:i HR 81 HR TOTAL 2-1/2 DAYS 3-1/4 DAYS

3-10

PERFORMANCE ANALYSIS

A computerized analytical model was developed to synthesize and optimize the operational and hardware parameters of three different large space structural systems and their OTVs.

Five main considerations were evaluated to determine the characteristics of greatest influence. (See Figure 4-1. ) a. The large space structure weight/size vs. static load factor. b. The large space structure load amplification factor.

C. The low thrust trajectory AV required vs. T/W and number of burns. d. The OTV performance (function of engine, thrust, number of burns). e. The packaging in the Shuttle.

s — ` 020 CURRENT I C LL RANGE !Z RANGE 1 0.10 ^20 0 0B O E006 r DYNAMIC^

~ I ^ 001 0 5 3 2 00 10 20 70 2 U 16 1B TAT IDEAL VELOCITY. V I II OW foil (THRUST RISE TIME -STRUCTURE NATURAL PERIOD) SW_ r 20 000 _FT J / OTV PERFORMANCE • PACKAGING IN SHUTTLE 1 ^ STATICS i 10 000 —^ - O LL j 9 BURN PERFORMANCE ^ 0 J O j SBURN F 001 007005010 + 02 07 05 MAX TV - N AL , 10A0 FACTOR K SHUTTLE 2 BURN O^ITV.RL•ASEI AREA INCREASING j A• A3 A2 Al SI TR UC T URA THRUSTWEIGHT RATIO. T'W = CAPABILITY. '► INCL EFFECTS OF SYNTHESIS I ^ , • Is, VS THRUST o P 0 • MISSION LOSSES VS NUMBER p ! PERFORMANCE OF BURNS AND TRANSFER TIME (OPTIMUM POINT • MAXIMUM CHILLDOWN (161b p , slams AREA ► LEAKAGE M 1 IbIH1I Ma. TMIIU•T OW[IGNT ••Tip TN ATTITUDE CONTROL O.0 b %,I - POWEV (0.5 WIII') BOILOFF (MAXI 11 O b'MI Figure 4-1. Performance analysis. 4-1 4. 1 COMPUTER SYNTHESIS/OPMIIZATION

The OPTOTV computer program (Figure 4-2) is both a synthesis and optimization program for parametric and trade studies of LSS and OTV configurations operating out of the Shuttle. The program has the following features.

It accepts LSS truss structure material properties, and minimum member size and gage For purposes of this analysis, graphite composite having an E - 40 x 106 psi and an Fcy - 37, 000 psi, and aluminum (6061-T6) having an E = 107 psi and Fcy = 35, 000 psi are used. Minimum primary tube diameter and t1dckness are 2 and 0.05 inches, respectively.

The program accounts for the Shuttle payload weight and volume constraints as well as the configuration of the OTV (i.e., mass fraction and length vs. propellant weight) and its propulsion syst:m Isp vs. thrust characteristics.

The input also includes factors for weignt of joints, the LSS hub weight, dynamic amplification factors, and number of burns.

INPUT • MATERIAL PROPERTIES • MINIMUM INITIAL T/W ° .001 • MINIMUM SIZE AND GAGE LIMITATIONS • SHUTTLE PAYLOAD AND VOLUME CONSTRAINTS • OTV PROPULSION SYSTEM AND CONFIGURATION • NUMBER OF BUPNS • FACTORS: - WEIGHT OF JOINTS - HUB WEIGHT - ARRAY WEIGHT - THRUST CUTOFF AMPLIFICATION

OUTPUT • LSS - DEPLOYED AND STOWED STAGE GEOMETRIES - STRUCTURAL AND MASS PROPERTIES

• OTV - THRUST LEVEL - MASS PROPERTIES - AV REQUIREMENTS - SIZE - PAYLOAD CAPABILITY - STAGE WEIGHTS PAYLOAD WEIGHT AND OTV PAYLOAD CAPABILITY CHECK

PAYLOAD, OTV AND SHUTTLE, WEIGHTAND VOLUME FIT CHECK

OPTIMUM PAYLOAD AND OTV PARAMETERS

Figure 4-2. Computer program oven-iew.

4-2 Through an iterative computational process the program computes stowed and deployed sizes as well as structural and mass properties. It checks critical stresses including Euler column buckling of truss member tub — and also radar-array-membrane stresses. If stresses are unacceptable, the tube diameters are first iteratively increased up to the point at which volume limitation constraints are encountered. After this, the tube wall gages are increased as necessary up to the point at which weight limitation constraints are encountered. It then computes OTV length, mass, and performance parameters. To j.,erform thhese analyses, it must compute AV impulse velocity requirements to achieve orbital transfer for the selected input number of burns and initial acceleration.

Fit ch;cxs are performed to determine, for a given T/W and structure size, if the payload and volume limitations of the Shuttle are met and if the OTV payload capability matches the actual payload weight. The structure size is then systematically increased until either volume and/or weight limitations are encountered, at which point the maximum LSS size is assumed to have been achieved. The T/W is next increased and the above process is repeated to generate data for LSS size vs. T/W. For each T/W all characterizing parameters of the LSS and OTV are computed and printed out along with a factor for the fraction of the total Shuttle cargo bay length utilized. In all cases the full payload capabilities of the Shuttle are used.

Appendices 1 through 7 define all elements of the OPTOTV computer program.

4.2 PAYLOAD CONFIGURATIONS

The selected payloads, defined In Section 2, are summarized as follows:

a. Space Based Radar - Tetrahedral Truss Arms (SBR-A). Six radial expandable truss structures support an active lens array having a hexa6onal flat pattern. The hui, to which these truss structures are attached also mounts an antenna-feed support boom. Packages such as solar arrays, attitude controllers, expendables, and receiver/transmitter equipment are located at either the truss ends or at the hub. The OTV is also • , ',Ipted to the hub and its propulsion thrust vector is normal to the plane of the lens array.

b. Space leased Radar - Tetrahedral Truss Ring (SBR-R). The annular expandable truss supports a lens array. The hub, in this case, is located on the array periphery, and the OTV thrust vector acts on the hub along a radial in the plane of the array. The feed support boom is assumed to be retracted during LEO-to-GEO transfer and, as such, adds only to the hub weight.

4-3 s c. Geoplatform. The radial truss arrangement is similar to that of the SBR-A but it does not include a lens array and associated feed. Individual add-on packages, such as solar arrays, antennas, attitude controllers, expendables, and receiver/transmitter equipment are, as in the case of the SBR-A, simulated by equivalent masses located at either the truss ends or at the hub. The OTV thrust vector is coaxial with the hub and normal to the plane of the radial trusses.

4.3 PAYLOAD-OTV-SHUTTLE LENGTH FIT

Figure 4--3 shows the configuration of the stowed payload envelope and the OTV to which this , ,-load mates. As shown in the figure, the allowable mated length is 59.1 ft for 30 1, au Bch elevation with respect to the Shuttle and 55.9 ft for 75 0 elevation. For purpose:: of this study, a mated length (i. e., usable length of cargo bay) is taken as 57 ft.

se7^ r PAYLOAD -,.,.,,. X0 PAYLOAD_ PIVOT

1 11 55.9 FT /^1 1

750

T X 1302.0 30° 1302.0 PIVOT Q. 1307.0

Z° 414.0 Z° 400.0

Figure 4-3. Payload-OTV-Shuttle length flit.

4.4 OTV CHARACTERISTICS.

The OTV is an LO2/LH2 vehicle having a toroidal LO2 tank and a conventional LH2 tank. For purposes of this study the OTV outside shell diameter is held constant at 176 inches while its length is varied for stage sizing.

4-4 n n 4.4.1 VELOCITY REQUIREMENTS VS. INITIAL THRUST-TO-WEIGHT. Figure 4-4 shows velocity requirements for LEO-to-GEO transfer vs. initial thrust-to-weight (TP) for N = 9, 5, and 2 total burns based on constant thrust engine performance and for N = 2 based on constant thrust-to-weight (TW) engine performance. The trajectory methods originally developed by GDC for the Air Force in the "On-Orbit Assembly Study" reference 2, have been used in this study. Recent work by Martin has verified our original work. The algorithm describing the first set of these curves is contained In Appendix 1 under V, V1, and V2. Similar data are used for the constant TW curve. This algorithm defines a linear piecewise approximation of the curves in Figure 4-4 where V1, V2, and T3 and T4 (per listing in Appendix 3) are, respectively, the velocity and initial thrust-to-weight ranges in the algorithm. For constant thrust engine performance the final thrust-to-weight (TW) is related to initial thrust-to-weight in the TP algorithm.

20000

19000

SEPS Isp = 2900s

18000 \ \^ \^ \~ , T-1 W \\\ \\\ \ \\ , i °ERiGEE BURN — N_ \ CONST. ACCEL. tZ 1100E

0 W C7 O UJ 16000 ^^ \ 48 CONST. THRUST WITH 1, 2,4,& 8 PERIGEE BURNS — PLUS 1 APOGEE BURN 150vu `8 ---GOC TRAJECTORY CALCULATIONS GOC REPORTS — 697-0-78-002, JAN 1978 (ref 7) 14000 — 691.0-79.001, JAN 1979 (ref 8) " ---- INTERPOLATIONS (LBRC) ------MMC CALCULATIONS (450 is) REPORT — MCR 79.653 JAN 1980 (ref 9) 13000 0.00001 0.0001 0.0005 0.001 0.005 0.01 0.02 0.04 0.1 0.2 0.4 TIW, AT BEGINNING OF BURN Figure 4-4. Low thrust AV requirements - LEO to GEO.

4- 5 thrust 4.4.2 Isp VS. THRUST. The Isp (IS) vs. (TT) used in this study is shown in Figure 4-5 and the algorithm for IS, based on curve fitting, is given in Appendix 1.

Isp (IS), SEC

0 500 1000 1500 1000 2500 THRUST (TT), LB

Figure 4-5. Isp vs. thrust (low thrust engine performance).

4.4.3 OTIV LENGTH VS. PROPELLANT WEIGHT. 30 176""IA LL 176 DD12 The OTV length (LP) vs. propellant weight (PW) ford MR = 6:1 the toroidal configuration used in this study is shown = Y0 TOROID in Figure 4-6 and the describing algorithm for LP is in Appendix 1. W 1 10 0 20 40 60 4.4.4 MASS FRACTION VS. PROPELLANT PROPELLANT WEIGHT (PW), WEIGHT. The OTV mass fraction (MU) vs. pro- 1000 LB pellant weight for the toroidal configuration used in this study is shown in Figure 4-7, and the Figure 4-6. OTV length vs. describing algorithm for MU is in Appendix 1. propellant weight.

4.4.5 MISSION LOSSES. Table 4-1 summarizes mission losses (PL) analytically defined in Appendix 1 as a function of propellant weight loss per engine start (KS), per hour due to leakage, boiloff, and attitude control (KT), and for onboard power generation (PP). The latter is a function of the weight per hour for electric power up to the first 12 hours (KP), and after 12 hours (KQ). See Appendix 1 for definitions. KS = 15 lb/start KT = 4.1lb/hr KP = 21 lb/hr KQ = 0.5 lb/hr 4-6

0.90

z0 av o^ N Q 0.80

0.70 L 0 50 100 150 PROPELLANT WEIGHT (PW), 1000 LB Figure 4-7. OTV mass fraction vs. propellant weight.

Table 4-1. Mission dependent losses.

• LOSS RATE CHILLDOWN — 15 LB (MAX) PER START (ZERO IF 'LANK HEAD IDLE) LEAKAGE — 0.1 LF3 PER HOUR BOILOFF — 1.0 LB PER HOUR ATTITUDE CONTROL. — 3.0 LB PER HOUR POWER — 21 LB PER HOUR UP TO 12 IIR (BATTERIES) 0.5 LB PER HOUR AFTER 12 IIR (FUEL. CELL)

• MISSION DURATION = (COAST TIME) + (BURN TIKE)

NUMBER OF BURNS COAST TIME (INCL l APOGEE) (Im

2 5 5 10 9 25

PROPELLANT WEIGHT (LBM) CONSTANT THRUST BURN TIME (HR) = THRUST (LBF) x 3600 ISP AV (FT/SEC) CONSTANT ACCELERATION BURN TIME (HR) = (T/W) (32.2) (3600)

4-7 n 4.4.6 SUMMARY OF OTV PARAMETERS. Appendix 6 lists under GENERAL the OTV input and output parameters handled in the OPTOTV computer program. Analy- tical definitions for these parameters are given in Appendix 1. Note that the OTV payload capability (WY) has one of two definitions depending on (menu-selected) number of Shuttles used in a mission.

The OTV performance parameters described above are an integral part of the OPTOTV computer program and, as such, are iteratively evaluated in conjunction with iterative evaluations of payload-describing parameters

4.5 PAYLOAD CHARACTERISTICS

General and peculiar payload characteristics are described in this subsection for the three different generic LSS systems chosen for purposes of this study.

4.5.1 SPACE BASED RADAR - TETRAHEDRAL TRUSS ARM (SBR-A). The SBR-A consists of six deployable truss frame structures that mount a lens array. The hub to which the trusses are attached also mounts a deployabla feed assembly and an OTV adapter structure. The OTV thrust acts normal to the plane of the deployed trusses.

Attachments on the ends of the SBR-A trusses :uch as solar arrays, attitude controllers, and electronic equipment are taken as lum:)ed weights (WT) as are all weights associated with the hub (WH). Orbital transfer is assumed to take place with the WT and WH deployable structures in stowed states. Thus, a solar array on the end of a truss and a feed support boom on the hub would be in their retracted states during orbital transfer.

The stowed geometry of the truss structures is shown in Figure 4-8. Note that each bay of a truss folds into a volume C1, C2. a (or C1, C2, A in computer symbols).

The OPTOTV input and output SBR-A parameters are listed in Appendix 6 under GENERAL, Payloads and SBR-A Peculiar. Algorithms and definitions for these parameters are given in Appendix 1. The input parameters of particular importance are ZH (hub weight fraction) and FB (joint weight factor in truss), which are products, respectively, of system and structural design requirements. Input data under GENERAL, Shuttle, such as LQ and RS (available Shuttle cargo bay length and radius), also have an important influence on the SBR-A payload sizing.

The SBR-A Peculiar, dynamics output in Appendix 6 is primarily directed at com- puting K¢P (the worst case thrust amplification factor). The dynamics routine in OPTOTV is identified in Appendix 7, the OPTOTV computer program flow diagram. This routine is only applicable to the SBR-A. Dynamics analyses are not programmed for the geoplatform and SBR-R.

A. STOWED GEOMETRY B. STOWED GEOMETRY OF OF ONE TRUSS BAY SIX TRUSSES -- END VIEW

Figure 4-8. SBR-A stowed geometries of bay and trusses.

4.5.2 SPACE BASED RADAR - TETRAHEDRAL TRUSS RING (SBR-R). The tetrahedral ring structure has a triangular cross section (rather than a diamond cross section as in the cases of the SBR-A and geoplatform). The stowed geometry of the SBR- R is shown in Figure 4-9. Unlike the SBR-A and geoplatform, the SBR-R is not mounted on the OTV while stowed for launch in the Shuttle cargo bay. This is necessary because the length of the stowed package, a, is large compared to the stowed package diameter, and the thrust vector must finally be in db ^-- 25 --^ 152 (r,) the place of the deployed ring structure. It will, therefore, be necessary to per- form an on orbit mating or positioning of the SBR-R on the OTV before SBR-R deployment takes place. The remote mani- pulator system (RMS) may be needed for this assembly operation and, because of this need, the allowable radius of the payload Rs (or RS) is smEdler than that used for the SBR-A and geoplatform. Figure 4-9. Stowed geometry of SBR-R.

4-9 Describing input and output OPTOTV parameters for the SBR-R are listed in Appendix 6 under GENERAL, Payloads and SBR-R Peculiar. Definitions and algorithms for applicable GENERAL symbols are given in Appendix 1 and for SBR-R Peculiar symbols in Appendix 2. Note that, in Appendix 2, in addition to new SBR-R peculiar symbols, some of the symbols used in the SER-A and geoplatform analyses are redefined.

4.5.3 GEOPLATFORM. The geoplatform configuration consists of six radial trusses that support antennas, solar panels, and other equipment, similar to diamond cross section construction of the SBR-A. A lens array is not used in the geoplatform and its hub weights consist only of platform equipment packages. During orbital transfer, thrust-to-weight-critical packages such as solar panels and unfurled antennas are in their stowed positions.

Because of the relatively large weights of the onboard equipment, the size of the geoplatform structures tends to be significantly smaller than that of the SBR-A or SBR-B.

Describing input and output OPTOTV parameters for the geoplatform are listed in Appendix 6 under GENERAL, Payloads and Geoplatform Peculiar. Definitions and algorithms for symbols are given in Appendix 1.

4.6 SHUTTLE CHARACTERISTICS

As indicated in Appendix 6 under GENERAL, Shuttle, the only Shuttle input parameters to the OPTOTV computer program are its cargo bay length (LQ), radius (RS), and payload capability (WS). In selecting values for these parameters, allowances are made for part of the actual Shuttle capacity being used for auxiliary equipment such as the RMS and for related crew support activities.

4.7 OPTOTV COMPUTER PROGRAM

The OPTOTV computer program is both a synthesis and optimization program for parametric and trade studies of LSS and OTV configurations operating out of the Shuttle. The program is described in Section 4.1 and is further defined in this subsection.

Appendices 1 and 2 are alphabetical lists of computer and analytical symbols as well as algorithms and definitions for these symbols. Appendix 1 covers the SBR-A and geoplatform analyses, while Appendix 2 covers new terms and re-definitions of terms in Appendix 1 used in the SBR-R analysis.

Appendix 3 is a combined summary of the literal definitions of computer symbols in Appendices 1 and 2, as well as all iteration and input-output format control symbols.

R 4-10 Appendices 4 and 5 are the OPTOTV computer program listing (written in TRS-80 Disk Basic) and cross-reference listing of lines in which different symbols are used and which are referenced by other program lines. (This reference listing is generated by a Microsoft-Apparat NEW DOS utility program.)

Appendix 6 is a summary of the OPTOTV computer program input-output parameter categories. It indicates the menu of program anaiysis options which can be selected, as well as the computer symbols for input-output parameters according to the indicated GENERAL and Geoplatform, SBR-A, and SBR-R peculiar categories.

Appendix 7 shows the OPTOTV computer program flow diagram. Computer symbols are used in this diagram.

4.7.1 MISSIONS. Program options in Appendix 6 outline the different missions which are analyzed in this study (per ZZ = 0, 1, 2, and ZO = 1 menu selections). ZZ = 0 (1 payload, 1 OTV, 1 Shuttle) indicates that one payload and OTV assembly is delivered to LEO by one Shuttle flight. ZZ = 1 indicates that the entire Shuttle cargo bay is used to deliver the payload to LEO, and ZZ = 2 indicates that separate Shuttle flights are used for the payload and OTV.

Analyses are generally performed for N = 9, 5, and 2 burns, based on constant thrust engine performance. )wever, by selecting ZO = 1, the SBR analysis is run on the basis of constant 'M ( crust-to-weight) and N = 2.

The detailed OPTOTV computer program flow diagram is shown in Appendix 7.

The OPTOTV program can be used to determine a single TNV value at which the maximum payload size is achieved. This capability is, however, not used at this time; instead, the program provides data on, and a feel for, the penalties resulting from off-optimum TW operation. Constraints not considered, such as minimum free-free mode resonant frequencies of the deployed structure, may preclude selection of the maximum payload size as the best configuration. Docking loads on the aeployed structure and launch loads on the stowed structure, which would also influence maximum payload size selection, have also not been included in this analysis, but can be in future refinements.

4.7.2 GROWTH CAPABILITY. Although OPTOTV is primarily used in this study to evaluate the thrust-to-weight dependent design requirements for a low thrust OTV, it can also be used for comparative evaluation of the selected or alternative large space structure systems. The LSS systems must be Shuttle compatible and transported per program options in or similar to those in Appendix 6. To implement such additional LSS systems analyses, it would be necessary to add payload subroutines to those routines already in the program. The SEPS power module algorithms in Appendix 8 provide an example of the level of detail needed to add a payload option to the program.

4-11 OTV performance and weight estimating relationships used in OPTOTV can be further refined to account for weights of avionics, electr.cal and fluid lines and joints, engine, tanks and tank suspensions, shell structures, adapters, and thrust vector control system.

The analysis methodology presented here for comparative OTV and LSS systems offers the following advantagcs:

a. It can be used to handle almost any set of interrelated sets of system performance and weight characteristics that can be defined by algorithms amenable to closed-form or iterative solutions.

b. All aspects of the analysis (i.e. , synthesis and optimization of both the payload and OTV for different mission options) are fully automated in one program for efficient execution.

c. Fixed and optimized OTV and payload design parar:ieters can be used in, or generated by the analyses. Specific parameters or sets of parameters can also be included or excluded from the program's optimization process.

4.8 RESULTS

Representative plots of OTV and payload performance, weight, and size characteristics vs. final thrust-to-weight, TW, are presented in the following: Figures 4-10 to 4-26 for ilie space based radar tetrahedral truss arm (SBR-A) Figures 4-27 to 4-36 for the space based radar tetrahedral truss ring (SBR-R) Figures 4-37 to 4-43 for the geoplatform

Typical OPTOTV printouts from which data were taken to generate the above plots are contained, respectively, in Appendices 9, 10, and 11.

a. Baseline Parameters. Except where parametric variations are otherwise noted on individual plots (in Figures 4-15 through 4-46, or on data printouts in Appendices 8, 9, and 10), the following baseline parameters apply for the Shuttle, OTV, and payloads. The parameters in parentheses after the baseline parameters are some of the variations considered in this study.

1. Shuttle - Cargo bay length (LQ): 57 feet - Cargo bay radius (RS): 88 inches - Payload capability (WS): 60, 000 lb (90, 000) + 5, 000 lb ASE

4-12 2. OTV - Outside shell diameter (2 RS): 176 inches - Number of burns (N): 9 (5,2) - Velocity requirements vs. initial thrust-to-weight (V vs. TP): Figure 4-3 - Minimum initial thrust-to-weight: 0.001 - Isp vs. thrust (IS vs. TT): Figure 4-4 - Length vs. propellant weight (LP vs. PW): Figure 4-5 - Mass fraction vs. propellant weight (;NIU vs. PW): Figure 4-6 - Propellant losses: Section 2.2.5 - Thrust-to-weight dynamic amplification factor (KqS): 2 (1) - Engine thrust characteristics: Constant TT (Constant TW)

3. Payloads (a) General - Structural construction material: Graphite Composite (Aluminum) - Minimum primary strut diameter (DIM): 2 inches - Minimum primary strut wall thickness (Till): 0.05 inch - Secondary strut diameter (D2): 1 inch - Secondary strut wall thickness (T2): 0.025 inch - Structural joint weight factor (FB): 3.218 - Hub weight fraction (ZH): 0.47 (0.65)

(b) Space Based Radar-A (SBR-A) Unit - area weight of lens («'L): 0.048 lb/ft2 (0.095, 0.143) or 0.00033 lb/in2 (0.000666, 0.00099) Tip weight on trusses (WT): 1 lb (400, 1000)

(c) Space Based Radar-R (SBR-R) Power spider weight (WN): 10 lb (15) Lens thickness (TL): 0.125 in. (0.086) - Truss face width (A): 300 inches (150, 200, 400)

(d) GEO Stationary Platform - Tip weight on truss ends (%VT): 1400 lb (1200, 1500)

b. Analysis Iterations. The following list of iterative values for the more significant OPTOTV parameters provides an indication of the accuracy of the printout results. - Primary strut diameters (DD): 0.1 inch - Primary strut wall thickness (TD): 0.005 inch - Truss length (LD): 10 inches - OTV mass fraction (DV): 0.01 - Alembrane thickness (TG): 0.009 - Final thrust-to-weight (TF): 0.04 4-13 1 . c. Mission Configuration. The baseline mission configuration (or program option per Appendix 6) is ZZ - 0 for one payload, one OTV, and one Shuttle. In the SBR-A and geoplatform analyses the payload is assumed to be pre- assembled on the OTV for Shuttle delivery, while in the SBR-R analyses it is assumed that the SBR-R and OTV are delivered as separate packages in one Shuttle flight and are mated in LEO.

4.8.1 SPACE BASED RADAR-A ANALYSIS RESULTS. The following explanations and comments on the SBR-A analysis results in Figures 4-10 through 4-26 are intended to clarify the more important curve trends plus interrelationships between parameters.

4-14 14 u 16.2

m 12 40 - 16.0 A a 10- 214.8 Z 11 39 LU 14.6

Q at v 0- 14.1 -LL = J T

W 38 O14.2 IL W 21- > 14.6

0 L 37 THRUSTIWEIGHT ITWI Figure 4-10. SBR-A engine thrust, propellant weight, velocity increment, and specific impulse vs. TW.

lii w l','i 48 J 11 i l l Qo ^ '^^ II I I I 9i ^l^ II III g 2f X y 46 } ^tE z ~^ Q LU ?#1^ oJ J O i 43V 14 _11 11M) I 11111111al .006 .01 .02 .03 .04 .00 .10 .20 .30 .40 .60 THRUST/WEIGHT FINAL (TWI Figure 4-11. SBR-A OTV loaded weight and payload capability and weight vs. TW.

These figures show representative OTV performance Parameters vs. thrust-to-weight (TW1 for an SBR-A baseline payload. Note that the payload capability (WY) constantly it eases over the selected TW range; however, as will be seen later, the size of &.e _:A does not follow a similar trend. The parameters in these figures define OTVs that have been optimized at specific TW values along with mated SBR-A baseline configurations. Improvements in specific impulse (IS) and reductions in required velocity increments (V) are among the primary contributors to shapes of the PW, WX, and WY curves. 4-15 am

N _^ ICI +^ I I ^ ^^ ^ a^Ot ow

u i D THRUST/WEIGHT - FINAL (TWI Figure 4-12. SBR-A. 1 payload, 1 Shuttle. WS = 60, 000 lb, diameter vs. TW.

The largest SBR-As that can be delivered to LEO by the Shuttle (having a 60, 000 lb payload capability) are identified here as a function of TW. Primary strut diameters and wall thicknesses resulting from SBR optimization are shown parametrically at several points. In this case, the entire Shuttle's weight and volume payload capability are used.

This curve is intended to provide an upper reference limit to SBR-A size that is achievable exclusive of OTV payload capability.

Examination of the D1/T1 data shows that as TW increases the optimum Dl and Ti increase. These increases are made in a manner that causes the full Shuttle payload weight and volume capabilities to be used for each TW.

4-16 4^( ENGINE THRUST (LB) 1000 2000 3000 —^—^ 44( 2.4/.06 /.0ls 2.3/.Ob 40( DI

iD/R12/.05 DIAMETER39C OF SeR IM N - 5 ^+ ^' r+ i ^. 32( I 3.4!.08 2, y .^'^`L • .MINIMUM TUBES •INCREASED STRESS; 28( m'» t • INCREASED GRAVITY - • LARGER TUBES I ENGINE n LOSSES ^^^ I THRUST (i • REDUCED ISp :.: I .. I .... i. 24( %II 4 11000 LB) i^ i l i (,^ I, I I I ^ i I 20( D1 - STRUT DIAMETER I THRUST i i T1 - STRUT WALL THICKNESS. --•— WITH MISSION LOSSES 160 d .006 .91 .07 .03 V4 U6 Me . 10 Z 7 .4 .6 4 1. THRUSTIWEIGHT, FINAL

Figure 4-13. Effect of engine thrust & number of burns on size of SBR-A.

with the OTV and payload In one Shuttle, this figure shows OPTOTV-generated optimized (maximum) SBR-A diameters vs. TW for the SBR-A baseline and the SBR-A baseline with N = 5 and 2 burnt, The largest diameter for the baseline configuration is developed at about TW - 0. 07, at Tai' = 0.0S for N = 5, and T%%' = 0. C9 at N = 2. These curves also show the relative magnitudes of the mission losses (dashed lines) as well as the TW values at which the payload is limited by the Shuttle's weight and volume limitations.

A partial explanation for the shape of these curves, referring to Figures 4•-10, 4-11 and 4-13, is as follows:

For small values of TW the OTV's payload capability (WY) is low and stage weight is high. Both of these factors tend tc make the SBR-A diameter (2LA) emall and generally cause the OPTO'rV-selected primary strut diameters (D1) and wall thicknesses (TI) to be at minimum allowable values. As the TAY is increased between (). 01 and 0.07, the rates at which WY increases and WY decreases are more d(,minant in increasing allowable SBR-A weights and stowed volumes than the rate at which increasing engine thrust (TT) causes increases in required structural dimensions Dl, T1. These in turn increase the payload weight NW) and stowed length (LC). This process continues, with increasing TW values, until the rates of increase in T: and decrease In Wti sufficiently decelerate to reverse the trend. With TW values greater than 0. 10, gsmrovements in OTV stage perfonr=cc tend to be small compared to required increases in D1 and T1 to carry increasingly higher strut buckling loads. Therefore, she size of the SER-A decreases markedly with increasing TW values Leyond this point.

4-17 THRUSTMIEIGHT - FINAL (TW)

Figure 4-14. SBR-A. Mission dependent losses, bun time, and mission time vs. TW.

The mission-dependent losses and burn and mission times vs. TW are shown in Figure 4-14. As expected, these related parameters decrease with increasing TW.

THRUSTMIEIGHT, FINAL

Figure 4-15. Effect of dynamic factor (Kg)on size of SBR-A.

This graph shows the effects of changes in the thrust amplification factor (Kt$), Ampli- fication factors greater than 2 are not likely. If K~Iwere close to 1, the optimum. TW would be closer to 0.107 than TW - 2.07 based on Kp = 2. DIAMETER OF SBR (FT)

THRUST/WEIGHT, FINAL

Figure 4-16. Effect of tip weights on size of SBR-A.

THRUST/WEIGHT, FINAL Figure 4-17. Effect of lens density (WL) on size of SBR-A.

Figures 4-16 and 4-17 show the effects of truss end weights (WT) and unit area weights of lens (WL).

I ^ 4-19 I

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200

160 .006 .01 .02 .03 .04 U6 .08 .10 .2 .4 .6 .8 1 THRUSTIWEIGHT, FINAL

Figure 4-18. Effect of truss material on size of SBR-A.

The size penalties resulting from use of aluminum as a construction material are shown here.

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Figure 4-19. 5BR-A. (LC) length of stowed payload and (LP) length of OTV vs. TNN', compared to cargo bay upper length limits.

The packaged lengths of the SBR payload (LC) and OTV (LP) are shown here vs. TN and compared with the 5 71-foot upper limit available in Lhe Shuttle cargo bay. 4-20

= 0 m J 8 O

a O J 1-

THRUSTIWEIGHT - FINAL (TW) Figure 4-20. SBR-A. (PC) critical longeron buckling load and (P) induced longeron load vs. TW.

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Figures 4-20 and 4-21 show the OPTOTV- generated induced (P) and critical (PC) longeron loads and the stepwise increased D1 and TI values related to PC. The OPTOTV analysis strategy is designed to select D1 and T1 values that minimize structural weight and maximize (LA) area.

4-21 480

440

400 N-9. I - - - -

Ir 360 II I {III I II ,! I E Illlll I I I , ( I ^ 1 t llI I'I I I ' ' U. 320 N'2 I l ll. j, I I I I I ,_' O I I 111.1' III II I ^ I I I ,il II I I I I I II II ' I I 'I^ I ^' I 1 1 I ^^ II. I I C I II I ^^ I I II 1 I I ^ I II I 11 ^.1 I,. ' , II !^ i ^ I 1^1 :., I J.^ W 280 I r^ i I ^,II1 ^I I II I I 'I 14 'I•'l I I ^^, I IIII II ,Irl I I I I I l l l l 11 I I ' l l r ^ l I I II I I!' ' , 1 I I II" 1 I I ^; I I I I'I 1 I l j l Iri i ' O 240 I..r I .• _ `r.' I _ ! 1 ,I : I I ^ ^ 71 1 I I I '. I I. .I I.', li II I .,I , I ,I I I I III ,^. I I II I r' I ^ II, i ' ^ j I I I (,I I 200 ..I. ^!^ I 11 I 1 IIII 1 I

^. 160 ^ I ^ t I I I ^ I i rl .008 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUSTAVEIGHT - FINAL (TW)

Figure 4-22. SBR-A. (IS) specific impulse held constant at 450 sec and N = 9, 5, 2. Diameter vs. TW. Effects of holding the specific impulse (1S) constant at 450 seconds is shown here. When compared with Figure 4-13 it is evident that the greatest effects are at low TW values where, per Figure 4-10, the normally computed IS values are lower than 450 seconds.

DIAMETER ENGINE OF SBR, THRUST, FEET 1000 LB 560 16 ! II 100K) /INS' I I I!I I^, I ' I 1 1 : I .I I ',1 ! j%j i I ^ I ,; r I I I r I' I I I I 'u ,, i' I I;:, ,l ,aw..; 520 14 I 1 I( 1, 1 r 1,i 1 n t ^ ill' I I l l l it r t ' '^ I ^ j l ' I I II II I l j I I I I , I I I I •., I I r I I I I I I I. I, lil! I ^ , I I I' ' I I ' I î I 1 I ' it !I i t I ill, 11 480 12 30% 1: I I I ,r'i t ^ I, r I Î ÎI ,, I r I I^ , ! I 1. ' I ' I I I III I II I r r' I!! I I il! 1 II i , I 1 1 i I I I I II I I !Ili I I I II I I I , I r_ ^ 11^ 1 ^ I, I I i 440 10 , III ?'11 r II r 4 ,p f ^ rri I. til l I i I I I' I I, I f ,^` I ,^ I'll I,II ^ ;: II I I ! i!II I I III I I ' I ^, III. I ( i ' ^ III l I i, IIII DIA 1 I. ,,. r i t 1 1 j ! Ii •^;I •I r!II III I 400 8 «-__ I'II'I ^-.'' +^.^ I I I I I I ll' III II '1 ' I' 1 '' I III ^ i, I II ;f it I 1 I I I ! ^, I I 11 1 1 I'' I I' I ' 11r I I; i'l III i'I I i'ri I I il l I ! I ! I ill 'II Il l t is ll I I 360 6 65K) t li (Ili I t l If,l Ill! THRUST!. I ' t I f I°I it I' (WS r^ 111 II I IWS- 100K1 II I "' ' I I I I } tIi .l ' I I1 I I I I I ^ Îi I I I I ni I p. r ;, II' I I ^1' I I t I ' I ,r. I I l l !' I !I n ^, ^r Im r ^ ♦•} I''i + ,I ^'i I 4 +- ^ ^ tY( "Y I Î I ^ I ^^ r ll !I I I I I 1 I t l I Ili I III I l^ I. `; I,; BASELINE II II li l 1 l 'i II'Il 'll ,,I, r III I;Î' II I. 280 2 , III , I, ;^ ' + I:.. I •THRUSTI' ,1': "'^ t 1 ^, III I II Ir (WS' 65K) I -7 - 7 7 1, , ' , I I I , 240 0 r III .006 .01 .02 .03 .04 06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUST/WEIGHT - FINAL (TW)

Figure 4-23. Effect of Shuttle capability (WS) on size of SBR-A. Figure 4-23 shows the significant effects on SBR A size and OTV engine thrust (TT) of increasing Shuttle payload capability ("i

4-22 w 48t ''•IIlii I CONSTANT THRUST N 9 (BASELINE) I _ i ' i. i t l ' f ;;i I ,.::. ^ i I 44C jj I il^! I^L^^ ^ f I ^; ^ ^ , I I II II ICI ^ i l, ', I I^ i l 40C 4 1 \ y _il _,DL _. 1 ^ I I ;', l i. fl'i I I 'I I l ! I I I ` CONSTANT T/W. M • ? 36C ^ ^îi ^ 'IÎii ► I r I ^ ^ I I I I I^ ^'I, DIAMETER I 1 I I;I I Î I ^' I I ' I ;' ^ I ^ ^^ II^ '^^ '' Illi i !Ip OF SBR, FEET jji ili^l, j ^L' ;j'I;1111!Iii 11 it ''' I,li i 'il jj 32C CONSTANT THRUST N 2 . f II li i , I , j i I j I I I I' ^ '^ Il i I IÎII ' ^ ^ I I I II I ^ i '^ j ^ i• j II' I ^ I ^Î ^ i l,; ^ ^ i I I ^ ^ I ^^ I• I I Ill ^ I i ; j I I I ' I i I l ^ ^ l I l i^ I, 28( I a I ^: I ;— 1 III II^ ill I ^ I I ^ ,I ^Î^ +t,^ ^^ ^f^ r I 1 i+' I,^^ I^ I^,, I ^! I i l I f ' ^ ^î i 1 I f^ I ,•, ll lll l ;l I III II I, ^^ , I I,' i l l. I I ! ^ I! III Ii j ! Ii I 'lid ^'I ' ^ ^ .J I î I, ^ { I ^ + ^^ ; i,1 Ii l l ^Î i,I;,l ^ j ^;l IIIÎI I i;;i , ^ i ;I I, ;, ^, I 241 i ► ^ ; ^ ^ III ÎI^ I,II ; li I II I j ^ I j II I 'Il ^ IIIIIIi illl lil i j ' I I ilil^l i III: IIII III, I I I I I I I I Ili ^ i 20( l i +, I;+ Î I 1 I' 1 1' , ^' i I; I I i ', I I^ i ^Îl it l ^ I. I i ; I ! I! '^^^ÎI I, I II , Ti, : I 180 !^ ^^ I!i I ^ ► I^ I''I .006 .01 .02 03 .04 , 1.00 THRUST/WEIGHT, FINAL Figure 4-24. Effect of constant acceleration (variable thrust) on size of SBR-A. Shown here are relative effects of constant TW and constant-thrust OTV engine performance for N = 2 burs. It is evident that constant TW and N = 2 can almost produce as large an SBR as the baseline configuration, but is much more sensitive to thrust-to weight.

I I I "0 j I I ; i I 10% 400 ^.

360 DIAMETER OF SBR, FEET (BASELINE) 320 - (-30 SEC ISp)

280

240 _ _. _._...... ^_ {... DIA' ^ I i 200 I i

160 .006 .01 .02 .03 U4 .06 .08 .10 .20 .30 .40 .6 .80 1.0 D THRUST/WEIGHT, FINAL

Figure 4-25. Effect of reduced engine performance on size of SBR-A.

Reducing the computed specific impulse (IS) by 30 seconds has an important effect on reducing SBR size (MA).

4-23

2SHUTTLES 0 FOR PL) . . II L. I I; 11 FOR OTV) I -+ --

^i, I i 140 I ^^ I ^^ DIAMETER ^ I , l' I ' OF SBR, FEET ^ I 400 ^^ ,Illlill

I !I, I 1 SHUTTLE (B6K)' 360 (BASELINE ► I , ` I T t41- -- 320 ICI ^i^ 280 it

THRUST/WEIGHT, FINAL

Figure 4-26. Effect of number of Shuttles on size of SBR-A-

Figure 4-26 shows the effects of using two Shuttle flights, one for the OTV and one for the SBR-A, The SBR-A sizes that can be achieved are smaller than those in Figure 4-12 in which the payload limitation of the OTV is not a limiting factor as it is in this case.

The graph also indicates that the LSS sizes that can be achieved are smaller in terms of total area than would result from two baseline flights (OTV and LSS together in each Shuttle). Two baseline flights would produce a 200 % area increase rather than a 144% increase (which the indicated 20% size increase indicates). This neglects the docking requirements for two separate payloads as well as stowing efficiency losses and payload losses for rendezvous AV. Further study would be needed to assess these requirements.

ORIGINAL PAGE 11 OF PuOR QIJAI. TY

4.8.2 SPACE BASED RADAR-R ANALYSIS RESULTS: The SBR-R analysis results resemble those of the SBR -A as the following figures indicate.

4-24

41, 46C Pi 11 'll ) , I I I II 1 1 1 111 - I I li! ! ' ^ II IIII I ' l l . ^^ l I 1 I I I ! ,15.2 1 I ril; ro ^^ t + LM I^I'^'f I, II II i1 j' II ; I ICI ,I,I> II ^ I ►

1: 00 IT, CO 40 15.0 45C .j I IS 000 -: 14.8 L4 IIII ^^ ^ ij ^ -^ I ; ; J , II I ,li' I C I 1 i Iji I I^j ^ 1 ; ► ► , II I I III I 'li I ( ^ ^ 11111 z I I ; ' Il i ^ I ► I I ► 39 - uj 14.0 44C 2 U 4, jl L UJI f 111iII! ! rY C11 q" IVI I' I l - U z 14.4 Q III I I1 II.I ' .j I I I I' , II ^ ^ I ^ I^ z LU lil t Ij I IIII l I I I I I II I I I L I I II'll' " I i 1 I III ^^ Ili II -4 WTt I u' 38 -814.2 35 43C U t II L -j I I III, ' I^ l I III I I 1 ! i I , ^ I LU II 11 1111 I ^ 1 I^ I (^^! I ( ,i I ► 1^ liI' ^ I t I I - ^^_i r• i ;;I^ j I;.,( II 1i:. III' ^^II ^'11 >_ ^ ^ - ' ? 1! I Ilrl !1►, iT JI I III III I III I i l'' , III ill I^ . I !III I•I '^ ( I I I I!l 0 37 L 13.8..8L8L L 20 .ON .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 10 THRUST/WEIGHT - FINAL (TW)

Figure 4-27. SBR-R. Engine thrust, propellant weight, velocity increment, and specific impulse vs. TW.

461 17

i ^Wx I I1

I, can I ltll Itl II^;{Il ^ ^_ ^ Iî II:Î(IIII I L I ^ ' IIIIIIIIII x 45- I 1 1' II F II^ -- I ^ I Il 1 ^^^`, IIµ: ( !:r I { ^ ^ I, l I II , I'll (1 ^ IIII III; ' I I;' I, I ^ Ii^l II; II I i II I II ;I'I III t I I t j I n I^ ^ i i t I I' 1 1 i i , !^ îr

CL II III, I 44 LU it I,I II,,, -^ -1 F ^; ' : , , ; { I ^, I,1, , r • ., t,ll'H1. - ,- -^ n,, rif l;! LU 7 I I II,;,1 ► r -^ I I;I, t'1111 Î III' ' i 1 1. I I ^ 1 1• (I 1 1 ^ 1 11 I 1 'I I r' II I rl I l ^, I I IIj I I I I 11 I, Î I qll f I ► î

> 0 14 :^I ,I, I I ;, II,. I ! I ,,:I 43- I It .;II ► ! i lI ^ ^. „11 :I1i , I' _^ t t :;^ a.. I I II I Ili'' I 1,11 II ' I !^ Irll I I^1^ ; I ^ I , I I , { ^ ^ II I 1 ^I 111 II II III , I I II' I ,I , ;;I ;

.006 .01 .03 .04 .06 .08 .10 .20 .30 .40 .60 THRUST/WEIGHT - FINAL (TWI

Figure 4-28. SBR-R. OTV loaded weight and payload capability and weight vs. TW.

These are very similar to Figures 4-10 and 4-11.

4-25 ENGINE THRUST ILB) 1000 2000 L000 8000 8000

I I I I 80 ^ ;,. I ,l I I ^ ' i I n II ^. Imo..... 40 OI/TI II ^ l i III 3.5/.On,I l 2.5/.050 _; N 9 00 it .120. _ ._ BASELINE f T,1 W DI W il' I ^ I II I N-5 j 4.4/140' . U. it i I I ^^ i l (I I I N-2 5.0/.155 ^ II , ^ ^ LI,; n I I I^ ^ I O . 20 -- — 8 m III Cr I I W j I li li il I I 'i j' 11 gJ W }III `ll l so 1. I I , I .. 6 a ^^^ fl' II^ `h ^ ^' il^ ,. ,11 I i l , I !^,^; I , ^ ^I ^; fl Il ,^ I„! ^ i ,^, I I ^ i it p ! III lill! ^.^. i', i ^ , - I_ is III I;. ^II I I III! Ij I If N 40 N^ 4 Q ^,I l ail I ! n THRUST I , i ^ i i ^ I I I I^ ^ ,'!^ x i I I ^' I r II:I ;I 00 ! . ,:i. 2 Z v ,, I Z I I i W so 0 .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 .60 .801.00 THRUSTIWEIGHT, FINAL

NO EXTENSIONAL STIFFNESS ---INFINITE EXTENSIONAL STIFFNESS (LENS ELIMINATES STRUCTURE BENDING)

Figure 4-29. Effect of engine thrust and number of burns on size of SBR-R.

In comparison with Figure 4-13 it is evident that SBR-R size variations (2LA) for different number of burns (N) are not as TW dependent as they are for the SBR-A. 'Phis is due to structural (stress) considerations and the relatively smaller percent total LSS weight in the truss structure.

4-26 ^" I,, I uo !! I I ! ', I;; d i I I , ' II jl I 100 "^ .! .. ^ _ I I " ' 'I' WL -.049 IkIASELINE^ W I I DIA, W I W l i ( , II 1lI I ; . I ^l l ij, } i^',' ,^ ; I I i Ijl '' '^ i ,; m¢ I ; I I I;II I I.^ I I ^^^ I I I i H j) I I I I ^^ II I !,I' IIII ! .II I , 320 III it }^ a O WL .099 t ii lI I III I , ^ 1 I W l ijl ! I , , I I I ! III ^ ' , H W too (,I , I .I I L. t .. i . I. I I , ,^, II ; : ^ I I ^Î, I I , j i^ I I C I I I l i I I I ! ÎI I I (^ II t I I .^!' L, I Î i j __. _i,_ ' I ^ .L. 240 IIL; III I; III' III ,, I I I I I I 11 jl II _I_ 200 {I{ IIII } : i } ^4 f I1 I. t I.- I I Î' I I' ,III I I I' I I ,III , I i, I I I ,, I II I I IIII 'l I !' ' I, ^ ' ' ^ j I I l i I80 I I. I .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUST/WEIGHT, FINAL

Figure 4-30. Effect of lens density (WI.) on size of SBR-R.

,. I I II I, I^ ;I I I ' i I.^ I III I 1 , I I ^ ^ ^ ^ ' IIII I' IIII I'll j , , 'I I I 140 t I lI }I,i r ', ^ I ,I ; I II I n ^ ,i 1'f' ,Î I. I l I I. I r II I W BASELINE t I I! W I; INI - 10 I I W 400 » ». » ' I , I I I I WN.16 11 I I! ;; I J î I I I ', i l i iY, I I N 380 1 I I ',I , If I^ Ii i' uII rill i,l:; Ih,ll: I L. I II I 1 I '.I i ^ I , I i ! I I ^ ,I ,I, ,^ ,' I I I I I ^ i.. ', 1 II I!II ( IIII C it l I +h i ' , I I, 1 11 i I'I ++ ! a0 ^ î+ ', i ^.. N 320 Ci IP'1 I,jj T JI, I , U. ^ I III Î ^',I II I .:II , I I II^ II ( I j j I +I lj ,III O I Iljl li -, i i 11 I I 1 I ll i li lt I ,..^ III, j ' ^, ' I l I I llj j I it x 280 ^ I . r l I III! " I ! { l.: , 1 } I ^{; I I I. W 1 ^ t t .h ^t }rffI I, I' ., ^ ^ ^ r W I I II II! I I ji I I j ^ I I^^ „I 111 ' j, {I iI; II I I I ': I ;II ^ W i ^lli,! I I II i I ' ' I I i' i j' I I ICI , I I I^ I ^ ^Î I II I i I I I I I^Î i1 , 240 rr .t tt, a l «- 1. I 1 + i, I , I f 11 I II `r^^ì ' I j,I ' ^ '„ I ' j, 1 I l l i I I II;I I r I I I i I ' I I . I ^ I ,;; 1 ,.I II ,I I , too

I 11 , I ,i i I I I I I I I^'I' , I', I I '!' w 'll u !j I, , 160 I 006 .01 .0Z 03 .04 .06 .08 .10 .20 .30 .40 .60 .801.00

THRUSTIWEIGHT - FINAL (TWI Figure 4-31. SBR-R. (WN) power spider weight = 10 and 15 lb. Diameter vs. InV.

Figures 4-30 and 4-31 indicate, respectively, the influences, relative to the SBR-R baseline configuration, of changes in lens unit area weight (WL) and power spider weight (N%N).

4-2i 440 WLU LL 400 J .N. 360 Gof 320 LL O cc W 260 W

240

160" .006 THRUST/WEIGHT — FINAL ITW) Figure 4-32. SBR-R. (TL) lens thickness 0.086 and 0. 125 inch. Diameter vs. TW.

10 THRUST/WEIGHT — FINAL ITWI

Figure 4-33. SBR-R. (ZH) hub weight fraction 0.47 and 0.65 inch. Diameter vs. TW.

Figures 4-32 and 4-33 indicate, respectively, the influences, relative to the SBR-R baseline configuration, of changes in lens thickness (TL) and hub weight (ZH).

4-28 •aw I I I ^ I I 1 I I' I l il ' II I I ii ^ ^ i I i I I ,,: I ^ ^ ! I 1 NO I 1 III I II i. Ili i l^' I li Ili I I I I I I I I I IIII q I I ' II .A-^00 'I I W '6ASELIN W 100 ! + Ill • jl. _ III' Ili l l l i A 300 j i it ^ I ^ I Î I ^Î I I .liî j ll l I I !^' l l i l , I '' I I 1 l il IlIijll'ii 3 ^ ZVU I ^^ l l^ II Ii' a itl I I IÎ ÎIAI' IÎ I ' ^'I Q Itli l I II' I II i II I20 il ^.^ ( . j, ,. " 1 1 II I I l I^ 11 'I',. ^ I 1 - I I: I I 1 i, , I^ I ^ ';I I I 1 0 I) 1 1 I t i f l ill. t60 I I ( I ll^ - j I I l I li I 1 1A1 •i ^j^ i `ti iI^7 i^^^ III III I I I IÎ III I ., I I I W ,I 1 1 Î II ICI I I I I 11 . I .III I I I'II ^ I I ^ I '..I I I I j 1 , I^. r ' I I I II! I ^ II 'I ^ I I it II^' ' I I I ' I I II„ 1' l'I W li I ill III ^ I^ î I III I l i l t I I I^ I I'I^ ^ II II +I .! li! 1 140 1'}' ! t *^ I l,l r •- ^w I^. I ' I I Ali' l i Î' I '`: I^. it I IY I I I I I' III' I I I I ÎI( ' I ^ I I ( ' i i lI' (I t I ' III I I li I ^L. I ../^ II Il i l'IIIIIIil,l^l .I il^ I bo l l III h^" I F1^ I14 i I) 1^;III II II 1 !^ I ^ I ^' f 1 I i l l iÎ l ^^ I^ iltil^îÎlll'^'I^ ^ l illl I it i ntl I Î I 1! (1111 I i III i I 11 I I I I I I I ^ I l,l ÎII1' 160 I .006 .01 .02 .03 .04 .06 .80 .10 .20 .30 .40 .60 .60 1.00 THRUST/WEIGHT — FINAL ITWI Figure 4-34. SBR-R. (A) truss face width = 150, 200, 300, and 400 inches. Diametcr vs. TNV.

The effect of truss face width (A) in this figure is explained by the fact that, as A is increased, a greater volumetric portion of the Shuttle cargo bay is used. The size of the SBR-R is accordingly increased up to the point at which SBR-R and O'1'V weight and performance limitations become significant factors.

410 i I 100 _.. I. G/E (6ASELINE) DIA W W W 380 3.6/.050 m D1/T1 ' N 6085 ALUMINUM I' W 0 320 1 _ ^. 5.2/.100 }•. Q W h 15.9/.120 W 250 I ... I I' 1 t' I I I i,; I I I 240

200 I.is 160 006 .01 .02 .03 a .06 .08 .10 .20 .30 .40 .60 .60 1.00 THRUST/WEIGHT, FINAL Figure 4-35. Effect of truss material on size of SBR-R.

Effects of change from graphite epoxy to aluminum as a construction material, as shown in this figure, are comparable to those shown in Figure 4-23 for the SBR-A. 4-29 s 4!( i, ! +' i 1 !.; I 1 ,' ill' I,': ! I

r W 44( PO W ^ 1. 1 !.; Î ,^' el I !I ' ' îil ' ..i I ( ,'' 1 it II c I ,I I II , 11!^ 1'. ,I I ^ I^ ( I ll' I 1 ^ Ij +^,II, I I ^ ^j l 'll j l!I u _ KO.) jf i Î iiI 111 + 1^1 i i ITI J 38( { III ill , d ^ 17, f t. 1I I f .I I I I ^^ , I^^ ^' I I f ^^ I !, + ! I ,' II i^l BASELINE ' 1 + j I l i I ! , ! I ! (I I +^' I ^.. " II li 3211 KO- 2 1 1 1 1 1 1 i I I I I{ I+^ +11 I I I, l l I I ! , i I i I II , ,,, I, 1 ^ .I ( Ili! i f ^ .1. i I !111: ^ ^ E ,^! I'^ '^' ^ ' 1 j {,, I^, 0 I; !, ',I ^ i i 1! f i ^ I, i ,, i ^l'I ,III E r ^ 1 ¢ 284 I II , 1'' I I'. f 1 I ^` ^^ 'i, ICI d } i• 1 Iji. ^ '^ 11 ' ' I ^ I 1 j { ! 1 2411 " I jail ! 11'', ^;i i l , 0 l' '1111! ^! ;, ^' ! L ; I t I , I' .i; , I , i^ 111 'ir. ^^+. ; , 20 t7 , 7 ;

18 .80 .80 1.00 .008 .01 oz w w .w Ma .lu du .30 .40 THRUST/WEIGHT — FINAL (TWI Figure 4-36. SBR-R. KS6 = 2 and 1.

The effects of reducing thrust amplification (K95) from 2 to 1 for the SBR-R are shown in this fLgure. The benF.Sts of low thrust ampliScation tend to occur at higher TW values in this case than they do in the SBR-A because of differences in structural characteristics and ^mailer traction of the overall SBR-R comprising the load-carrying structure.

4.8.3 GEOPLATFORM ANALYSIS RESULTS. The following explanations and comments on the geoplatform analysis result; parallel some of those for the SBR -A and SBR-R. 4-30 16 411- t6.4-

I 14- 15.2 I f 11 i m J ` I j Ile iS j 12F $ 40-1 1S.0^ J 460 , .^ LM ^- ^10^ ^ X14.6, -N F I [[.... W ^i X ^,._rt ^ 8 36 W 14.6 440 4 • ;• .a^ hr -+ .i:a 1 1^.:., r ; Cr X V'' W 6 ~ r Z 14.4 Wv i'I ^ ^ I t I i ^^ i J ~ ^z W c^ W I W4- o 14.2 $ 38 430

IL > I 1. i, 2 14.0 I ,. ^^. I TT iI ,.ate I_

37 L- 13.6 120 .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 Bn THRUST/WEIGHT - FINAL (TW) Figure 4-37. Geoplatform. Engine thrust, propellant weight, velocity increment, and specific impulse vs. M.

46 m 17 10 J J Q ^ Sm g } O 45 16 } H ~I W m I

44 V 3 15 O Q O 00 0 O o d 43r 14

.006 11 .02 .03 .04 .06 .08 " .20 .30 40 .60 THRUSTIWEIGHT - FINAL (TW) Figure 4-38. Geoplatform. OTV loaded weight and payload capability and % •eight vs. T\V.

These figures are similar to Figures 4-10 and 4-11. The same explanations and comments apply.

4-31 324 ! I I I I u l d I I 281 ENGINE THRUST (LB) IN Dj 20M 30004000 6 8000 244 01

9, LINE 1 7 .- LL 20( ,2.2 /.M iii ÎI II 2. 1 F. 05 u I I„' I l I IIp I I III Îli''II I,^ X 1 I I II 11'^;, U. Illi^ ' i ^ I ,, .11 II 1 1 .. 0 161 77 I II I I Î :I I I, ^,p I^ 1 1^ I , C ` ' i - ^ I I I I I.I I, , ` IIII I I^^ I -J_ LU i Ì 1 I IÎ

12( :' I i,:.. 6 'II^ IIII ^'1 ^I Ilii ^1 I 1 i III I ^; I,, I ^ D1R1 I ^; ^ ^ (^_ ^ Ij Ii 1 I I ^ : ^ 8( ": THRUST 4 x I I I I ^ I i^: III f II I I l i L I, I ^ II ^ I , IIII ! ; I i LU 4( 2 + Z I , I 1 IIII I I ^ I :. II ^ I 1 I 1 5 II Z W 0 , .006 .01 .02 03 .04 .06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUSTIWEIGHT, FINAL

Figure 4-39. Effect of engine thrust and number of burns on size of geoplatform.

These curves, like those in Figure 4-13, identify the TW values producing the maximum size structures for number of burns N = 9 and 5. In this case, 11V = 0.13 produces the peak geoplatform. size (compared with TW = 0.07 for the peak SBR-A size with N = 9).

Note that the structure sizes (2LA) are considerably smaller than those for the SBR-A or SBR-B. This being primarily due to the larger truss end weights (WT = 1400 lb). Fall-off of size with increasing TW above 0.17 is also not as pronounced as in the case of the SBR-A for TW values above 0.07.

4-32

280

= 240 h . ! --^.t _ .. ^.: DIA

W I I ^ i I I I I i .. a 200 I .'.^. I L. WT - 1200 (BASELINE) 160 l _ I' ¢ 1 I S. 1 WT 1400 W I ^ J I ' I' I W 120 00 -1500I . I WT i I

I j I,. _.. I I I I

I' I 40

' I A- 11 .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .410 .60 .80 THRUST/WEIGHT, FINAL

Figure 4-40. Effect of tip weights (WT) on size of geoplatform.

W 280 W f I .. W Q J 240 + ZH-.47 i 2D0 j : BASELINE I LL ZH .66 a I aJ 160 _.. O W 0 U. 20 ^.. O I i ^, I ^ i 1 i I I ¢ 'II W H 80 - i .. 1.. I W ^ Q I it I I

a fl i

0 1 .006 .01 .02 .03 .04 .06 .08 .10 20 30 io so Ro 1 n 0 THRUST/WEIGHT — FINAL ITW)

Figure 4-41. Geoplatform. (ZH) hub weight fraction = 0.4? and 0.65 inch, WT = 1200 lb. Diameter vs. 'nV.

Figures 4-40 and 4-41 shcw the effects of changes in truss end weights (WT) and hub mass fraction (ZH).

4-33 R 321

281 I i .I II I I I j I I I II, II !^^ ^ + 1 I I '

24( j I :,' III _}.DIA^ H I i^ il^ '. i II ^'jl ^ i , , I I W II ^ I III{ I ' ^^ ^ ^ ,I I I I I i Il ^ l l r I I I ^ ^: W ^4u I l' 1- } { I.^^ i ih I j •I ; ^ j I I L, W 20( i I ' 'i t I Ir ^ , I r I I I I I,'I Î i'I I I I I I . I i I i I I 'II I11 j,'I I 'I I I C7 (I^ I II I I ^ I ! I II Ij I ul I II i..^ j_. I, it li I^ I I, I I^ ! tj ^ II. I I I,,, I i + ^ C 16( _ 7-1 j.l I l l I, 1 1 1 1 j ^' BASELINE Q jllj f li III j j l l Î^ GRAPHITE EPDXY I 12( 1 :1 i.. up 2. 1/050' fI j l I l "" 2 0/.050 I Q 2.6050 I 2.0/.050 III + I Î L -- I 7 8( ' --j-^-- . I D11T1^- ^! I'II I I , I t III II I ', ^ I II ! „ - •i i 11 . L' I,I a ^^' I ll ill ;,I 71 ,I^ I I' ,LUMINUM .^ ^^ ^ j III i cl i I îl I II ( I I 4( ^ll I I ' ,2.0/ 050. III j l i I, jil' I•I i I I I I I; ^ III III! Îjl I i jl I ^ j i I .I 0 .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUST/WEIGHT. FINAL

Figure 4-42. Effect of truss material on size of geoplatform.

Influence of aluminum as the construction material is shown here. Compared to the SBR-A and SBR-B, the structural weight, in this case, is a relatively small fraction of the total LSS weight. The OPTOTV-selected strut sizes (D1 and T1) also tend to be relatively closer to their minimum values. As a consequence, the percent reduction in size due to use of aluminum is greater than in the cases of the SBR-A and SBR-B. If the minimums, DM and T1VI, were reduced, then OPT OTV would have selected smaller D1 and T1 values (i. e., D1 = DM = 2.00 in. and T1 = TM = 0.05 in.) between TW = 0.01 to 0.1.

4-34

n e 32C

l ; , U'-1 29C ;jI I ;Ij;N;I'LI,;I Ij ! i i l jl;;!! i' ;.I TV1 W 't..IJ^I'j;'jl; I , ,; j,Il 1 Iji ' ,i;; !^ i ' i i I U. I j^^ I Ij ..I t I. III ' i I ^, I ;I ^, ; I I,I, ., la; , i ^ I t ^' ^I ^ j ' I• , Il '' jI;I ^ ►I 24C ►

ki T ^l i! i t ^ Ili 11 lj!^ 1111, , I I !J 1 !;I! I 7 ! i 20C tIT It I ;T 0 ^ I I II ^ 1 ; ^ ^ t 'fl; IC I ^ ^ I i.' j'1 1^ (i ;I ► I I1^} I ^ i IJi , II IIII LL t ;^^ I I^I^,: I,;I I, II 'tltll^i, ' II I ^ j iI!l ^^ 1 1 1 II; II•• , I I. i,^ r C I '^t ^t 16( i ► ^11

0 it I l jl lli l I fll t!`j^I^ i' i '.kllllll i ' Iin_ ill jil ^ ijlll,llill,l,^i!i,i},,,,II, 12( ^ t ^ ,I h I', „, I t till ;:il 1"I ^1 ill Tt ^ I ^. 0 ,ri i t II li ^^!'. III ^ ^ l III I ; I t i I t li l I III ^j^ ^ I ► ICI ^'^' ^ i^' '«, LU

T tL i ^I T '1 ^ ► 0 .006 .01 .02 .03 .04 .06 .08 .10 .20 .30 .40 .60 .80 1.00 THRUST/WEIGHT - FINAL ITW)

Figure 4-43. Geoplatform. KO = 2 and 1.

The benefits of reducing the thrust amplification factor (KO) from 2 to 1 start at TNV > 0.15. Since the structure forms a smaller part of the geoplatform than it does in the SBR-A, the effects of increasing M, occur at larger TW values, where minimum strut sizes (DM) and wall thicknesses (TM) are exceeded. The benefits of low KO at sufficiently high TW values are approximately proportional to the mass fraction contained in the GP and SBR-A load-carrying structures.

4-35 4.8.4 SUMMARY ANALYSIS RESULTS

The baseline weight summaries at optimum TW in Table 4-2 are based on data in Figures 4-15 through 4-43, as well as Appendices 9, 10, and 11. Significant depar- tures are possible from the values in Table 4-2 depending on differences between selected mission, payload, and OTV parameters and those parameters used to define the baseline configurations.

The SBR-A and SBR-R results indicate that the optimum-thrust OTV design is at or near final thrust- to-weiguti (TW) = 0.05 to 0.07 or thrust levels (TT) of 1000 to 1500 lb.

For the geoplatform application with structural sizes pealdng at TW = 0. 13, the optimum engine thrust (TT) is near 3000 lb.

Table 4-2. Weight summaries for baseline configurations at optimum thrust-to-weight.

SBR-A Geoplatform SBR-R

Thrust-to-weight 0.07 0.13 0.05

Weights, lb Total Payload 15,930 16,799 15,204 Structure 2,069 816 1,062 Nodes (joints) 4,590 1,811 2,990 Lens (antenna) 4,245 -- 6,291 Hub (core) 5,021 6,972 4,861 Truss End, Total 6 7,200 — OTV 44,070 44,201 44, i . 6

Diameter, feet 420 220 396

4-36 n R 05

BASELINE VEHICLE CONCEPT IDENTIFICATION AND DEFINITION

The baseline vehicle selected for definition is the short expendable (torus L02 tank) configuration shown in Figure 5-1.

EXPENDABLE LOW THRUST OTV 138K PROPELLANT CW-61 MR - • PUMP FEU 16K ENGINE • ENGINE MUUNTED/DRIVEN PUMPS (NO VEHICLE - MOUNTED BOOST PUMPS) • 16 PSIA MIN INLET PRESSURE • NPSH L02 - 1 PSI L11 2 --05PS1 • AUTOGENOUS H 2 BLEED • COMPOSIIE STRUCTURE • ALUMINUM TANKS 18 FT • PROPELLANT ACQUISITION • PARTIALSETTLING • SCREENS • MLI TANK INSULATION (15 LAYERS) • PRESSUMZATION • HELIUM PRE PRESS. 02 RUN • AUTOGENOUSH2RUN • ZERO G VENT/MIXER • FILL AND DRAIN THROUGH SIDES OfORBITER • 300 SEC A B ORiUUMP 1 8 14 1/2 FT -^ • N2H4 AITITUDE CONTROL • FUEL CELL POWER 11 KWI • MISSION (DESIGNED FOR 3 g IN : HUTTLE). • 40-HR ORBITER CIO • 20-HR IHANSFER • BBURNS • 6 HR BURN 1IME

Figure 5-1. Baseline design definition.

5.1 BASELINE CONFIGURATION, DESCRIPTION, AND WEIGHT

Figure 5-2 shows the detail layout for a short (18-ft length) OTV using a conventional liquid hydrogen tank and a toroidal liquid oxygen tank. The RL10 engine (short-low thrust) was used for layout/ interface definition since new low thrust engines are yet to be defined. Both tanks are suspended from an outer body structure. A separate conical thrust structure is located between the two tanks and intersects with the outer body with a kick ring. This thrust structure also provides a second support system for the toroidal tank at the inboard side. Figure 5-3 shows a model built to approximate scale. Figure 5- 4 shows dimensional data. Table 5-1 gives a summary weight state- ment for the baseline low thrust OTV. Table 5-2 gives a detailed weight breakdown by subsystems. Table 5-3 gives a detailed summary of propellants and fluids. 5-1

N W V Q on U. W

rl / Ne^t^ "I' L W A W z zW I. W r c ^' ^^ \1 `^' va, v V y 1 r`0 W ^^ ^^ ti ^:'^ ^aZ^ y ► 3^ Lo^ ^ ^\ id^^ i^ t 33 •_♦ O \L a N^ ^^ o^ ^y- ^°.--- •^/ \h d_ J IJ 0 N

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5-2 0 F'iure 5-3.. Low thrust OTV (model). 5-3 ^^

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1.50 ^3R 1 Xo 1054.07 i 1 ; q--X0 1065.Q Q FWD 60.145 I SUPPORT

148.02 25.73 ; LH2 TANK

2V 18. 1 1 1I 247.96" 60. 145 i 57.5 20.66 FT

L_ 22.0 L02 28.03 TANK ^r Xo 1226.98 48.0 28.03 Xo 1271.59 Xo 1299.8 12R ^.• PIVOT Q 54.97R--^ -- -- - Xo 1302

Figure 5-4. Low thrust OTV dimensions.

Table 5-1. Weight summary low thrust OTV.

WEIGHT DATA (LB) STRUCTURE 2,177 THERMAL CONTROL 635 MAIN PROPULSION 782 ATTITUDE CONTROL 206 AVIONICS 388 ELECTRICAL POWER 380 CONTINGENCY 668 TOTAL DRY WEIGHT 5,124 RESIDUALS 382 RESERVES 4 BURNOUT WEIGHT 6,936 INFLIGHT LOSSES 319 MAIN IMPULSE PROPELLANT 37,434 ACS PROPELLANT IINCL DISPOSAL AV) 55i STAGE TOTAL WEIGHT 44,240 PAYLOAD TO GEOSYNCHRONOUS ORBIT (MAX) 15.760 STAGE PLUS PAYLOAD WEIGHT — —8000 AIRBORNE SUPPORT EOUIPMENT 5,000 TOTAL LAUNCH WEIGHT 65,000 MASS FRACTION 0.866

5-4 Table 5-2. Detailed dry weigh: breakdown.

STRUCTURE 12,1771 BODY STRUCTURE 816 FUEL TANK AND SUPPORTS 409 OXIDIZER TANK AND SUPPORTS 628 THRUST STRUCTURE 171 EQUIPMENT MOUNTING 40 PAYLUAU INTERFACE 83 DEPLOY ADAPTER INTERFACE 66 UMBILICAL PANE. 66 THERMAL CONTROL ( 5361 FUEL TANK INSULATION 108 FUEL TANK PURGE ENCLOSURE 82 FUEL TANK PURGE SYSTEM 89 OXIDIZER TANK INSULATION 39 OXIDIZER TANK PURGE ENCLOSURE 69 OXIDIZER TANK PURGE SYSTEM 72 RADIATORS, ETC. 86 MAIN PROPULSION ( 752) MAIN ENGINE 100 THRUST VECTOR CONTROL 30 FEED SYSTEMS 90 FILL AND DRAIN SYSTEMS 64 GROUND VENT SYSTEMS 111 ZERO-G VENT SYSTEMS 41 ABORT DUMP SYSTEMS 136 PRESSURIZATION AND PURGE SYSTEMS 141 PROPELLANT MANAGEMENT 49

ATTITUDE CONTROL PROPULSION ( 206) THRUSTER MODULES 62 PROPELLANT TANKAGE 80 PROPELLANT FEED AND FILL 49 PRESSURIZATION AND PURGE SEE MAIN PROP PLUME IMPINGEMENT, ETC. 15 AVIONICS 1 3961 ACS SENSORS AND ELECTRONICS 88 TELEMETRY, TRACKING AND COMMUNICATIONS 45 GUIDANCE 39 CENTRAL PROCESSING 45 RENDEZVOUS AND DOCKING 0 SERVO ELECTRONICS 30 SEQUENCE AND PYRO CONTROL 101 INSTRUMENTATION 48 ELECTRICAL POWER SYSTEM ( 3801 FUEL CELL 158 FUEL CELL INSTALLATION 60 BATTERY 31 BATTERY INSTALLATION 6 INVERTER - BOOST PUMP 0 POWER CONTROL UNIT 15 BUS INTERFACE UNITS 10 HARNESS AND CONNECTORS 100 SUBTOTAL 4,456 CONTINGENCY 668 TOTAL DRY WEIGHT 6,124

5-5 Table 5-3. Detailed weight breakdown of propellants and fluids.

TOTAL DRY WEIGHT - BROUGHT FORWARD 6,124 RESIDUALS ( 362) TRAPPED LH2/LO2 142 TRAPPED GHyy/G02 226 TRAPPED N2 TRAPPED HELIUM 2 TRAPPED H2O 3 TRAPPED N2 3 BURNOUT WEIGHT WITHOUT FPR 6,606 RESERVES 12% AV FPR LH2/102) 430 BURNOUT WEIGHT WITH FPR 6,036 INFLIGHT LOSSES ( 3161 ME START/STOP 136 ME LEAKAGE 8 BOILOFF GH2/GO2 133 HELIUM LOSSES 3 FUEL CELL H2/02 40 IMPULSE PROPELLANTS ME LH21LO2 37,434 ACS N2H4 661 STAGE TOTAL WEIGHT 44,240 PAYLOAD TO GEOSYNC. (MAX.) 16,780 STAGE PLUS PAYLOAD WEIGHT 60,000 AIRBORNE SUPPORT EQUIPMENT 6,000 TOTAL LAUNCH WEIGHT IN ORBITER 66,000

5.2 SIT13SYSTMIS 5.2.1 TORUS L02 TANK. A design for a 468-0 conventional toroidal tank for L02 service is shown in Figure 5-5 (Layout 59). The tank features a structural arrangement which permits access to the interior and provisions for a low conductive support system located at the inside and outside diameters. Also included is an acquisition system, a pressurization bubbler manifold, a ground vent duct, a fill and drain port, abort dump sump, a ring baffle, and a boss for an electrical penetration fitting. Internal bracketry is also provided for mounting a zero-g vent apparatus, a propellant utilization system, and electrical harnessing.

5.2.1.1 Structure. Design approaches for the torus tank are shown in Figures 5-6 through 5-10.

The primary structural members are eight rings equally spaced at 45 = , and eight shell segments. The rings also serve as radial baffles and have a tee cross section. The 24-inch inside diameter was chosen for accessibility between compartments and the webs are perforated with holes. Tests will be required to determine any effects the 24-inch access holes may have on the baffling characteristics. If required, removable perforated doors may be added to each ring.

The flange ends are step machined to fit the weld zones on the shell segments. 5-6

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1 5— 7 R • OENERALCASE • CANDIDATE CONTOURS_ AXIS MERIOIONAL CONTOUR • CIRCLE ^0 i ^1 I r 111

W—Rw.m 1 • MAK R IS OOOO • ELLIPSE H • INCREASESRI • DECREASES A • SHORTENSSKIRT L ^^ • CYLINDRICAL O.D. IS 0000 • REPLACESPORTION t OF SKIRT • KIDNEY IS CROS&SECTION SYMMETRY • TYPICAL OTV APPLICATtON NOT MANDATORY

o IF REQUIRED FOR CLEARANCE CIgES AANCE WITH CENTER ENGINE. -ENVELOPE SKIRT • WEINER t

t t Figure 5-6. Torus tank considerations - I. Geometry.

IS GAS PRESSURE ONLY • FLUID IFFECTS HOOP LOAD VS. CONTOUR 1

EXAMPLE (REF II: i • V - 1]3.1 F73 • Ro-t10UIN 1 u 1 +r^ --0 r -O -=R---^.-a--_4 1 NI IC)OP P EXAMPLE IREF 21 LO! • Ia • D1a - 3 5 ^• _ _ _ n • not • all - 400 a• U HUOP • -- -—_ — --_ ----- ASII

N • !• • 1 ! 171 1• a• cr AUDITION OF FLUID CAUSES • SIGNIFICANT "DROOP" 1 ••• • •r 1 (ANALOGY. WATER-FILLED INNER TUSkI n . • HIGH MER10IUNAL StNOING AT • • HIGH 01N)UP AND WIDESPREAD BUCKLING AT 11 PNESSURt INVUCtU N1100P IS • INHERtNT IN SOME SLENUtH CUN70UR4 Natwwwn • Vk RY HIGH IF INNER SURFACE IS SINGLY CURVED i Twowr, N M, 0"1 P•tatlwtw& too tNrpmad Tuoo1d•1 • AVOIUAbtt b y MkR101ONAL TAILORING Ptauulr Y•.w!•, Anuµ+Ka Erlyltw•ou1y, NuranlDa. lY•A 2 Sport. E E , N W. Luwat bosidi ry Atwlrun ut a Thm Tuwurlal pun"110+w Tam. J Spnt•0•11 i Rock"n. May INS Figure 5-7. Torus tank considerations - II. Membrane behavior. 5- is n . • HOOP

PERIODIC ^. CONTINUOUS ' } (TRUSS) ISHE LL )

10 OR 10 AN OD ^' OO i SAG t

vis MERIDIONAL SUSPENDERS • SURFACE PAD OR PILLOW

G;— ) ure 5-8. Torus tank considerations - III. Membrane support options.

• ANNULAR SEGMENTS 2 PC: WASHERS 2 PC: CYLINDERS l -, --I

/ I

I _^ 1 /^ OPTIONS

3 PC: CONESICYLINDER 4PC: OPTIONS / \^ .

I^ A i i1

DESIREMENTS REALITIES + / LIKELY RESUI T • Llyht, Tuuyh Matatat • Spu•tuta Imlwsa H1911 Strain V • WaWa12210 AL • Low Par Count —+ Nut Form/*iI*i-Anneal • Gore Conutuctwn • Mutunum Woid Lath —+ Lary Goons; AS • Stuck Must 11• Plot -WeWad S,wrr Formal Alumatwn • Truk is Radwily Rallied Seynwus Look GoW Figure 5-9. Torus tank considerations - N. Membrane forming and joining. 5-9 __ __..

• RETAIN "UNIFORM" -SHELL CONSTRUCTION

THICK STIFFENED SANDWICH SKIN SKIN

+ / OR OR

T"U • HEAVY • LIGHTER • LIGHT • STILLSOFT • FORMABLE? • STIFF • CHEAP • STIFFER • CRYO BOND • OA? • PROVIDE A SKELETON • LIFE? SKIN + RADIAL BULKHEADS + CIRCUMFERENTIAL RINGS CYLINDER + MEMBRANE FRAME WEB FRAME BEAM ^ r 00 ^r 000 OR I OR 000_ °o I o0 \ ti

• CONTAIN FLUID & GAS • SUPPORT SKIN • MAINTAIN BULKHEAD • TRANSFER FLUID INERTIA • MAINTAIN CROSS-SECTION SHAPE ORIENTATION (PREVENT TO BULKHEADS • TRANSFER FLUID INERTIA TO PIVOT) • PROVIDE SHELL INTERFACE SUPPORTS • REPLACE SHELL PORTION I?) • PROVIDE TRUSS INTERFACE Figure 5-10. Torus tank considerations - V. Structural concept opti-nns

Each of the eight shell segments consists of two 180° sections welded at the girth. Each segment also has weld lands for wall penetrations and for the disconnect panel support fittings. The largest penetration is the 24-inch access opening. This =access opening plus five additional penetration fittings are located in one shell segment. This arra_^_;ement allows ready access tn most of the system components without entering add icent compartments. T..:e access opening door also contains two outlets and sumping equipment for abort dump, fill, and drain. Each of the remaining seven shell segments has one hand hole for penetration which serves the acquisition system. In addition to the hand holes, two of the segments have weld lands for the disconnect panel support fittings.

After installation of the penetration fittings, the shell segments are fitted to a rings and held in place by fixtures. To allow for tolerances, the weld land widths on one segment are made oversize so that trimming can be performed. The complete assembly is then welded.

5-10 The tank is supported at the outside diameter with eight pairs of low conductive struts arranged in a "V" pattern and oriented so that the line of actions are tangent to the tank wall. Additional support is provided at the inside diameter with eight single struts which are also arranged so that the reaction loads are directly tangentially to the shell. To accommodate this support system, 16 fittings are welded to the rings.

5.2.1.2 Internal Structure

a. Pressurization Bubbler Manifold. The bubbler manifold is a 133-inch diameter tubular ring located forward of the acquisition system. The man±fold has a series of holes equally spaced over the entire length and is supported from the ring baffles with slip collars. Helium gas is supplied through a tank wall penetration fitting which is connected to the manifold with a tubular flex loop.

b. Ground Vent Duct. A tank wall penetration fitting located adjacent to the access opening serves as the ground vent outlet. This fitting is attached to an internal duct which follows the tank wall contour and routes forward to the ullage area. The forward end is radially restrained only with a collar fitting.

c. Fill and Drain Provisions. The tank is filled or draired through a flange penetration fitting located in the access door.

d. Abort Dump Provisions. The abort dump circuit is separate from the fill and drain; therefore, a flanged opening is provided at the center of the access door. A pull-through plate mounted on two cross webs is attached on the inside of the access door. A hole at the center of this pull-through plate allows the acquisition system capi llary device to be located inside the sump.

P_ Ring Baffle. A 109-inch diameter x 10-inch width ring baffle is provided at the tank center. The baffle is constructed in eight 45' sections and is attached to the radial baffles with angle clips. Each section is aluminum alloy sheet with stiffener beads.

f. Electrical Penetrations. Internal wiring is required for the zero-g vent, PU system, and instrumentation. The tank wall is equipped with a boss-type penetration fitting which interfaces with a removable receptacle. The receptacle is sealed to this boss with a metal radial seating seal using a backup flange.

g. Internal Bracketry. Internal bracketry is required for mounting the zero-g vent apparatus, ver" ",.,:t support, the PU assembly, and wiring harnesses. This bracketry consists of small z-rings, tee fittings, angles, and collars. Tne tank shell incorporates weld lands for each bracket which, in turn, are fillet welded to the tank wall. 5-. 1 h. Seals. All flange-type connections have metal radial seating primary seals and secondary metal O-ring seals. The cavities between these seals are vented overboard through a tubular manifold. A reduction in the number of mechanical connections is possible by replacing the flanged covers on the hand holes with welded caps. The use of welded caps, however, requires installing or replacing the acquisition capillary devices from the inside of the tank. For small tank wall penetrations requiring tube connections, in- duction brazing or orbit are welding will be used.

For connections inside the tank, AFRPL connectors using bobbin seals are employed for the helium feed tube running from the tank wall penetration fitting to the bubbler manifold. V-band connectors using metal seals are used for the larger lines. Some typical examples are the vent duct, and the removable section for the acquisition ring. Both the bubbler and acquisition manifold will have to be installed in sections. Where possible, the joints between sections will be orbit arc welded.

5.2.2 PROPELLANT ACQUISITION. Propellant acquisition systems were defined for use with a 1000-1b thrust LS2/LO2 engine. System selection was made after evaluation of feasibility and weight penalty for several propellant management techniques.

5.2.2.1 Propellant Acquisition Concepts. A propellant acquisition system operates by providing subcooled propellants to the main engine feed system pumps prior to each main engine start, and during main engine firing to prevent pump cavitation. This liquid can be supplied through either of two approaches. The most developed approach is to use settling motors for collecting propellant over the tank outlet prior to each engine start. The other approach is to use screens or capillary devices for maintaining liquid over the outlet during the entire vehicle mission. Capillary acquisition devices are divided into two general areas: (1) partial acquisition devices (start baskets) which do not conta:-t the liquid pool during a coast but contain enough liquid propellant to start the engine and setLle out the remaining propellant and (2) total acquisition devices which maintain contact with the liquid pool at all times.

The mission considered in the acquisition system design was an eight-burn transfer from LEO to GEO with rather long coast durations between burns. Two characteristics of this mission that differentiate the acquisition system design from previous studies are the low vehicle acceleration levels of 0.02 to 0.05 g and the possibility of a thrust vector misalignment caused by payload flexibility at final wain engine cutoff (MECO). The influence of these factors is dependent upon the particular propellant management technique under consideration.

5-12 f a. Total Capillary Acquisition. During the coast periods, propellant may migrate toward the forward end of the tank due to vehicle drag. In order for a total capillary acquisition device to maintain contact with the liquid pool at all times, it must extend throughout much of the tank. The resulting rather severe weight penalty of this system does not make it competitive with either of the two other acquisition methods for this mission. Therefore, it was eliminated from further consideration. b. Propulsive Settling. This approach uses a propulsive system to provide low acceleration for propellant collection following each coast period, prior to main engine operation. Consequently, the system weight penalty is proportional to the number of burns during a mission. Propulsive settling does not provide a means for acquiring the propellants for the case of a thrust misalignment during the final stages of draining. Additionally, because of the low acceleration environment of the OTV, rather severe suction dip of the propellants can occur during final draining resulting in vapor ingestion at relatively high propellant levels. Both may result in rather large propellant residuals at MECO. c. _Partial Capillary Acquisition. Partial acquisition devices, or start baskets, function by maintaining wetted screen barriers over the tank outlet. The start basket is sized to retain propellant in sufficient quantity during a vehicle coast to accommodate engine startup, propellant settling, and basket refill without supplying vapor to the engine feed line. Vapor will enter the start basket during a coast if heat input results in evaporation of some of the liquid. Vapor will also enter during engine startup as liquid is drawn from the basket before the propellant has been settled. Most of this vapor is then expelled from the start basket during the refilling operation under high acceleration. The amount of vapor which remains in the start basket depends on the screen surface retention pressure which must be overcome before vapor can pene- trate the wetted screen. The final vapor head trapped inside of the start basket is equal to: 4o- Hv = PL g DBP where: Hv = trapped vapor head a- = surface tension of the propellant P L = liquid propellant density g = acceleration in g's DBP = screen bubble point

L^ previous partial acquisition device studies (References 10 and 11, the trapped vapor head presented only a minor problem for the 0. 1 to 1 g acceleration range. This problem was alleviated through the use of a screened stand pipe placed at the top of the capillary device. The stand pipe is constructed of 5-13 screen having a higher bubble print diameter and, therefore, lower surface tension pressure than the remainder of the basket, thereby reducing the trapped vapor head. However, the 0.02 to 0.05 g acceleration range under consideration in this study yields a high trapped vapor head no matter which screen is selected for use with the acquisition device. This can be seen from Figures 5-11 and 5-12 which show trapped vapor head plotted versus acceleration for various screen meshes. With no practical means available to eliminate the trapped vapor head from the start basket, for low thrust vehicle missions, vapor penetration must be totally eliminated or minimized as much as possible. This makes the low thrust start basket design significantly different from previous start baskets designed for higher acceleration environments.

The weight penalty for a partial capillary acquisition device is relatively insensitive to the number of burns in a mission. It, therefore, becomes a more attractive system as the number of burns increase. Additionally, the system may be designed to acquire the propellants during a thrust misalignment, which a propulsive settling system is unable to do. Finally, suction dip of the propellants does not present a large problem since vapor ingestion will not occur until the surface retention pressure of the screen is exceeded, which will occur at low propellant levels.

5.2.2.2 System Selection. Because neither of the propellant management systems is without problems, it was decided to use the attributes of each system to design a combined propulsive settling-p.--- - dal capillary acquisition system. This system uses propulsive settling to initiate propellant acquisition formain engine start; however, by placing a system of screened channels at the bottom of the tank it is no longer necessary to have the propellants completely settled before main engine start can occur. As long as part of the screened device is in contact with the liquid pool, vapor will not be drawn into the acquisition device. The system also provides a means of acquiring propellants during thrust misalignment and minimizes the problems associated with suction dip of the propellants.

5.2.2.3 Torus Tank Acquisition Device. The acquisition de'. 1.;e for the toroidal LO2 tank consists of a ring manifold located at the bottom of the LO2 tank with eight equally spaced screened branch channels (Figure 5-13), which supplies a single outlet to the engine. This device is based on a concept tested by GDC. No matter what orientation the thrust offset, liquid is supplied and residuals are greatly reduced. (See Figure 5-14.) This design eliminates the complication of multiple outlets/sumps. Complete thermal isolation ensures liquid at all times. Design details are shown in Figure 5-15, and further described in the following paragraphs.

5-14 10.1

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do Af Jy -to Jrfp, I 4

4 'tp p9p 1pr rpp ,b/ ^ I.1 t J4 /•prB ,/•p0 ,►r /p

M 1 U 1 LI 1.0 11.2 ACCUEIAnOL" 171 -ZU Figure 5-11. LH2 trapped vapor head vs. accele:^ation for various screens.

10.1

W i

U e.11 at 1.0 l00 ACCELEAATM. e'1 9716.911 Figure 5-12. L02 trapped vapor head vs. acceleration for various screens. 5-15 JO MANIFOLD

SINGLE OUTLE TO ENGINE ANCH CHANNELS

TOROIDAL TANN

Figure 5-13. L02 tank propellant acquisition device.

QUID B

T 100 LIQUID

TORUS PROPELLANT ACQUISITION DEVICE MINIMIZES RESIDUALS WITH C.O. MISALIGNMENT. Figure 5-14. L02 acquisition with thrust misalignment. 5-16 c v a^ r: b ^^ 2 Oy^ 1 ^f a. LO A J1 M rW, V a 1 '^"er. t

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uI I ^^. ^ 1 a

5-17 a. Acquisition System. The acquisition system is a 109-inch diameter tubular ring manifold with nine tee fittings, eight capillary devices, and an outlet duct. The ring is located at the tank center and oriented so that eight tee fittings are opposite the tank access door and hand holes. The ring is supported with eight slip collars which are attached to the web sections of the ring baffles. A 30-inch removable section is provided at the access door to permit entry to the tank interior. Each of the eight tee fittings has a flanged side outlet which is attached to a capillary device. These capillary devices are 4-inch diameter x 2-inch long cylindrical screen ass,;mblies with a flange at cne end. The devices are installed from the outside of the tank through the hand holes. An alternative method would permit installation from inside the tank, which would delete the need for seals at the hand holes. When installed, the capillary devices protrude into the full depth of the hand holes.

The outlet for the manifold is the ninth tee fitting located adjacent to the access opening. This tee fitting is connected to a tank wall penetration fitting with a removable CRES duct equipped with flex joints for absorbing tolerances and flexures. b. LO2 Acquisition Device Sizing. The LO2 tank acquisition device waE sized for a worst case condition at the end of the last burn when a thrust misalignment of 10° might position the propellant in contact with only the branch channel located farthest from the tank outlet. The requirement to prevent vapor ingestion during main engine firing under this condition is that the system flow losses not exceed the surface retention pressure of any of the screened branch channels. That is

APB > A Pb + QPr + A Ph + APs (5-1)

AP.., screen retention pressure = '00- _ (5-2) Dgp

where = a dimensionless constant dependent on the individual screen and fluid being used 0- = the surface tension Dgp = the screen bubble point diameter OPb = the pressure loss due to bending from the branch channel

into the ring manifold = Kb PL V2 (5-3) 2 where Kb = a pressure loss coefficient taken from Reference 12 PL = the fluid density V = the branch channel fluid velocity OPr = the pressure loss due to flow in the ring V2 manifold = Kr PL' (5-4) 2 5-18 where Kr = a pressure loss coefficient accounting for both friction and bending in the manifold taken from Reference 13. APh = the hydrostatic pressure difference between branch P channels - gh (5-5) gc where g - acceleration h = the differential head supported by the screen gc a dimensional constant µLPLH2kL I, BVZ APs = the screen pressure loss = PH2 (5-6) where µL = the propellant viscosity µLH2 = the viscosity of 50 psi LH2 PLH2 = the density of 50 psi LH2 A&B = viscous and inertial constants determined in Reference 14.

The screens listed in Table 5-4 were initially considered for use in the acquisition system. They represent a broad sample of screen meshes available.

Table 5-4. Screen types.

Bubble Point Diameter, Screen ;Mesh Weave In. x 10-4

325 x 2300 Twilled Dutch 's. 94 200 x 1400 Twilled Dutch 5.51 500 x 500 Twilled Square 10.00 720 x 140 Reverse Dutch 10.24 165 x 800 Twilled Dutch 14.17 50 x 250 Plain Dutch 25.59 150 x 150 Square 40.55 24 x 110 Plain Dutch 61.02 60 x 60 Square 91.73 40 x 40 Square 150.00

The surface retention pressure (APO.) of each of the screens is plotted versus screen bubble point diameter in Figure 5-16. It is obvious from the figure that the two finest screen meshes, the 325 x 2300 and the 200 x 1400, are far superior to any of the other screens in terms of retention capability. This is not the only criterion for selection, however. The flow pressure loss across the screen is also an important consideration and it generally increases as the surface retention pressure increases. Figure 5-17 shows screen flow pressure loss plotted against surface retention pressure for two different screen flow areas. 5-19 r° ,..- ---

; I I I ! tl III fI U i II' I II lif ! j i 't I } ' I I I Ilii j ^t Ii, , l î { t ^(^ II 3X25 2300 1 Ifl t ' ^' 1` i ' j l^ i '{; t I ! ( ^ r ^ f ^ f ^Î ^ f I I I '(^ 11 j+ ? ^ f 41 ffii III ! ÎI^ I ^ ^ ^ , , , , i 7( '! I ^ I ; ^ t II l l i I i. i it i ' l ; I^^^ ^ ., Ij^ I ( , ' ,;, 1 ^ ^ ^ ^: I' u I 1 (, ;I l I I Il î' I' ,I t ^ , ^ {t' i i ` }+ f I ^ I I ,iî ' ' k I I ^ ^( ! ^` , 1 f ! 11 i ^ }II I^ (^ jjt , , `' }!" , I I ^ ^ II I it I! ,i li r.. t• I ; ( fit t i 1111. ! I , ! iiE^ 1 I ,. } ^jf , , I I, I Ro ,} ; , f ^ ^ ! ^ ; ; t I ^ ' ^^ }^f OC j ► (i ' i ( ^

j' '^ ^ 1 f j f f f ! f ffj1 frff j} I II; 4 200 X 1400 } l 1 11 i'I j I 1 III (} 5C ' 4 i ^ ! ! I! ! li I! I ^^ !^ ' ! r III t f } f l . f {f ^{^ I } , ! f f t I 1 t }I { i Î I },i ^! I I I lili i t I f, fl„ ^ i fli l i' I! lllf ijl ; 1 ; jl i' ; ,t;; ^ ` ! ^ j ! }jl I Iltl (( (! ÎI ; i ^j.i i 4C if f( j i H i r v f + } f t; 'f i f i fl l i 1 I11 t 1 0 111111 11 } 1. !(! ' rl } '1 } ^1,' I i^ ! I Ì , T ' a l ' li t, 4 T ;11 t!! I I !^! !I' r j^ 1j W 3C '720X 140- Il }r. i n rlr , . I ^,ff I I I I I Lit I l r f ^ III t W I. I i I }^ ,iI i I' i! 1 I' t ^ ^ } l f! I .^ 1IIt 'll ^^ ^!j ^ i' f It: Q 500.00 X 5 f ! f, I !I ,^!' t^+ ,1 I ^ i .; r I j!1 f'; 2 '.^ ^!^ 1 i!^' .1t I. ^' r l ^^ .0 I I .I I iii: W 2C , fI il!Illliî: I 11}( }lîl!^îl i^f îlll{Iji Cr f l 185X800 1 I I I' j, ^^ l t t li lift I11; . 1 11 i "; !! '?:? j I( }1! 1 I i, I1 ff f ' 1 I f i f li t I'I'. ;i! i 1} I,'}I ii 1' 6oX250 !:} j j 1 i il iiullll1 ' 1( F ill,;, ! ! +) ^tl,,^^,. Î 11 , i( it i r^r I,I ^ ;I f^ i i f l j, , 11 I I t ri I S0 X 150 I I I ( ! (t l j ' 1 I, ` 24X 110 f t t ' Ij!( j11; ^t ttl ii t iiil li I I` it +-r fr ^ i I H, 80X60 ..: .r.. it t,: 40X10 I r 1 I ^ ^ i ( it i }f ; irlt ^f ^ I ^ I I I ^^Î ^ I I II hill Î ^- CI 111 II!+ III ' liii I'll ^î I, 0 ZU 40 60 BO 1UU 120 140 SCREEN BUBBLE POINT, IN. X 10-4

Figure 5-16. Screen surface retention pressure in 13 psi L02.

Looking at the pressure difference between the surface tension pressure and the screen flow pressure loss as the reserve retention pressure, the fine mesh screens offer a significani advantage over the coarse screens in terms of '_s reserve retention, pressure at the hither screen flow area. It was for this reason that the candidate screen field was narrowed down to four screens. Those selected for further study were the 325 x 2300, 200 x 1400, 720 x 140, and 500 x 500 screens.

The ring manifold diameter was selected on the basis of pressure loss in the manifold. Figure 5-18 shows pressure drop in the ring manifold for flow from branch A (figure 5-15) to the engine outlet for different manifold diameters. As can be seen from the figure, the pressure drop is extremely sensitive to the manifold diameter. The 1.5-inch diameter manifold was selected in order to keep the manifold pressure loss approximately an order of magnitude less than the screen retc ition pressure. A larger oiameter manifold was not chosen as it would increase system weight and liquid residuals without any appreciable decrease in the manifold pressure loss.

5-20 80

N c 50

0 J W Q y 40 W Q

O J Z 30 W W Q V N 20

0 D ,0 0, SURFACE RETENTI0 i CAPABILITY, PSF Figure 5-17. LO2 screen flow pressure loss, in = 1.863 lb/sec.

Figure 5-19 shows the branch channel configuration. The side and bottom of the disk at the end of the channel are screened, with the screen mesh, disk height (h) and disk diameter (d) chosen to maximize screen retention capability while minimizing flow losses. The total system pressure loss between branch A and branch B was evaluated fo g the four screens and various combi- nations of d and h. Figure 5-20 shows data for the two most extreme screens in terms of system pressure drop as a percentage of surface retentior, pressure.

5-21 k K > 1W z MANIFOLD DIAMETER, IN.

Figure 5-18. L02 tank acquisition system ring manifold pressure loss, m = 1.363 lb/sec.

Based on these data, the 325 x 2300 screen was selected for use with the 4-inch by 1-inch disk in order to retain a factor of safety of about 2 against vapor penetration. The surface retentior pressure will be exceeded in the most severe case of flow through only one branch channel with this branch channel selection only when the side of the disk is completely uncovered, greatly reducing tank residuals.

5-22 ORIGINAL PAGE IS 11P' PI N W QUAL17T

Figure 5-19. Branch channel configuration.

100 . T 1 1000 16 ...... 111 r! 7!.. 1 lop :AOSFC t t' • I MII ......

120 x 140, 1, 0 1W OR

t 11 , I V t - ' , I I. #I... I. I I I... It,, 1. 11 . . I I . . I I . f . I 6 1 1 3 4 S 6 DISK DIAMETER, 101

Figure 5-20. Total L02 acquisition device system pressure drop for flow into one branch channel, in = 1.663 lb/sec,

5-23 c. Thermal Conditioning. A primary objective of the L02 partial acquisition device is to provide adequate thermal protection to assure no vapor entry. In general, concern with any screen device is that conduction heat penetration or ullage heating may result in vapor formation within the screen volume. Of equal concern is the possibility that vapor generated in the feedline, as a result of engine heat soak-back and radiation heat transfer, will expand into the screen device. Fortanately, the toroidal tank configuration and pressurization system combine to provide an acceptable passive thermal conditioning system. The beneficial aspects of these systems is described below.

1. Heating Effects. Acquisition system heating will be from conduction heat transfer through the L02 feed duct and heat exchange with the ullage. Ullage heating will have a minimal effect upon the screen device because near-thermal-equilibrium conditions should exist throughout the mission. This is due to the MLI system and pressurization system. The approximate 15 MLI layers seiected for the vehicle will reduce tank heating to the level where large temperature gradients will not be possible. Bubbling of helium through liquid during engine firing will tend to create equilibrium conditions. Thus, it is expected that the propellant contained within the screen device and the ullage will reside at about the same temperature. Heat conducted to the propellant through the L02 feed duct should be insignificant because of the duct length and tank wall properties. The feed duct from tank penetration to acquisition system is of sufficient length that most of the heat conducted along the duct will be convected to the tank propellants or ullage. Of greater importance is the fact that the major portion of heat conducted through the penetration will flow along the tank walls in preference to the feed duct. Tank wall thermal conductivity (aluminum) will be an order of magnitude greater than for the CRES feed duct. Furthermore, the tank cross-sectional area will be substantially greater for heat flow.

2. Feed Duct Vapor Generation. The feedline will contain pure liquid at the beginning of each zero-g coast period. Much of this liquid will be forced back to the propellant tank through the screen device as liquid boils during the extended zero-g coast period. Normally, it would be difficult to provide assurance that vapor would not be returned to the screen device. For this toroidal tank configuration, however, it appears that it will not reach the acquisition volume. It is estimated that as much as 18 inches of vapor may be contained within the tank volume, at equilibrium. Any vapor beyond this penetration length will condense. Condensation should occur because propellants will be subcooled (due to the tank helium partial pressure) relative to the vapor. Furthermore, heat conduction from the tank penetration along the tank wall will tend to cool the ullage and enhance condensation.

5-24 5.2.2.4 LH2 Tank Acquisition Device. The acquisition device for the LH2 tank is shown in Figure 5-21. It is made up of six equally spaced screened branch channels manifolded into a solid walled raised duct connected to the outlet duct. The material is 304L CRES. The branch channels consist of fine mesh screen on the top and bottom of the channel with solid sidewalls. The screen is backed by perforated plate with 50% open area to withstand loads encountered during the relatively high acceleration (3 g) Shuttle ascent phase.

TANK WALL

SCREEN--!

PERFORATED--II^

Figure 5-21. LH2 tank acquisition device.

The number of branch channels chosen was based on the competing factors of decreasing the number of channels to minimize weight and increasing the number of channels to minimize residuals. The channel dimensions were chosen in order to maximize the ratio of screen surface area to channel flow area while keeping sufficient channel height to minimize channel flow pressure losses. The outlet duct was raised above the tank wall by 12 inches in order to prevent vapor from traveling up the propellant duct from the engine during a coast period and working its way into the screened channel. If vapor did migrate into the branch channels it might be ingested by the engine at an inopportune time. Intercepting it in the raised duct would allow the vapor to flow into the engine during the coo' down sequence when it would not be detrimental to engine operation. The raised duct diameter is larger than the propellant duct diameter in order to contain sufficient liquid in the raised duct for thermal conditioning during coast phases.

5-25 a. LH2 Acquisition Device Sizing. As was the case for the L02 acquisition device, the LH2 acquisition device was sized for a worst case 10 0 thrust misalignment and flow through only one branch channel. (See Figure 5-22.) The requirement to prevent vapor ingestion is again that the system flow losses not exceed the screen surface retention pressure at any point in the system. For the LH2 acquisition device, this requirement is represented by:

APa- > A Pc + APb + APh + OPs + APcr (5-7)

where APO., APb, APh, APs are as previously defined APc - channel pressure loss = fsLs f s Lus (5-8) 2 gc where fs = the friction factor for the screened channel section determined from Reference 14. Ls - the length of the screened channel section. fus - the friction factor for the unscreened channel section, also from Reference 14. Lus - the length of the unscreened channel section. APcr = the Pressure loss due to the reduction of flow area in

the branch channel = Ker P LV2 (5-9) 2 ge where Kcr = a pressure loss coefficient for flow area reduction taken from Reference 15.

All LH2 properties used were for a vapor pressure of 14 psi. The LH2 flow rate of 0.311 lb/sec assumes a 1000-1b thrust engine with an Isp of 460 seconds and a 6 to 1 mixture ratio of L02 to LH2.

The combined pressure loss due to APc, APb, OP}1, and APcr was determined to be 3.925 psf. The major contribution to the pressure loss is due to APb, the pressure loss due to bending the flow. This pressure loss could be decreased by reducing the angle of the 'wo bends, but doing so would increase the system size and weight in order .0 maintain the same screen area without. any appreciable decrease in residuals.

Figure 5-23 gives surface retention pressure, AP.., versus screen bubble point diameter for screens in 14 psi LH2. The lower surface tension of LH2 compared to LO2 results in lower retention pressures for the screens in LH2 than are plotted in Figure 5-16 for the screens in LO2. In order for the screen surface retention pressure to be greater than the system pressure loss, only the 325 x 2300 and 200 x 1400 screens were considered for use since the 3.925 psf loss already calculated exceeds all the other screen retention pressures.

5-26

0 B

\LIOUIO

Figure 5-22. LH2 acquisition with thrust misalignment.

325 230011.X t 1 , , ;n , I ! , I f ;^ ' „ 1 f I , I 1 , , 1 10

,, I t ` ' fi t' t I t, t ! It ;;1 ! ! f 111 t} 1 , ,1 t I , I,. , ! I ! ,1, ;f li( ^, ^' i l l, !1 , ! } i11 ;n ,, r 1 lei } ^;.++.. ,+.. tI 1.1'} ! ',: 111 '! 1 ^: ^ t >-' „ ! , 1;1 I t. n `11 !1111' ;t , I. Î • I (; 1 , ; 1! I l 1 I+ I ' 1 1 I+ C + i ! ( I +k III I!. I +, LL I I 11 I II 111 11 f 1 ; ! ^! II 'I I I 1 ,• ,t '1'} .” 1 } + 1 _ }' t. , , 1 , , 1 I r. ! I I ! I j1 , 1 , 1 1 it l xoo laoo' 1, CL ' , ,! 11 ^ , . I 1,111 I(! ^^ , • i , I I ^ ^ I ;I ' ! !! u^ l 1 ! 1 k ut t }}t rt t, 11: , tt}^ } i. tt :.. ,1 r I I^ II I ;, 1 111 ;f li !, l ^! I;, It) I , 1 1 IT 1 1 I^ 11 ++}, I}1 , ', I!'; 1111, ir^ I I ,i 1 , !^, I f1! ^ t+ i +' f^ !I+ ^^^ T w a H I I:tit 11 11' II ^ i:1 +'t l t 1tt 1i 11 tj1 L L m r i 1, 1 .^f +. r , t "'' , Î 11'1 i, ^, I. I I , ! I I I s ,1^ 1 1 ! ^ , ' I , . I fi l ^'ti l ,J , ! i; 1 1 1, I^ i i ^ II II II ; +` 1 ! 1,1 1 ! ;,I^ I ;, ;1,1 I! f •;; ,1^; 1 ,1 ,! , ! ', ' , ^^ 1,. ^' , ^. ail , I I ^) I ,I I'( 1 }l. 1( j ! 1 u I' 1 1 ' }I (^, u^ iN' 1 I !11 ' 1: ^ ^f I +'+1 , l ^ 1 , I z ^ ji 1 I^ I i' i I ` ^ ' ll, ' I tl 'i ' 1 (( I^ I ' 1 ,! ! i I''' '' 7 ^ ^ I ,Î (( i^ , 1 11 ! i i, '1 (j ', I .1, , „1 0 ^! ' }?i .l. ! ; 1.+, , I , , , I ,; 1, ' 1 ,; t! + ^, i^ ^ I1 I' 1^>^ + i 1^^ I ! ;il t+,, Il, , 1'I! ; 1 P , , ! f ,1 i! } If!I 1 ' 1 I,Î i 1 , 1 11 i 1 I ► }11', 1 ' I z 4 + 't' i2o x 1aol ; ^^ .:! I ' +f II i^{ 1 f 1^,^1f i^t rt!t .'111 1' t ^ } 1 i t W j ! !1 I ! t' i'i i 1 11 fit'' Ì t 1 , (~ 500 X 500 I I ill ,, W t , , . l 11 ^:11 ! ! 111 il! ,t'i I' 1,1 ,1 !x, , II , Î 1 11 r + S ''I!f 1 ,^ ! !^^ '.;; 1 Ii;l 11 , +1 '. ttt1ll+ i t } : 1 t }1 }lit 11 .1. , i. { .t}i t.' l r, l z .I III W In n 185 X 6001' I !, ,111 1 1 , i, I' II r 1 1 !I 1 !, ! l} 111 y+ { 1 1• }I f W + +, I I f, E X1111'1" 1 t. t. I ' t I lli . t' tl Iti: 1 S I 1 i „ I i i .IU 1 I i;, li. 1 1;'.t I,'I ,!I 1 1 t i i' 'i ; 8 11. it i! 1 i I , .I} 5o x 280f I i I! t I I 150X 11601 ! ' 11 ` 24 X 110 80 X 80 I ^:'40 X 40

0 0 zu au su nu iw lzu IOU SCREEN BUBBLE POINT, IN. X 10-4

Figure 5-23. Screen surface retention pressure in 14 psi LH2-

e 5- 21 i

The total system pressure drop between points A and B in Figure 5-22, as a percentage of surface retention pressure, was evaluated for the two screen meshes for various screen flow areas. These data, presented in Figure 5-24, show that the 325 x 2300 screen outperforms the 200 x 1400 screen for all screen flow areas. Based on this, the 325 x 2300 screen was chosen for the LH2 acquisition device. The surface retention pressure will be exceeded for the case of flow through one branch channel with the 325 x 2300 screen when less than 18 square inches of screen flow area is in contact with the liquid pool.

b. Thermal Conditioning. The LH2 tank screen acquisition device will not benefit from the subcooled propellant environment available to the L02 device. The partial pressure of helium in the tank will not be adequate to expect vapor condensation. Although a thorough analysis was not conducted during the study, it is suspected that vapor penetration into the screen device may be prevented only if the LH2 feed duct is cooled with an active heat exchanger. The heat exchanger would be in thermal contact w i ''- a one-foot length of feed duct downstream of the tank outlet. Liquid hydrogen would be expa::ded through a throttling device to a low pressure and temperature. The vent fluid would intercept penetration heat leaks as well as condense vapor.

W — T - 1000 L6 iw • 460 SEC 90 _MR•6:1

V s a 60 -- - -' "'—'—""'1200 x 1400. P, - 7.61 od W W so

tW^

W 3 40 — —72S x 2700. P,10.88• „r 0 W 30

V 20

0 10 20 3U 111 DU DQ /U OU VU SCREEN FLOW AREA. SO., IN.

Figure 5-24. Total LH2 acquisition device system pressure drop for flow into one branch channel, m = 0.311 lb/sec.

5-28 R 5.'2.5 INSULATION. Analyses ..ire conducted to determine tank insulation system requirements for the low thrust OTV. The baseline mission selected for this analysis is described in Figure 5-25. A 40-hour LTPS/LSS checkout and erection period was assumed to occur before first main engine firing. Additional requirements aad inputs employed for this analysis are contained in Table 5-3 and Figures 5-26 through 5-28. These data were used for insulation system optimization.

Vehicle subsystem optimization will be based upon maximizing OTV payload capability. Consequently, it is necessary to include the interaction between subsystems. That is, an MLI system with very few radiation shields may be selected for minimum weight. But. the weight savings could be considerably less than the increased propellant boiloff. Thus, it may be necessary to compromise on a subsystem design point in order to optimize vehicle design.

1 11111 EUNKI AFnC H IVAN 10611M

• TRANSFER TIME 23 S Not t ^ • 1CCILINAT10R RANGE oil — 141 Tilt • MEIENI AT ENO OF APOGEE IIAN 21444 111

TRARt4FR AMT PARRYAT12t

ALT AT COAST COAST I ^ I \ IOAN :41TOF ArOCEE ^ Iii'_. ^\^ 1 I IIAN TIME WN COAST r1M( ALT 1 i^ ., — ^•^ X11 N1 INASI IN MI I NO ;HAS) IN W1 1 let 1 043 No 1 13 714 1 1 1 0.41 311 2 I S 1334 ^/ 1 I 3 141 443 3 11 tllo 1 1 ►EAIEEE IuANS / 1 ^ 4 0143 $II 4 23 3423 I 1.41 114 S 30 5405 1 110 1001 i 47 1211

2 ILIA 1311 T 44 11241 1 1 134 1 11971 1311.41A

Figure 5-25. ltission profile selected for subsystems definition.

a. RTLS Influence Upon OTV Design. Return-to-launch-site (RTLS) requirements are imposed upon the OTV. These requirements will influence allowable vehicle tank pressure, which will influence 1ILI selection. Propellant tank pressures of 19 psia (Lli_, tank) and .15 psis (LO` tank) will be required to expel propellants during RTLS abort. Thus, the 1%ILI system can be penalized only for propellant tank pressure increase in excess of the above stated values.

b. 1ILI System Optimization. The 1iLI system influence upon propellant tank design pressures. vapor residuals, and vent mass requirements has been assessed. This assessment was made for the main engine pressure requirements of Table 5-5, Item A, and for the pressurization system described by Item C.

I 5_219 t w

Table 5-5. Assumptions and requirements for insulation optimization study.

A. Main Engine Requirements Selected Tank Minimum Inlet Propellant NPSP, PSID AP, PSID Pressure, PSIA LO2 1.0 1.3 16 LH2 0.5 0.8 16

B. Mission Payload Partials

Item b PL/6 W, lbm/lbm Hardware weight (tanks, insulation, etc.) -1.0 Vent mass (T-0 to first main engine start) -0.6 Vapor Residuals @ final DIECO -1.6

C. Selected Pressurization Systems LH2 Tank 1. Engine start pressures provided by direct helium injection into ullage. 2 Main engine burn pressures maintained by autogenous pressurization.

LO,_) Tank - Engine start and main engine bu rn pressures provided by direct helium injection into liquid.

D. OTV Configuration Design* Propellant Tank Surface Area, ft2 Volume, ft3 Pressure, PSIA LH,) 599 1360 19 LO2 425.8 505.S 25

*Determined from RTLS Abort Dump Calculations.

5-30 .1

IN SYSTEM (HELIUM PURGED) I 1.11 UME INTERVAL FROM ASCENT a<

e e (LIS) ^i r 2.16 • 11.011 r O

NUMBER OF RADIATION SHIELDS

Figure 5-26. OTV tanks average boiloff flux during ascent.

NOTES:...._ ...... » _...... _ _^...... _._.._ ^' ^ 1. STEADY-STATE CONDITIONS ARE ASSUMED O.E 2. APPROXIMATELY 2 HOURS ARE REQUIRED TO ' ACHIEVE STEADY-STATE HEAT FLUX. .... :i...... 0.6

m ;, . - 3 0.4 X :2 » J I N.-

= 0.2 c_ -,• :MLI+PURGE BAG: 1 : -r M L I.:

2 —"t ^ _ 0 0 Zo au 60 80 NUMBER OF RADIATION SHIELDS

Figure 5-27. Space boiloff rates for insulated LH2 and L02 tanks.

R ..... 1, ...... _... i. NOTC3: 9. INSULATIGN SYSTEM IS MLI * PURGE BAG HELIUM PURGED fI-;2. {? : L02 TANK 0.8 LHTANK s 0 3 W 0.5 N N z 0 ONE BLANKET`TWO BLANKETS' 0.4 N Z 7`5 t _ -w -..

0.3 0 20 40 60 NUMBER OF RADIATION SH!'LOS

Figure 5-28. MLI insulation system weight. c. LTPS/LSS Checkout. The first trade performed was to determine whether the propellant tanks should be vented during the 40-hour checkout and erection period. The alternative to venting is to lock up the propellant tanks and absorb heat input, which increases tank pressures during this period and increases final vapor residuals. Heat 'Input to the propellant tanks was determined from Figures 5-26 and 5--27 and tank surface area given in Table 5-3, Item D. The resultant vent mast- is given in ,Figure 5-29 as a function of MLI layers. If, he Never, a no-vent option is used, an increased vapor residual mass will be experienced.

Payload penalties are given in Figure 5--s0 for the vent and no-vent options. The no-vent option is always preferred to the vent option. This result is due in part to the fact that tank pressure inc: ease during the 40-hour checkout will not exceed the tank pressures required for RTLS abort. Consequently, a tank weight increase is not needeC to accommodate the increased tank pressure. The no--vent option does not apply during Shuttle powered phase because rapid pressure increases resulting from hig:i heating rates occurring during this period preclude not venting. Thus, the no-vent option is applicable from SSME cutoff until OTV first engine burn.

5-32 L02 TANK. ?i: tt 00 LH2 TANK'

i... .: so 11..^VA OR RESIDUAL INCREASE OCCURS ONLY IF PROPELLANT TANK VENTING IS NOT PERMITTED. 2. NO LH2 TANK WEIGHT INCREASE CONSTANT PRESSURE r { =^ REQUIRED SINCE PRESSURE WILL NOT ;VENT MASS PRIOR TO RISE ABOVE ALLOWANCE VALUE OF 60 ;1ST ENGINE BURN 18 PSIA. m ^. •.^: ?::: ? :i :^..- :iii •^:?.:: 3. NO L02 T4NK WEIGHT INCREASE yJ N REQUIRED3 SINCE PRESSURE WILL NOT RISE ABOVE ALLOWABLE VALUE OF 34 ' . OR -:::: 25 PSIA. 40 . ; 4. VAPOR RESIDUAL INCREASE CAUSED BY TANK VAPOR PRESSURE INCREASE S: i PRIOR TO MES 1. 5. VENT MASS BASED UPON HEAT INPUT FROM SSME CUTOFF TO CUTOFF r 42 HOURS. 20 INCREASE IN FINAL iiii VAPOR RESIDUALS ...... : :::c: Of - 0 10 20 30 41 NUMBER IF MLI LAYERS

Figure 5-29. Vent option vs. no-vent option during LTPS / LSS checkout period.

NO-VENT CONDITION - PRESSURES DO NOT PAYLOAD EXCEED DESIGN VALUES PENALTY, (SET BY ABORT LB DUMP REOUIREMENISI 19 PSIA L142 25 PSIA — L02

NUMBER OF MLI LAYERS

L02 TANK ZERO "G" VENT AA, XER — --- L14 2 TANK

PAYLOAD PENALTY FOR 40 HR CHECKOUT AT LEO IS MINIMIZED WITH NO-VENT OPTION

Figure 5-30. 4 0-hr checkout (payload penalty).

5-33

5.2.3.1 Torus Tank Insulation. Figure 5-31 shows how MLI system weight and vapor residual mass will be influenced by number of MLI layers. Also included is the increased vapor residual mass caused by the propellant tank vapor pressure rise during the no-vent heating period from T-0 to MES1. Payload penalty is given in Figure 5-32 as a function of MLI layers, and is minimum for about 15 layers of MLI. It should be emphasized, however, that variations in payload penalty are sufficiently small that any MLI system selection between 10 and 30 layers should be acceptable.

400 W _ _. Design details of the insulation and purge system are presented in - TOTALZY Figures 5-33 through 5-35 (Layout 60) for the LO2 (torus) tank showing - - w NOTE: 1MG IS THE INCREASED general arrangements, construction 300 VAPOR RESIDUAL THAT RESULTS `w FROM THE NO-VENT OPTION details, and techniques for penetra- PRIOR TO MES 1. tions such as support struts and access holes. 2 CoJ The insulation is basically a s 20C lay-up of multi- layer radiation W 3 shields separated by dacron flocking spaced on 3/8 inch centers in a tri- - = — Obh4im— :VAPU angular pattern. This system (called r :RESIDUAL~ "Superfloc") was developed and 10C tested by GDC on an 87-inch Ellip- ...... soidal 2219 aluminum alloy tank. The first " Superfloc" system developed under 1968 IRAD funds used double aluminized mylar shields 0 0 10 20 30 with " Lexan" fasteners. A second NUMBER OF MLI LAYERS system tested used double goldized I Kapton" with P. P.O. (polyphenylene Figure 5-31. MLI influence upon LO2 tank o)ide) fasteners and an external vapor residuals. purge enclosure. This second effort

450 was accomplished under Contract ' NOfE: PAYLOAD PENALTY DATA NAS8-27419 for MSFC in 1975. A 71 C7'21[L-1t:: BASED UPON FIGURE 5 .31 AND third system was recently constructed {{ TABLE 54, ITEM C. _ -t 1 f. 3 using coated double aluminized r} .. "Kapton" under Contract NAS8-31778 d 445 ^^ .^. for I%ISFC. This third system, cm - c together with an acquisition device, # # is scheduled to be tested at MSFC.

440 !.. 0 10 20 30 Figure 5-32. Payload optimization NUMBER OF BALI LAYERS of L02 tank bILI system. 1 w 5-34

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5-37

The design shown features a tank-mounted purge gas distribution system. The complete tank, including the purge system, is enveloped with the multi-layer insulation (MLI). Although the configuration is different than those previously constructed, the same basic techniques are used in this application. The purge gas enclosure is not required for the oxidizer tank.

a. Purge System. Gaseous nitrogen is injected into the MLI lay-up ft approximately 36 points. (See Figure 5-36.) To accomplish this, two ring-type supply manifolds equipped with branch tubes are mounted on the tank. Each branch tube is equipped with a purge pin which engages with a hole in '.he MLI. Both the manifolds and the branch tubes are attached using fiberglass (F. G. ) bosses equipped with self-locking CRES inserts. The bosses are bonded to the tank wall and the tubes are attached with CRES clamps and screws. Upon pressurizing the manifolds, the nitrogen gas flows through the purge pins, through the MLI layers, and exhausts at the edges of the AILI blankets.

HELIUM PURGE ENCLOSURE TANK MOUNTED MEMBRANE BH'D MANIFOLD

MLI- TANK'`` 1 WALL HELIUM PURGE CAVITY I (302.0 FT3)

PURGE PINS ^ _ 1 ; -

(1 HELIUM SUPPLY t0 I. ^ 1 TANK MOUNTED MANIFOLD FOAM WAFFLE r STAND-OFF pt "'t I ` HELIUM VENT VALVE

GN2 SUPPLY TO TANK MOUNTED MANIFOLD PAYLOAD BAY FNVELOPE

PAYLOAD BAY GN2PUnGE

Figure 5-36. Purge system enclosure.

The purge manifolds, support bosses, clamps, and fittings protrude approximately 3/4 inch above the surface of the tank. MLI installed over an irregular surface of this type would cause local compression between layers and wo•-,d prevent a uniform fit. To avoid these surface discontinuities, foam blocks cast in a waffle pattern are located between the purge tubing and bonded to the tank. The foam blocks are not part of the insulation system and serve

5-38 only as a uniform base for mounting the MLI. The oxidizer tank purge system, which uses gaseous nitrogen without a purge enclosure, reduces vehicle weight, saves design and construction costs, and reduces ground-hold operation expenses. b. M_I. System. The MLI is applied to the tank in eight 45 0 preformed blankets. I flat pattern layout of a typical blanket is shown in Figure 5-34. The cross section is a series of 1/4 MLI core sheets sandwiched between two scrim- reinforced face sheets. Both face sheets and the core sheets are coated double aluminized "Kapton". The coating is an organic film which preserves the thermo characteristics of the aluminized surfaces. GDC developed coated aluminized "Kapton" under Contract NAS8-31778 for MSFC in 1978.

The purpose for the scrim-reinforced face sheets is to provide load carrying membranes and to improve general handling without damage to the 1/4 MIL core sheets. To prevent core sheet shifting relative to the face sheets, the blanket is interlayer spot bonded at 34 places. These spot bonds also serve as hard points for attaching fasteners. The typical cross section uses fiberglass washers between each layer for maintaining a uniform spacing. The fiberglass washers are coated on both sides with adhesive, stacked between the NILI layers, and bonded into a single hardpoint. "Velcro" hook sections on one end are engaged with the pile sections on the opposite end. The fasteners are engaged simply by applv,ng thumb pressure at the hard points.

Several cutouts in the blankets are required for tank support struts, plumbing penetrations, disconnect panel support fittings, and access holes. To prevent tearouts, the ccre sheets are locally reinforced with 1 MIL aluminized Kapton tap strips at the perimeter of each hole. Each blanket is purged at four points with holes cut in the core sheets and through one face sheet. c. ?enetrations. The access opening, the hand holes, and the support fittings are significant penetrations which must be insulated with the same system used on the tank walls. The entire opening is enveloped with a fiberglass "can-shaped" fairing which is attached to the door with fiberglass bosses which, In turn, are bonded to the access door. The fairing, which forms a plenum chamber around the access opening, is insulated with one wrap-around girth blanket and one circular cap blanket. The blankets are constructed similar to the large core sections and are attached to the fairing with "Velcro" fasteners. Purge pins bonded to the fairing engage with the blankets at three points. Overlap butt joints are used between blankets and held in position with 3/4-inch wide aluminized Kapton tape strips spaced on approximately 3-inch centers. A branch from the helium supply manifold injects helium gas into the plenum cha;nber, through the purge pins, through the DILI layers, and exhausts at the blanket edges between the tape strips. Should access to the tank be required, the edge tape strips are cut. the cap blanket removed and the fairing (with girth blanket) detached from the access door. 5-39 A similar arrangement is used for the hand holes, except the fairing is a shallow-pan-type configuration. The tank wall insulation terminates under the opening flange, and the fairing blankets intersect with mitered joints equipped with cover strips. Since the blanket sections are small, no purge pins are used. Purging is accomplished by diffusion from the helium atmo- sphere inside the purge enclosure.

The tank is attached to the vehicle structure with low conducw.c iiruts which are enveloped with MLI. A flared or half-boot area at one end of the strut insulation overlaps the tank fittings and is held in position with aluminized Kapton tape strips. The cavity around the fittings and the strut MLI is purged by helium gas flow from the blankets and by diffusion. The blanket stops at the penetration with a simple circular cutout. The insulation on the duct or tube forms an overlap butt joint with the tank blanket.

5.2.3.2 LH2 Tank Insulation. Figure 5-37 shows how MLI syc;.em weight and vapor residual mass will be influenced by the number of MLI layers. Also included in the figure is the increased vapor residual mass caused by propellant tank vapor pressure rise during the no-vent heating period from SSME cutoff to MES 1. Payload penalty is given in Figure 5-38 as a function of AILI layers. This system is optimized at about 17 layers. However, 30 layers of MLI was selected to satisfy prelaunch insulation system requirements.

The basic insulation and purge system arrangement is schematically shown in Figure 5-36. MLI materials, construction, lay-up, and attachment techniques are simi-

300 lar to those for the torus tank. The hydrogen tank requires a helium purge distribution system with a purge gas enclosure. The purge gas enclosure consists of a lightweight 250 membrane (scrim reinforced Kapton). The

mJ 450 = 100 c^ W O

RJ ^ X75 50 s 0J 1 sR

400 00 10 20 70 40 NUMBER OF MLI LAVERS NUMBER OF MLI LAVERS Figure 5-37. MLI influence upon LH2 Figure 5-38. Payload Optimization of tank vapor residuals. LH2 tank MLI system. 5-40

A structural shell of the vehicle is a part of the enclosure in order to save weight. The purge gas volume is 302 ft3. Helium is supplied to the purge pins through a tank- mounted manifold. During ascent, the purge system MLI is required to vent to an ambient pressure environment. To maximize thermal performance, the MLI inter- stitial pressure must vent down to 10- 4 torr within one hour after launch. An analysis is required to determine a venting approach that satisfies the above requirements.

5.2.4 PRESSURIZATION, TANK PRESSURE CONTROL, AND ABORT DUMP. The purpose of these systems is to satisfy main engine NPSP requirements for transient start and steady-state operation, and to maintain propellant tank pressure control throughout the mission. The pressurization system includes pressurant storage, lines, and valves needed for pressurant transfer to the propellant tanks, and software to monitor and command pressurization. The tank pressure control system includes hardware and software needed to maintain propellant tank pressures within prescribed limits during the mission, especially during coast periods between engine firings. It is expected that the same software System will monitor and command both pressurization and tank pressure control systems.

5.2.4.1 Pressurization System. The pressurization system selected for this vehicle will provide helium to each propellant tank to satisfy main engine NPSP requirements, autogenous pressurant for the LH2 tank during engine firing, and helium pressurant for the LO2 tank during engine firing. A pressurization system schematic is shown in Figure 5-39. Helium pressurization for main engine start was selected to simplify requirements for the main engine system design, i. e., eliminating the need for engine boot-strap capability.

GHZ BLE ENGINE

Figure 5-39. Propellant tan'. pressurization system.

5-41 a. Main Engine Requirements. The pressurization system will be designed to provide main engine NPSP throughout the low thrust OTV mission. This means that both engine NPSP and absolute pressure requirements must be satisfied. Tank pressurization AP can be determined from the following equation:

AP = OPNPSP + AP L + APACC + APS where AP - required tank pressure &PNpSp - engine NPSP AP F = feedllne frictional losses from the tank outlet to engine inlet APACC - Propellant acceleration losses during engine start pP transient (APACC - 0 for steady state engine firing) LAPS = pressurization system AP consisting of pressure sensing accuracy, deadband AP, etc.

APN PSp values of 0.5 psi (LH2 side) and 1.0 psi (L02 side) were selected for pressurization system sizing. The low propellant flowrates should result in low values for OPF and APACC. Consequently, it was decided to combine these terms with LAPS and use a representative value of 0.3 psid. Thus, required tank APS of 0. 3 psi (LH2 tank) and 1.3 psi (L02 tank) were selected for this study. (See Table 5-5. )

b. Software Controls. In addition to NPSP, main engines may also have a minimum operating pressure level. For this study, a minimum operating pressure of 16 psia was selected. Software must be available to satisfy both requi re- ments. Software logic will be relatively simple: pressurization valves will be commanded open to satisfy the absolute pressure requirement. At other times it mad• be necessary to reduce propellant tank pressures prior to pressurizing with helium, to avoid exceeding tank pressure allowables. The software control system will be capable of discriminating between pressurization and tank pressure control requirements.

c. LH2 Tank Pressurization. The hydrogen tank will be pressurized with helium from the ambient storage bottle prior to main engine start. Helium will be introduced into the tank through a diffuser to avoid the possibility of liquid spray created by a gas jet. Tank 6P will be maintained with helium until after main engine start, when autogenous pressurant is available from the main engine. Helium mass usages for the eight-burn mission were determined using empirical relations developed from Centaur vehicle flight experience. These mass quantities are summarized in Table 5-6.

.5-42

R Table 5-6. Helium pressurant mass quantities for low thrust OTV mission.

Press'n Engine Start Helium, lb Steady-State Helium, lb Period LH2 Tank L02 Tank L02 Tank

1 0.26 0.04 0.20 2 0.19 0.05 0.19 3 0.26 0.07 0.05 4 0.33 0.09 0.04 5 0.41 0.12 0.03 6 0.49 0.15 0.00 7 0.56 0.18 0.00 8 0.70 0.21 1.42 Total 3.2 0.91 1.91

Autogenous pressurant flow will commence shortly after steady-state engine conditions are established. A pressurant temperature of 350°R was selected for this study. As with helium, hydrogen will be introduced through the pressurization diffuser to minimize interaction with the propellant. During engine firing, heat exchange will occur between the ullage and tank walls, and between the ullage and liquid surface. As a result, both the tank walls and liquid surface will increase in temperature during engine firing. Warm tank wall temperatures present the potential for a sudden pressure rise after MECO because propellant will evaporate upon contact. Also, a stratified liquid surface could adversely affect main engine NPSP requirements during the final OTV engine firing. Fortunately, analysis has shown that these conditions are substantially less serious for an eight-burn mission that for a one- or two-burn mission. It was concluded that autogenous pressurization was suitable for the low thrust vehicle.

d. L02 Tank Pressurization. Helium pressurization of the L02 tank will be different from that of the LH2 tank; helium will be injected beneath the liquid surface rather than into the ullage. The advantage of this technique (which has been proven on the Centaur vehicle) is that less helium is required than for direct ullage injection. Reduced helium usage is due to the considerable oxygen evaporation into the helium bubbles that occurs during pressurization. In fact, the, evaporated oxygen is responsible for a major portion of tank AP.

5-43 f

Because of the small quantity of helium required, the liquid injection technique was also used during each engine burn. This method b.As the advantage of maintaining near-thermal equilibrium conditions during engine firing, so that pressure shifts following each MECO will be minimal. Another advantage for this tank pressurization method is than helium introduced throughout the mission will have an accumulative effect so that pressurant mass requirements can be reduced for subsequent pressurizations. For this study, the accumulative effects of helium for engine start pressurization were not considered so that conservatively high helium usages could be calculated. How- ever, the engine start helium was accounted for in calculating engine burn helium requirements. These quantities are given in Table 5-4. Note that a substantial quantity of helium is required during the final engine burn. This results from the need to maintain a minimum engine inlet pressure of 16 psia.

5.2.4.2 Tank Pressure Control. Propellant tank pressure control during each of the zero-g coast periods will be maintained with a thermod ynamic vent system (TVS). The primary components of the TVS are a heat exchanger and a mixing device (Figure 5-40). The vent aide, or cold side, of the heat exchanger will accept any combination of liquid and vapor, expand it to a reduced pressure and temperature, which allows for heat exchange with the tank side, or hot side, fluid. The heat exchanger is sized to guar- antee that vapor is always vented, even with pure liquid at the cold side inlet.

TANK WALL

1 5 TANK FLUID -^

i VENT FLUID THROTTLING I SHUTOFF VALVE HEATEXC14ANGER i VALVE

PERFORMANCE REQUIREMENTS • VENT PURE VAPOR REGARDLESS OF FLUID QUALITY AT SYSTEM INLET • MAINTAIN PROPELLANT TANK PRESSURE CONTROL OVER NARROW PRESSURE BAND

NT TEMPFRATURE SURF

ENTROPY

Figure 5-40. Zero-g thermodynamic vent system.

5-44 The mixing unit has a two-fold function: to pump the tank side fluid over the heat exchanger surfaces, and to mix the fluid with the propellant bulk. Forced convection flow over the heat exchanger will provide the heat transfer mechanism needed to completely vaporize the vent fluid. However, propellant tank pressure decay will not necessarily occur until the hot side vent fluid is mixed with and cools the bulk propellants. Because the TVS can satisfactorily operate with either liquid or vapor at the vent inlet, propellant tank pressure control can be maintained in a zero-g environment.

There may be instances where venting may be required in order to maintain propellant tank pressures within acceptable limits. This was not the case for the baseline eight-burn mission.

Mission propellant tank pressure profiles are given in Figure 5-41. The technique of bubbling helium through L02 causes vapor pressure decay during engine firing. Heat input during the coast periods will increase vapor pressure, but not enough to compensate for the decay during pressurization. Consequently, tank pressure will gradually decay throughout the mission as liquid vapor pressure decays.

Liquid hydrogen tank pressure during the multi-burn OTV mission will be influenced by the autogenous pressurant and propellant tank heating. At each MECO, the possibility of a pressure spike and pressure decay exists. As previously discussed, the pressure spike can occur when liquid quenches the warm tank walls. A pressure decay would occur once propellant mixing with the ullage established thermal equilibrium. These pressure spikes and decays are shown at each MECO condition.

L02TANK PROPELLANT TANK 14 PRESSURE, 1R PSIA ^; ^ ^11 lil h I ^i^ iI ► ! ^ ^.^ ^.

18 I i; I iii III I ^ ) ► I !►^ f ! ! +II I ! II I LHZ TANK I Ill^i I i ^ll 1 I^ Nl A i l } N I SSS I M 1f I ;,^ I 0 MAIN ENGINE CUTOFF IMECO) 14 ((i` I I ^ } I'I} j i' ! I I i I ii I l I ,I,, III ! I, ! I I . ^' li,,' TIME. NO SCALE • L02 TANK PRESSURIZED WITH HELIUM FOR ENGINE START AND ENGINE BURN • L14 2 TANK PRESSURIZED WITH HELIUM FOR ENGINE START; AUTOGENOUS PRESSURIZATION FOR ENGINE BURN • ENGINENPSPREOUIREMENT • 1,0PSI L02 • 0.5 PSI LN2 igure 5-41. Propellant tank pressure histories for eight-burn OTV mission. 5-45 5.2.4.3 Abort Lump. The return-to-launch-site (RTLS) requirements imposed upon the OTV are given in Table 5-7. These requirements may influence vehicle tank F^essure all.owables and pressurization system selection. The most demanding condition is for dumpirg propellants within 300 seconds. By using applicable data available from the Centaur-in-Shuttle Integration Study (Reference 16), helium pressurant requirements and propellant tank pressures during RTLS were determined for OTV. These data are summarized in the table and are based upon 5-inch I. D. propellant dump lines desigr.ad for Centaur. Note that propellant tank pressures of 19 psia (LH2 tank) and 25 psia (LO2 tank) will be required to expel propellants during RTLS abort.

Table 5-7. OT-V abort dump pressurization system requirements.

OTV Abort Dump Requirements 1. Dump cryogens in -- 300 seconds. 2. Repressurize propellant tanks following dump. 3. :Maintain MLI and engine purges from initiate RTLS to post-landing + 15 minutes.

Assumed Hardware 1. 5-inch I. D. dump lines 2. Ambient helium storage (btl. vol. = 3, 008 in 3 , btl. wt. _ 26.5 lb) - Initial conditions! P - 4000 psia, T = 540°R - Final conditions: P = 200 psia, T - 540°R 3. Cryo helium storage - Initial conditions: P = 3000 psia. T = 380R - Final conditions: P - 200 psia, T = 38°R

LH2 Tank L02 Tank

Initial liquid vapor pressure, psia 16 16

Dump pressure, psia 19 25 Helium for propellant dump, lb 22.3 13.3 Helium for rep.essurization, lb 3.3 ;Jot req'd

MLI purge = 9.95 lb Engine system purge - 3.64 lb Total helium required = 52.5 lb _ (13) ambient helium bottles or 6 f,l cryo-storage bot les

5-46 OTV pressurization requirements were determined for both ambient storage and cryo-storage of helium. Although 13 helium bottles are needed for ambient storage, they will be Shuttle-mounted and not affect OTV performance. However, this particular vehicle-payload combination is length-limited within the cargo bay, and may not accommodate a large number of helium storage bottles. By comparison, cryo-storage of helium will not impact OTV design because only about 6 0 of bottles are needed, which could be stored within the liquid hydrogen tank. Furthermore, this helitu^ «ill also be available for the mission if abort is not required. A payload penalty due to excessive pressurization system weight will be incurred for the mission, however, because the cryo-storage system is designed for abort dump helium requirements rather than mission helium requirements.

Design details of an alternate abort dump pressurization system were defined. Instead of using ambient helium bordes, a helium bottle is installed inside the liquid hydrogen tank with an external hot gas heat exchanger using either hydrazine mono- propellant or solid propellant as the heat source, exhausting overboard through the Orbiter skin. This results in a simplification of Shuttle accommodation systern design. (See Figure 5-42.)

To supply heated helium at 22 psi for 5 minutes, a total weight (hot gas generator + heat exchanger) of —90 lb has been determined. Hot gas flow rate is about 0.2 lb/sec.

Hamilton Standard has done similar work on a NASA/JSC CRAD for a freon/H2 system for the Orbiter. A tube-in-tube, counterflow heat exchanger with N2H4 hot gas generation and throttling (demonstrated) configuration was defined.

5.2.4.4 Propellant Tank Conditions./Mission/Trade. The optimization analyses previously discussed to select the MLl requirements were for a baseline 8-burn mission (1000 lbf engine). In this section the influence of number of burns and vehicle thrust level is examined. Four missions were used to determine final vapor residuals and maximum and minimum propellant tank pressure.

a. L02 Tank Conditions. Vapor residuals are given in Figure 5-43 as a function of initial liquid vapor pressure for each of the OTV missions identified in Table 5-6. It is evident that initial vapor pressure will have a substantially greater influence upon vapor residuals. The influence of engine thrust level or munber of engine burns is virtually insignificant. The same conclusions can be drawn from Figure 5-43, which gives maximum and minimum tank pressures plotted as a function of initial liquid vapor pressure and mission parameters.

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5-48 E b. LH2 Tank Conditions. Hydrogen vapor residual and maximum tank pressure data are shown in Figure 5-44. Results are substantially the same as for the L02 tank. That is, the mission parameters identified in Table 5-8 will have a minimal influence upon vapor residuals and mission pressures. Minimum hydrogen tank pressures at MECO were not identified because autogenous pressurization can provide any pressure level at little or no cost or impact upon system design.

+ I^I I ' li {I I,^I ,II ' ;^ ;,j^ II III' ' i I ilj I( i I I I i t ' 'I '' ill! I' I lilt I I , I ^ i Ij (i( il; { I I (; ' ( 111 21 ,^ ,I ' I VIII 1 I ,' 'll II I I 01 21 IN `' I 1 II III'II

; I; I; I,I ; II:''. I^I .t II I . ' I ''I I MAXIMUM ,', II 2 (I ^' III I I ^ '

PRESSURE I^ I' ! I;! I' Î(i I^ III (' I ? AT ME ' I'I Iii' ,I! II ' . a' : I I (r II II ^f III 'iiI ili i^ f ' i1 ' if" a X . il , ; II III i< .I I^l I(i! ' 1 ' ItpA N ^' .III I II IIII fl ;I I. ''-I I I1'^ I ÎI :^ _II tll I II ... ^.i ( ^ 1..! ^ I ` I ^ i t It IIr IIII,^: L FIII; i I ; III I , Ill VAPOR 1 Ifi^ L. ' I ; I I I I ;III PRESSURE, Ê IIIIj ^ 12.î( RESIDUAL, I _ ' ^ î I I I I 1^^'i PSIA I I I I :Ii I' '(1 ^ IIII IIII Lem I ; • IIII I'-I ^ ^ I I III , ,^.'I' ^!' ' i :'' 1 II I' I IC I {^ l i (^ ^ I i u 'r III i Î j ^;^Î N I 1 1 ' I I ,^ I I I , I i%''^li t 'ii I ( i 1 III I If I }I I 1 , I I ;III, I I„ ^''ii ;ir it ! ^ (I II ii( .p „I I ` II ;, j; ,, ^^ I II i i II MINIMUM ' I i^ r IIi^ I'' I'I' I I II(^ I 'I( II I lli i I ilia '' PRESSURE ' I I I ^ I' !'I I ' I (, AT MfCO Iu ,ilk 4 ^ I I I ';I II I I I I ,III I I'`^('^ 14 lil II I I 1 ,^ 1 i III 120 ''II , •I' ^ I, I`'(; 1i I { I '' it 1s IB 20 le 1S 20 INITIAL LIQUID VAPOR PRESSURE, PSIA INITIAL LIOU.0 VAPOR PRESSURE, PSIA

J 6- BURN, 1000 LBF T S-BURN, ]00 LBF C') S BURN, ItMI-OF (D 2-BURN. X000 LBF

LOZ TANK VAPOR RESIDUALS OR PRESSURES LITTLE AFFECTED BY MISSION — ENGINE THRUST OR NUMBER OF BURNS

Figure 5-43. OTV mission parameters influence upon L02 tank.

5-49

Table 5-8. OTV mission transfer orbit parameters.

BASELINE Mission 1 Mission 22 Mission (3 Minion 4 8-burn, 1 Ibf 8-burn, 3001bf 5-burn, 1000 Ibf 2-burn, 300^1bf Burn Coast Burn Coast Burn Coast Burn Coast Full and Time Time Time Time Time Time Time Time Coast No. (hours) (hours) (hours) (hours) (hours) (hours) (hours) (hours)

1 0.43 1.3 1.43 1.3 0.79 2.7 1.15 5.0 2 0.41 1.5 1.37 1.5 0.79 2.7 0.46 - 3 0.40 1.8 1.33 1.8 0.79 2.7 - - 4 0.43 2.3 0.43 2.3 0.79 2.7 -- - 5 0.46 3.0 1.53 3.0 1.3: - - - 6 0.50 4.7 1.67 4.7 - - - - 7 0.54 4.4 1.80 4.4 ------8 1.34 - 4.47 - - - - -

PRESSURE EXCURSIONS VAPOR RESIDUALS

t I ^ r• j l { l f { tj 1 :iI .. ' ! i 1111111 (f I^ i jii ^^ i l if { 1 ^; I :^ i' j 2 ^ If 1 11 101 ^! I^^ t ' ul (,; i I ! t^ t, I j ^^ !;t I i r^ l 111 ., ^.ir PRESSURE 11 'I 1 + i 1 ^ ,I VAPOR ll I \ wi ! I j^! I I \J RESIDUAL, 9( x"T"• 4 PSIA +^ IU T i i i l jjt ^,I ^ I I t Son X ^ ' { i I l 7 1' { O! 111 1 4 t,; tjl j!J1 I' 1 j1 I' II fl!I ' ^!II ^I^ I t r j I I^j I I' I }i^{!j^ II . I I+ ,I. 11!1 i;l ^... u fir! 11 {t '{I^ i^ ^}ii lilt 16 'II' 1^i ^1 II en I ,i i 1 II i . ^ 1^ 16 is INITIAL LIOUID VAPOR PRESSURE, PSIA INITIAI 11101110 VAPOR PRESSURF PSIA

TANK P RESSURE IS MAXIMUM VALUE TO OCCUR PRIOR TO HELIUM PRESSURIZATION

lO 9 BURN. I rA)O t OF V 4 BURN, 700 1 BF C3)) 6 BURN, II X10 L BF L 1 7-BURN. JOOO L BF

LH Z TANK VAPOR RESIDUALS OR PRESSURES LITTLE AFFECTIUO BY MISSIO►. - ENGINE THRUST OR NUMBER OF BURNS

Figure 5-44. OTV mission parameters influence upon LF2 tank.

1 ^ 15-50 5.2.5 ENGINE FEED DUCTS. Engine feed ducts for both fuel and oxidizer have axially restrained high flexure joints for absorbing the engine gimbal motions. Mate- rials for both the flex joints and the ducts are 304L CRES, A286 CRES, and 718 Inconel. The fuel duct routes from a flanged outlet in the fuel tank access door to the engine inlet. The duct has an offset configuration using three flex joints. Future detailed analysis may show the need for a fourth flex joint.

The oxidizer feed duct consists of two sections. One section is an "elephant trunk" inside the tank, running from the acquisition ring to the tank wall. The second section is located outside the tank and is routed from a tank wall fitting to the engine inlet using a wrap-around configuration. Two flex joints are located in the gimbal plane and a third flex joint is located at the engine inlet. A fourth low-flexure joint is located at the tank wall to compensate for small angular misalignments.

5.2.6 FILL AND DRAIN. The fill and drain ducts for both tanks are separate from the abort dump systems. For the fuel tank, the circuit starts at a flanged outlet located off center in the access door. A shutoff valve is located at the outlet and a duct is routed from this valve to the disconnect panel located at the aft end of the oxidizer tank. The duct is 304 L CRES and incorporates three axially restrained low-flexure joints. The duct terminates at the disconnect panel through a disconnect fitting. During the fill or drain modes, the fuel does not flow through the acquisition device. The flow path is through the annulus area between the access door ring and the acquisition device.

The fill and drain for the oxidizer tank is similar to that described for the fuel, excep* the duct is short since the tank outlet is near the disconnect panel. This short duct section consists of three flex joints, two vaned mitered elbows, and a disconnect fitting. A shutoff valve is also provided at the outlet. Additional details and final location for the fill and drain are presented in the oxidizer tank design.

5.2.7 PROPELLANT UTILIZATION. The propellant utilization device is basically two concentric tubes insulated from each other and equipped with sensors and wiring over the ercire length. A single straight section running from top to bottom inside the tank is desirable. For the fuel tank, however, a single straight section would have a span of 11.6 ft which presents support problems. The configuration has three straight sections positioned so that short support members from the tank wall can be used. A combination of support c,,Ilars, drag links and one tongue/clevis connection at the aft end is used. This support system allows dimensional changes between the tank and the PU assembly. The wiring is routed through the tank wall at the aft. bulkhead using a penetration fitting.

5-51 A straight bayonet-type PU assembly is used for the oxidizer tank. The assembly, complete with wiring, a blind flange, and one electrical penetration fitting is inserted through an access hole. The forward end mates with a conical fitting attached to the tank wall. The sliding fit between this conical fitting and the PU probe provides radial restraint only. The flanged connection at the aft end provides restraint in all directions. An alternate arrangement del &'. s the blind flange by installing the PU through a tank access hole. Provisions for ais are shown in the L02 tank design.

5.2.8 AUXILIARY PROPULSION/ATTITUDE CONTROL. Four hydrazine attitude control modules are located between the fuel and L02 tanks and are supported from the main body structure. Each module consists of a spherical propellant tank and four thrusters arranged in a cluster. The thrusters are supported from the bottle and the nozzles are scarfed so that the exhausts do not protrude beyond the body structure. Each bottle has an acquisition device and a pair of external support trunnions. Propellant requirements are shown in Tattle 5-9 for an 8-burn mission.

Table 5-9. Low thrust OTN' mission profile.

ACS Prop. Req. Time QV Approx. Event 1(hr:min) (ft/sec) Weight N2-N4 (lb)

Deploy OTV 10 58K 88 Coast #1 0:50 0.5 Burn #1 0:26 1.2 Coast #3 1:18 54K 0.8 Burn 43 0:25 1.1 Coast #3 1:30 0.9 Burn #3 0:24 1. 1 Coast #4 1:48 51K 1. 1 Burn #4 0:26 1.2 Coast #5 2:18 1.4 Burn #5 0:28 1.3 Coast #G 3:00 1.3 Burn #G 0:30 1.3 Coast #7 4:42 2.8 Burn #i 0:32 1.4 Coast #8 4:43 2.9 Burn #8 1:20 3.6 Deploy Payload 40 128 Coast #9 0:20 0.2 Burn #9 0:0S — Coast #10 12:00 7.3 Bu rn # 10 0:08 0. 1 Rend. & Dock 15 GK ;G Total Inipuise :z:75.000 lb-sec 323.S 5-52 0 5.2.9 AVIONICS/PDXVER. The low thrust expendable OTV has two operating characteristics which affect the avionics configuration. These characteristics are the longer transfer orbit duration and the reduced thrust acceleration and vibration as compared to a conventional thrust level expendable OTV.

The mission transfer orbit duration is increased from approximately 6 hours for Centaur, for example, to approximately 24 hours for a low-thrust OTV. The longer mission duration requires a means for maintaining the vehicle attitude reference accuracy and the guidance accuracy. An attitude update from a star tracker can satisfy the attitude reference accuracy. For the known mission guidance accuracy requirements, an attitude reference update by a star tracker will suffice in meeting the guidance accuracy requirements. For missions with unknown, more stringent guidance tolerances, a guidance update could be obtained from a TDRSS or GPS transponder yielding range and range rate updates.

A long transfer orbit results in Increased radiation exposure in the Van Allen Belt. For a single-trip, expendable OTV with low integrated total dosage, this is not a severe design requirement.

The electrical power system has to have a long-term energy source. As will be shown in the detail of the Electrical Power System discussion, a fuel cell system has the highest energy her emit mace ; ,c IL for missions In the 10- to 100-hour duration class. This includes dual redundancy for the fuel cell. A standard thrust OTN' also has some mission durations in this range, so the EPS for both low and standard thrust OTN' could use fuel cells, with higher reactant tankage requirements for the low thrust OTC'.

The extended duration low thrust OTV mission implies the need for higher avionics reliability. This, along with the increased ability to add redundancy without excessive weight penalty due to integrated circuit state-of-the-art advances, leads to an OTV avionics using advanced redundancy techniques and distributed processing.

The reduced thrust acceleration and vibration levels are of significance primarily for items that may be deployed during the transfer orbit such as solar arrays and antennas. The FPS will not have solar arrays for the standard missions, and deployable antennas are part of the payloads, so the OTV avionics will not be influenced by this design relaxation.

A representative OTV avionics configration is shown in Figure 5-45. Maintenance of high reliability for long durations can be met with advanced redundancy techniques using fault-tolerant computers with: power supply and logic cross strapping; memory bank sparing; improved self test, error detection, and soft failure detection algorithms; and increased component and sensor redundancy. Illustration of reliability improvement from these techniques is given in Figure 5-46. The advent of the micro-processo_- era and optical data transmission facilitates distributed computer architecture with

5-53 improved performance, simpler segregated software, lightweight componentry and harnessing, high EMI tolerance, and theoretical optical data rate limit. Many functions previously performed by analog circuitry or special purpose digital logic will now be accomplished by micro-processors. A Charged Coupled Device type of star tracker provides improved false star discrimination. if accurate na, igation update is needed, the transponder can be a TDRSS unit with ranging, or a special GPS unit can be added, assuming GPS operation is extended out to GEO. A weight and power summary of a typical low thrust expendable OTV avionics configuration is presented in Table 5-10.

In Figure 5-47, a comparison is made of three power source alternatives (i. e. , battery, solar array, and fuel cell), with regard to minimum weight for various power levels and mission durations. The fuel cells use propellant grade fuel with the tank weight assigned to propulsion. Fuel cells in a dual configuration have a margin over solar array source for 1 kW systems in the 10- to 100-hour mission duration. In addition, fuel cells do not have the solar orientation and thrust acceleration operating constraints of solar arrays.

DECRYPTOR/ ORBITER 1ENCRYPTORTRANSPONDER L----J

MSE ORBITER S/C MSE r BIU, BIU *, r- IU *1 REDUNDANT FIBEROPTIC BUS G,N,&C U PROCESSOR PIU BI

L BIU *j L BIU *J L BIU *J RIMU SIU SCU I!U

PIU PROCESSOR INTERFACE UNIT ACTUATORS AFT INSTRU. FLUID FWD INSTR. I I STAR * BIU BUS INTERFACE UNIT, DUAL SENSORS VALVES if PCU SENSORS TRACKER RIMU REDUNDANTIMU SIU SERVO INTERFACE UNIT SCU SEQUENCE CONTROL UNIT * µPROCESSOR ELEMENTS USED IN IMPLEMENTATION. IIU INSTRUMENT INTERFACE UNIT DATA BUS **ADVANCED INTERNAL REDUNDANCY TECHNIQUES. MSE MISSION SUPPORT EQUIPMENT 4" HARDWIRE NOTE: POWER AND SOME HARDWIRE SIGNALS NOT SHOWN.

Figure 5-45. Data buses and advanced redundancy techniques.

5-54 R n 0.95191 6 MOs. POWER 2GYROS SIMPLEX SUPPLY 2ACCEL COMPUTER IUS X- 57.3 X 1(1-6 (ORIGINALSYSTEM PAIR PLUS SPARE, POWER 2 GYROS PRESENT PAIR SYSTEM SUPPLY 2 ACCEL HAS EVEN LOWER SIMPLEX RELIABILITY) COMPUTER POWER 1GYR0 SUPPLY tACCEI

SIMPLEX COMPUTER ACTIVE STANDBY)

POWER 1 GYRO NOTE: THIS IS FOR SINGLE-AXIS GYROS. SUPPLY 1 AC '.EL USE ADAPTATION OF NASA STD IRW FOR DUAL-GIMBAL-AXIS GYROS. SEMI- 1 GYRO INDEPENDENT 1 ACCEL R - 0.99977 6 MOs. POWER 1GYRO ADVANCED OTV SUPPLY 1 ACCEL REDUNDANCY COMPUTER LOW THRUST EXPENDABLE SEMI- 1 GYRO (SIMPLEX INDEPENDENT 1 ACCEL X= 42.6 X ,0-6)

POWER j 1GYR0 SUPPLY IACCEL

SEMI- 1 GYRO INDEPENDENT 1ACCEL

Figure 5-46. Enhanced reliability for OTV.

5-55 I r 3.0

NOTES: 1. UNMANNED OTV EPS OPERATING REGIMES BASED ON WEIGHT ONLY. 2.5 NO CONSIDERATION OF RELIABILITY, OPERATING CONSTRAINTS, OR COST. 2. FUEL CELL IS MODIFIED ORBITER.

2.0

FUEL CELL HAS PERFORMANCE POWER MARGIN IN POWER LEVEL FOR LEVEL 1.5 MISSION DURATION IN 0.5-7.0 Ikwl DAY RANGE.

1.0

SOLAR ARRAY

0.5 \ BATTERY

0. 011\\\A , I 1 1 ! 1 1 1 1 1 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 1 MISSION DURATION (DAYS)

Figure 5-47. Power/mass efficiency operating regimes for various power sources.

5-56 IN Table 5-10. Low-thrust OTV avionics weight and power summary.

Weight Power (kg) (^^') ACS RINIU, 6 cha. A 33.0 215 CCD Star Tracker, Dual 8.4 52 Sun Sensor, Dual 2.2 5 RCS (covered by propulsion) 10 Bus Interface, Dual 4.5 10

48. 1 292 T, T. &C TDRSS Transponder, Dual 15.0 60 Premed Processor, Dual 3.6 10 Hemispherical Antennas 0.2 — RF Hardware 1.4 —

20.2 70 G, N, &C Processing G&N Processor Int. , Redundant 16.8 71 Processor Interface Unit Int. , Redund. 3.4 15

20.2 S6 Servo Electronics Servo Interface Int. , Redundant 9.1 20 Bus Interface, Dual 4.5 10

13.6 30 Sequence and Pyro Control Sequence Control Int. , Redundant 26.S 20 Pyrotechnic Control Int. , Redundant 14.5 20 Bus Interface, Dual 4.5 10

45.3 50 Instrumentation (Additional) Signal Conditioning 8.2 10 Transducers 9.1 — Bus Interface, Dual is 5 10

21.8 20 DOD Communications Decrvptor- Fneryptor 5.0 10

c5_ 57 Table 5-10. Low-thrust OTV avionics weight and power summary. (Concluded)

Weight Power

Propulsion Switching Modules (allocated SCU) 20 Vent Valves (allocated Fluidics) &.VV Hydrazine Heaters (allocated Propulsion) 50 270 Orbiter Interface Harnessing and Connectors 2.0 Bus Interface, Dual 4.5 10 (Fluidic and Propulsion Allocated to those Subsystems) 6.5 10

SUBTOTAL TO POWER SYSTE:II 181.2 838

Power Fuel Cell System, Dual S6.0 — (Including Transient Battery) 0.34 kg/hr — Power Control 6.8 10 Power Harnessing 2.5 — Bus Interface, Dual 4.5 10 99.8 20 + 0.34 kg/hr

VEHICLE TOTALS 281 853 +0.34 kg,"hr

5-58 0 5.3 INSTALLATION IN SHUTTLE. The forward end of the OTV is supported at Xo 1065.07 from a pair of trunnion fittings and a keel fitting which are integral with the OTV body structure. The aft end is supported from a rotary adapter which interfaces with a frame on the OTV and structurally connected with a system of latches.

The purpose for the adapter is to provide structural support, provisions for abort dump, systems interfaces with the Shuttle, and deployment of the OTV. The adapter is a cylindrical structure having one primary box frame, one interface ring (with latches), three stabilizing rings, one truss-type crossover structure, and a pair of disconnect panels. The adapter is also equipped with helium storage bottles for abort dump, astrionics equipment, and plumbing which runs from the disconnect panel to the Shuttle interfaces.

5.3.1 Adapter Structure. The primary box frame is equipped with two trunnion fittings which interface with Shuttle at Xo 1269.6. A keel fitting is also provided forward of this box frame. The truss-type crossover structure spans the adapter diameter and provides support for the engine, the plumbing rotary joints, astrionics equipment, and serves as a structural tie between the primary support point areas. The two disconnect panels (one for LH2 and one for LO2) are supported from the box ring with a system of struts. These panels are mounted on a rail or linkage system which permits retraction from the mating OTV panels.

5.3.2 Helium Storage. The ambient helium storage bottles provide gas to the OTV propellant tanks during abort dump. These high pressure bottles (4500 psi) are supported from the aft side of the primary box frame with yokes and struts. The bottles are interconnected with tubing through a control system. Three bottles are located between the disconnect panels at the bottom area and two are positioned above the pivot centerline near the plumbing rotary joints.

5.3.3 Plumbing. The plumbing system provides overboard circuits for ground vents, flight vents, fill and drain, helium fill, helium purge, and abort dump (Figure 5-48). There are six duct assemblies and two tubular lines. Each duct assembl% has two sections. The first section runs from the disconnect panel to a rotary joint and the second section routs from the rotary joint to the overboard interface. A typical duct is a 304L CRES welded assembly containing three axially restrained flex joints, a flange at one end, and a disconnect half at the opposite end. The flanges and disconnects are designed for dual seals with cavities vented overboard. The ducts for fill, drain, and ground vent are routed overboard through umbilical panels located between Z. 367.3/Zo 350.8 and Xo 1307/Xo 1278.0. The abort dump ducts have overboard locations at Zo 331.9 and Xo 1295.0. The flight vent and electrical cables attach to the Shuttle service panels located at the 1307 bulkhead.

Q zW w ^' w^ a^ Q = V z w

W^ F O x Y +^ I ^

^ I V] W ^ Z ., s\^ .4 i v^ I I ^ ¢b..U -- -- - . -- w^ a FWa F zz In ^W 1ìzl U W _a _ , ^, ^J^^,^-Qa c^ ova ^ c o e^ `,coolocc' F A o c :1.o, â aâ

i w • si ;v

E ° 5-60 To accommodate deployment, rotary joints are used. These joints are mounted on the adapter crossover structure and feature dual seals with vented cavities routed overboard. The static sections of these joints are attached to the 1307 bulkhead with struts for reacting the breakout forces. The two helium supply lines and the electrical cables use flex loops to absorb adapter rotation.

5.3.4 Deployment. The OTV and adapter can be rotated to 30 0 and 750 positions. (Figure 5-49). The 30 ° position is the minimum for clearing the OTV and payload package from V e Shuttle. This 30° position is also limited by the adapter box ring and the helium bottles. Some payloads are deployed before the OTC' is released from the adapter, therefore, the 75° position may be required, which, in turn, reduces the available payload length. For the 75 ° position, the pivot center is moved forward to Xo 1230.3 which is one of the standard locations. The relation between available payload length and deployment angle is, therefore, influenced by available support points.

During normal landings and abort operations, the center of gravity must fall within specified envelopes. Out-of-envelope conditions are permissible during launch and space flight. However, the conditions must be correctable before reentry or in the event of an abort on launch. The baseline OTV and payloads satisfy these requirements. (See Figure 5-50.)

A - D-

39.9FT AVAILABLE PAYLOAD LSS IRINGI X 1269.6 ° PIVOT t 1 `• 35.7 FT 1 AVAILABLE 1 PAYLOAD LSSIARMS) 300 1_ It

_ I ' X 1302 0 30" PIVOT X, 1302 0 2 1307.0

Z" 414 0 2,4000

()ItIG AL YA i;F; 1S OF pl w'.IR Qf 'ALITY

Figure 5-49. Deployment.

5-61

36K Wp 1AK PL 6K OTV 6K ASE

FULL DESIGN ASCENT CARGO LO21LH2 TORUS CONFIGURAT

CARGO WEIGHT DESIGN LANDED 11.000 61 CARGO EMPTY

0 L r r i ^ 0 10 20 30 t0 50 60 CARGO C.G. POSITION (fml

SNORT (TORUS) OTV SATISFIES SHUTTLE CONSTRAINTS

Figure 5-50. Shuttle c. g.

5-G'2 6 PROP*TL G /atTBSYSTEM TECHNOLOGY REQUIREMENTS

Systems identified in the baseline concept that requ:.—. technology development are the engine, the torus L02 tank, the propellant acquisition syetem for the torus tank, at.z! the insulatior. system for the torus tank. These are described !n standard SRT format aL the end of this section.

Estimated investmt: * needed is: Torus tank $3 to 5 M Propellant acquisition $1.0 Al i Fabrication and Test Insulation $0.5 M Low thrust engine $3 to 7 M Both n3w low thrust and punq)c-d idle TOTAL $7.5 to 13.5 M

6.1 'TORUS TANK

The United States has ao large torus tank space applications (France and Russia do: Ariane, Soyuz, Cosmos, Zond); however, Convair has built torus tanks to evaluate individual technologies (sloshing, residuals, and manufacturing, Figures 6-1 and 6-21 but a complete system development is required to determine the interaction of each technology, as fol ows:

a. Design - Optimization - Structural support - Manufacturing/producibility/forming/welding/assembly - Component assembly/installation/checkout

b. Sloshing - C ontrol/c. g. - Baffles

c. Propellant (1,02) Acquisition/Residuals Offset c. g. Thrust transient Low flow/high flow Zero-g

6-1

d. Pressurization - Main engine operation/multiple burns - Abort dump

e. Insulation - MLI applied to tows shape-'efficienc y vs. conventional tanks - Purge

E ^'^7 1 r .-A LIQUID 1 HLSILMAI S, I e WAILH

MAX)H DIAMETER BA IN ^ MATERIAL J,h IN YIEXIGLASS SLOSH TEST BAFFLES OUTFLOW TEST CONFIGURATION 18 SLUSH0 F R L UUE N0 16 (1 Y Hi

1 °r ANAL Y1ICAI 1 Hk OtANCIES I o O TILSI FRLUUkNlakb Ip \ 91 CANILU TUHUS I (NO HAI F LEST b \ :ND MOUE

o U

TLST UAIA 04 IST Mol-A

,1 .mil U 0 1 7 J 4 ., 6 7 d 9 10 0 1 8 17 16 :0 :4 :8 u UEV I HLL SURFACE STA NO IN LIOUID RESIDUALS SLOSH TEST"ESULTS

Figure G-1. Torus tank tests (Ref. AL.0 0-1325).

ti— rr

THIN WALL - ALUM I NUM CRYOGENIC TORUS TANKS HAVE BEEN BUILT BY GDC

Figure 6-2. :Aluminum torus tank.

6-' 6.2 LOW THRUST ENGINE

The United States has no high performance, low thrust hydrogen-oxygen space engine. Studies currently being sponsored by NASA/LeRC and NASA MSFC are investigating two concepts. Technology development/demonstration is recommended.

a. New Low Thrust Engine. This concept being studied by NASA/LeRC repre- sents an optimized engine, speeificaRy designed for low thrust application. Its advantages are its small size and low weight. Technology development concerns are cooling, very small pumps, and performance.

b. Pumped Idle/OTV Engine. The OTV engine concepts being studied by NASA MS FC Include the option o, operating at reduced thrust (10 c). Its advantage Is common development/utilization in the C)TV mission model for high thrust missions. Technology concerns are performance and stability at 10% thrust. The engine is also larger and heavier than a new low thrust engine.

6-4 I CC-INITION Of TECHNOLOGY RECU;AEMENT I. TECHNOLOGY RECUIFIF-MEVT (TITLE: Torus Tank Pigs 1 of 3

2 TECHNOLOGY CATEGORY: Low Thrust OTV grated System 3. CSJECTIVF-AOVANCc'.I1EN,T R ►;CUIREO: Inte

d. CUSRENT STATE CF APT: Convair has built torus tanks for evaluation. France and Russia have large torus space applications.

5. CESCRIPMON CF TEC7-+NCLCGY' Short OTVs are needed if accompanying large payloads are to be transported in the space shuttle orbiter. The shortest stage is achieved with a torus L02 tank surrounding the engine. Development of a full scale flight weight aluminum alloy toroidal tank for long term cryogenic storage for space application has never been done. Internal baffle systems, acquisition devices, supports arrangement (low conductive struts), and access openings are necessary complications.

3. PATiCNAL ANO ANALYSIS: The torus is a required configuration for aerospace vehicles for efficient use of available volume determined by the shuttle payload bay restraints. The torus yields the shortest OTV stage which in turn increases the available length for payloads.

6-5

R DERNMON OF TECHNOLOGY REQUIREMENT t. TECHNOLOGY REQUIREMENT ( TTLQ: Torus Tank Page 2 of 3

T. TECHNOLOGY OPTION& Various cross sectional shapes such as circular, kidney (tension membranes), and elliptical.

8. TECHNICAL PROBLEMS. Construction and assembly of gores (membrane forming and joining). Minimize residuals and sloshing. Numerous penetrations in shell necessary for acquisition system, outboard and inboard supports, abort dump, vent, etc.

9. PO i NT1AL ALTcRNATIVES. Nested tanks which result in a longer ,.tage and present structural penalties associated with the compressive bulkhe^..:.

10. PLANNED PROGRAMS CA UNPERTUREED TECHNOLOGY ADVANCEMENT:

t1. P"LATED TECHNOLOGY REQUIRE?,IENTS. J

6-6 CEFFNl^. CN CF —tC%-.;-tNCLCGY RECUIREMENT No.

I. _I7 NCLCGY ;_Cl;, gE.^,1c.NT (T1 i ^^: Torus Tank Pipe 2 of 3

12 T=C.%4NCLCGY ;E.UlREMF_NTS SC:iEOUL=: CAL_NOAR YEA R

^C4EOUL 17112A -9 9219319-4191 1

T_-CHNOLCGY I I I I :adapt current toroidal tank tech- ' I I nology to low thrust OTV. I 1. Develop mfg techniques, analyze structural and fluid characteristics. I ( I I ` 2. Fabricate test tank and per- I I I form evaluation tests li I

I F UNCING I—EVEL I I (Millions 19805) 0. 1.52.0 1 l I I i I I I

I i I^ I I' i I ^ I I i I I I 11 USAGE SCrE^u:.= cIAro LOGY G` n 0A—Id I T o7 A L I I I Y I I i I I I{ I{ `UMBER OF LAL-.Nc.MrI 1 1 1 1 1 1 1 1 1 1 1 1 1 I ► 1 1s. REF_RENCES

1_K LNEL OF STA—_ CF THE ART: S. Component or treadbcard- tested in relevant environment in lawratcrI I. Basic onenomena ctnerved and resorted S. Model tested in aircraft environment Z Thecry tormuiatec to describe 7. Model tested in 3oace environment -henomena & New capability derived !rem a muc't 3. TTtecry tested t:y ;nysical txceriment lesser cperatianal model or mathematical model 9. Reliat:ility uograding of an ccera- OPertinent `unctions or c:taractenstic ticnal model demonstrated. e.;.. material. 10. Lifetime exter.3icn of an cceraticnai Comoonent mcdel sea;-;a 6-7 OFPNI11ON OP TECHNOLOGY REQUiAEMENT

1. TECHNOLOGY REQUIREMENT MTLM: Acquisition Device for ftge 1 of _3 Torus Tank 2 TECHNOLOGY CATEGORY: Low Thrust OTV 3. 08JECTIVE1A0VANCEMENT REQUIRED: Cryogenic propellant management under low g

d. CURRENT STATE OF ART: Testing of cryogenic capillary devices performed under NAS3-20092 and planned under NASB-31778

S. DESCRIPTION OF TECHNOLOGY: Acquisition systems are required to assure liquid propellant flow to engine feed lines during engine operation. The system chosen is a tubular ring manifold at the bottom of the toroidal tank with screened branch channels positioned inside small sumps. A single outflow line supplies propellant to the engine. The screened branch channels prevent vapor from entering the device under most conditions due to surface tension effects. Vapor may not be allowed in the device due to large vapor head trapped under low accelerations.

8. RATIONALE ANO ANALYSIS: Because it can acquire propellants in any part of the tank, the acquisition system chosen will significantly reduce liquid residuals in the toroidal tank should a thrust misalignment occur during final draining. It will also reduce residuals due to rather severe propellant suction dip which occurs under low accelerations.

i i I

6-8 OERNMON OF TECHNOLOGY REQUIREMENT t. TECHNOLOW REQUIREMENT (TITLE):Acquisition Device for Torus Tankspsge 2 of 3

7. TECHNOLOGY OPTIONS. a. Develop full-scale acquisition system as presently conceptualized. b. Analyze a continuous screened ring with no branch channels.

S. TECHNICAL PROBLEMS. a. Suction dip for flow up into channel not well studied. b. Dlechanism of screen breakdown in a heating environment not well understood. c. Startup and shutdown transients flow tests required (NASS-31778 should provide some data).

PQ cNT1AL ALTERNATIVES. 3. ! Propulsive settling only with multiple outlets.

10. Pt„ANNEO PQCGRAMS OR UNPERTURBED TECHNOLOGY ADVANCEMENT.

11. RELATED TECHNCLCG', REQUIREMENTS: I I Ii

6-9 CE+NI!;!C,V CF :LH NCLCGY ;;_CUi REMENT Nc. t. (TI w;: Acquisition Device/Torus Rage 3 of 3

12 r=C),NCLCGY RIECUIREIMEINI i 5 SC:-iEZUL:-- CALSNCAA YSkR cCu U^ 1 ^a,^ -91 Sol Ills-.I 13 IS-L! $$1 961 V: 331 391 `?0 1 91 [9,219319-LI 9S `CHNCLCGY 1. Analyze fh.id flow mechanics G develop fabricate test c ompo - nents to demonstrate character- ( I istics. '. System test torus I UNCING L_"VEL I I I I (Millions 30 $) 0. 0.5 1 I I I I i I

i f 1 ^ I i 1 ^ { ( ^ ^ I ^ I ^ I 13. USAGE SCHECUI.E•

;zC*ANCLCCY `EM OA-1 E I I + ; CTAL

.Nti'MM OF LNLNC ZS ( I I + + + I ( I I I I I A. REFERENCES

S. LEVEL CF ST, aTS CF THE ART: r. Cemponeni or trwlteard•tested in relevant environment in %tcratery ), Basic onenomena coservect and reporter & Model tested in aircraft environment Z Thecry f ermulatee to Cescrit;e 1. Model tested in s pace snvircnment phenomena 8. New cacaciiity denvec !rem a muc.01 3. h ecry vested ry :rrysical excenment ►ewer operational mccel er mathematical rnccel 9. Feliat:illty upgrading of an ocera- s Pertinent ` unctions or viarac:enstic tienal model eemenstratec. e.S.. matenal, 10. Ufetime extension of an ccerat,cnal component mccel 163.141

6-10 I DEFINITION Of TECHNOLOGY REQUIREMENT i

1. TECHNOLOGY REQUIREMENT MTLA: Insulation/Torus Tank psae 1 a f 3

2 TECHNOLOGY CATEGORY: Low Thrust OTV

3. 041ECT1VEiA0'VANCEMEN7 REQUIRED: C mplete thermal insulation at minimal weight to assure liquid at all times and minimize propellant loss due to boiloff.

d. CURRENT STATE OF ART. MLI has been successfully_ applied to more convention al geometries. but not the torus.

S. CESCR1PT10N OF TECHNOLOGY: 'MiAtilayer insulation (bILI) consisting of radiation shields, separated by low conductive spacers are required for upper stage vehicles to provide theinnal protection of the cryogenic propellant tanks. The insulation system must be capable of providing adequate performance for the required mission cycles including ground operations •,to prevent moisture condensation). launch, be compatible with the torus tank design and the structural environment to which it is exposed, such as structural bending, flexing and bucki.ing resulting from thermal stresses and launch acceleration, vibration, and acoustic loading.

o. RATIONALE ANO ANALYSIS: Thera are INILI system designs available, together with analytical and experimental results. These systems include: (1) double goldized Kapton and double coated aluminized Kapton radiation shields, separated by Dacron tufts, utilizing purge bag and purge/repressurization systems; (2) double aluminized Mylar shields with silk and Dacron net spacers, utilizing a helium diffusion system (no purge bag). None of these systems has been applied to a torus tank. Specific requirements for the torus tank are: (1) to establish thermally efficient AILI design concept; (2) develop an efficient, lightweight MLI purge concept; and (3) establish thermal performance data.

6-11 DEFINITION OF TECHNOLOGY REQUIREMENT 1. TECHNOLOGY REQUIREMENT (ITnA: Insulation/Torus Tank Pigs 2 of 3

T. TECHNOLOGY OPTIONS: 1. Use purged, high performance, moisture resistant, coated I%ILI. 2. Consider MLI/foam composites to reduce prelaunch heat flux. 3. Consider MLI with one or two vapor cooled shields within a double wall tank to reduce boiloff such, that little or no venting will be required for short term missions. Disadvantages are high cost, high weight, and difficulties to construct lightweight double wall dewars.

S. TECHNICAL PROBLEMS. 1. Effective purging of MLI requires a purge enclosure and purging hardware which is expensive and heavy. 2. Radiation shield materials must have the ability to withstand exposure to a humid environment. 3. Complexity of MLI blanket fabrication due to the severe curvature of the torus tank. 9. POTENTIAL ALTERNATIVES. I 1. MLI with auawd double aluminized Kapton radiation shields, separated by Dacron tufts, utilizing a purge and repressurization system (with purge bag). 2. MLI with coated double aluminized Kapton radiation shields with Dacron net spacers (no purge bag).

1 0. PLANNED PROGRAMS OR UNPERTURBED TECHNOLOGY ADVANCEMENT.

it. RELATED TECHNOLOG`' REQUIREMENTS.

6-12 CE-F3NIMC.N OF i=C^4NCLCG4Y q 9-MUiAEME.xIT Va.

1. =Ci4NCLCGY T (717U. Insulation/Torus Tank Page 3 -W 3

12 TEtr+NOLCGY q ECUTAE'vtE.'vTS S>r)4Ei,UL=: CAL-2VOAA YEAR 7 SC:^CUL: I-VA -9 30 1 31 1 3:I U 1 I-► 1 931 961 r: IS 1 391 90 1 91 92 931 9r 91 TECHNOLCGY 1 I I . %1LI design L fabrication I I I i j 2. Thermal performance L it purge test I i I 3 System test torus M4 I l

i I I I I ?%- NClNG LEVEL I i i I I (Millions of 1980 dollars) 0 2 .11 I i !I

! I I I I! I

I i( f I I i I' I I I I 1 1 I I I I l i ^ l i l ^ ^ I II USAGE SCNcCU: 2 I

^C.-ih'o ^eeY e` D ^,^►;e I I ( x ( f t I i i I TOTAL `'UMBER OF LALN. CHES " I ) I I' I I 1 I nE= EREINCES

1S. LVEL OF STATE OF THE ART: S Component or tareactcard•tested in relevant environment in :ancratcr/ 1. Basic onencmena coserved ono reoom" 5. Model tested in aircraft environment Z Theory ferfrulatec to Cescrllre 7. Model tested in s pace environment ;henomena L New capability Cenvea trem a much 3. Theory tested Cy :rtysical !xcenment lesser operational mccei or mathematiui mcdei 9. Feliaoillty 4cgraciny of an ccera- d ?!rtinent 'unctions cr cnaractenstic tienal model cemcnstratec. e.;.. material, 10. Lifetime extension of an cceraticrai comoonent model ls::44 6-13 DEFINITION Of TECHNOLOGY REQUiAEMENT

1. TECHNOLOGY REQUIREMENT (TiTLA:_ LowTh,rustEngnePage 1 o} 3

26 TECHNOLOGY CATEGORY: Low Thrust OTV

3. 08JECTIVFIAOVANCEMENT REQUIRED: Low Thrust (i - 3 K) Hlgh Isp (450 sec) Engine

A. CURPENT STATE OF ART: No hi h performance low thrust engine exists

& OESCAIPTION OF TECHNOLOGY: Many of the large space systems which have been identified as candidates for transportation to GEO have minimum capability to withstand transfer acceleration loads. The development of a low thrust engine in the 1- to 3-K range is required. The engine must be of high performance and be capable of multiple burns.

6. RATIONALE ANO ANALYSIS. The low thrust engine options include a new low thrust design, or pumped idle mode of a larger (15K) OTV engine. While a pumped idle mode derivative engine could have lower (6) development costs, a new low thrust engine has advantages in weight, size, design simplicity, performance, and lower recurring cost.

E-14 DEFINMON OR TECHNOLOGY REQUIREMENT

1. TECHNOLOGY REQUIREMENT MTLE): Lc -v Thrust Engine Pip 2 Of 3

T. TECHNOLOGY OPTIONS: New low thrust engine Pumped idle GTV derivative engine

8. TECHNICAL PROSUMS. 1. Small engine technology (cooling, pumps) 2. Pumped idle technology (stabilih ) 3. Multiple starts (2-9 burns)

3. POTENTIAL ALi _RNATIVES. a. New low thrust engine. b. Low thrust (pumped idle made) of larger engine. I

10. PLANNE? P gCGRAMS CR UNPERTURBED r.CHNCLCGYi AOVANCEMEN T.

11. RELATED T—r-CHNCLCGY REQUIREMENTS.

f i

G-15

C1:= NI 7C.%4 Cr TI C4NCLCGY .; ECU;AV%AE34' Nci.

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12 TrC7-vNCLCWGY ;;E•:U1AF-M FN T S SC:IECUL.: CAL-ENCAA YSAP1 SCY0UL. I'"--M I -9 30 1 st 11: l !3 i !+I 151 36 1 37' 3i 1391 ?01 91 Ini gilul ill I

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Development l I i ( I 1

Pumped Idle Technolo gy for RLIO

FUNCING I-=VEL I i j I I I i I ( millions 1430 dollars) I , 1 I j I ^^ ^ I I I

Small Engine Technology I i 1 010

Pumped'.dle Technology for RL10 i i 10• 5 ^ ^ ^O 4 I I I I I^ j^ ^ i t ^^ ^ I t3. U5.>GE S^^c^UL= I TECHNCLCGY ` W 3a 7c l I I I . v ai. ► j .l- I ( I ' i I

`L-MBER QE LAL-NCHES I 1 I i I I I I ' a. FE=cAENCc3

1R L.eVEL OF STATE ".IF THE ART: S. Comcenent or breacecarc-testec in relevant environment in ;accratcrr I. Basic onencmena coserved ane rtcertec Q, Mccef tested in a+rc dtt environment Z. T rieery lermulatec to Cescrlte T. McCel tested in soace envircnment nencmena Pl. NOW Cacac +lity cenvec 'rcrr, a ruct1 3. , `*cry Astea try :rys+w sxCenment Ieeaer cCeratlena. ^nccel or .,,attlemaacal mccel 9. ReliaC+ lity uCgracing of an cceraticnal Mccel OPirtlnent `urivicns cr Viarac'trtatic _err+cnstratec. e.;.. material, 10. Ufstime exters+cn of an cceraticnal ccmpenent mceel M., is 6-16 7 COSTS AND SCHEDULE ESTIMATES

The purpose of this task was to estimate the costs of the low thrust OTV and prov program schedule for development, production, and operation of the vehicle.

7.1 COST METHODOLOGY

Tne low thrust OTV costs were determined using the General Dynamics Orbital T fer Vehicle (OTV) computerized cost model. The model was developed under NA MSFC Contract NAS8-33533, "Orbital Transfer Vehicle Concept Definition Study. model is presented in detail in the final report of that study. Its basis is primarily from the General Dynamics Space Tug Cost Model and the Space Systems Life Cycle Cost Model. The OTV cost model is computerized for quick turnaround with an output format which follows the work breakdown structure (WBS).

The organization of the cost model is centered around the WBS, with a cost estimating relationship for each WBS element. A summary WBS is shown in Figure 7-1. More detailed breakouts of the cost elements are provided in Figures 7-2, 7-3, 7-4, and 7-5. OTV hardware is presented at a Level 5 for maximum cost visibility of the vehicle system. The WBS numbering system is as follows:

WBS Number Series Cost Element

1000 DDT&E phase 2000 Production phase 3000 Operations phase 4000 First unit cost

The OTV cost model is a parametric costing technique. The parametric cost estimating relationships (CERs) for the individual cost elements were determined through analysis of historical cost data and cost study results. Cost estimating relationships for hardware provide cost as a function of the most influential parameter in the system, such as weight, thrust, or power. Other relationships, such as the cost for command and control in the operations phase, are price-quantity relationships. Given the size of ground crew, the model computes the total crew cost.

7-1 7.2 GROUNDRULES/ASSUMPTIONS

1. Costs are expressed in 1979 dollars.

2. Contractor fee is excluded.

3. All low thrust OTVs are expendable. Twenty-five units are produced and launched. Launch schedule based on NASA/MSFC Nominal OTV Mission Model, dated 2/29/80 (reference 1).

4. Ground Test. 1.75 equivalent hardware units were assumed for costing purposes. Ground testing includes: propulsion system and structural system testing including fatigue and vibration; Deployment Adapter Functional Test, Avionics Functional Tests, Thermal Vacuum Tests, and launch site verifica- tion. The production cost of one set of ASE was included in ground test hardware.

5. Flight Test. One proto-flight unit and one test flight were included for each development. Full price of a Shuttle launch ($25.4 million) plus operations was charged to the OTV program for each test flight. The production cost of one set of ASE was included for flight test.

G. Ground Support Equipment. Development cost plus the cost to produce three set,- was included.

7. Facilities and equipment are excluded from the cost estimates.

S. 0TV/Orbiter integration costs are excluded (Orbiter mods, KSC mods, overall program management and integration).

ORR1TAl.'rRANSFF.R MincLF^ TOTAL PROGRAM COST

DD1'&F, FIRST KNIT.% I OPFRA11ONS PIIASF^ ------L PRODUCTION PR;ISF,S FUGRT RARDWARF UO1'V F_ lXllT RARDwARF GR((I7ND RASFD (TFRAIIONS S1 AGF 1 SHI I I'1'Lr RASFD oPF'RAIl(INS STAGE 2. SPACr RASFD 0111-1MIONS SF. H LOCIS11CSSlIPPOPT 1N111AL 1`0K01,ING SYSTFM TFST ASF. VIF. FA(1 L111 F'S I'll Al Nl NG PROGRADI MANAOF'MFNT

Figure 7-1. Summary work breakdown structure.

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17-6

A 9. 40`(1 of Avionics system components are off-the-shelf or already developed. Fuel cell is a modiSed shuttle design. All other components are new development.

10. Shuttle user charges are $25.4 million per launch. It was assumed that no additional charges would be incurred for extravehicular activity or additional on-orbit stay time.

11. Production learning rates are shown below. Production was based on 25 units at a 2 per year rate.

LEARNING CURVES FOR OrTV

SUBSYSTEM V LC = n -+- 1

Structures 90 .343 Thermal Control S5 .766 Avionics 95 .926 Power Supply 95 .926 Main Engine 99 .985 Propulsion 90 .843 Solid Rockets 90 . S48 Attitude Control 90 .848 I, A, C /O, Test 35 .76666

12. Initial spares were based on 10 1;, of hudware development costs.

-i - 7.3 COSTS FOR THE LOW THRUST OTV

The finalized conceptual definition of the low thrust OTV is shown in Section 5. The design is totally current state of the art. Some features, however, will require a great deal more development than others. For example, the toroidal tank concept has been proven but still will require the development of manufacturing techniques and testing to assure its performance under all required conditions. The engine will be a new design using current technology. The avionics system will be a combination of both new and off-the-shelf hardware.

Costs for the optimized low thrust OTV program are provided in Tables 7-1 and 7-2. Costs are provided down to the hardware subsystem level. Appendix 13 contains the computer printout from which Tables 7-1 and 7-2 are derived. Costs in the Appen- dix are presented down to WBS Level 5, or one level lower. Development costs are $536.6 million. The largest contributors to this are the engine, avionics, and structures development. Production costs for 25 units average $12.82 million per unit. Avionics and propulsion constitute over half of that total. Operations cost, excluding the Shuttle user charge, is $5.13 million per launch. This cost assumes a launch rate of about 2 vehicles per year. Total program cost, for 25 units, is $1622.22 million, or $64.89 million per vehicle if DDT&E is amortized over 25 units.

Table 7-1. Low thrust CTV development and theoretical first unit cost's. (MILLIONS OF 1979 $)

COST ELEMENT DDT&E FIRST UNIT FLIGHT HARDWARE 292.6 13.19 STRUCTURES 54.6 1.41 THERMAL CONTROL 16.7 1.01 AVIONICS 53.0 4.29 POWER SUPPLY AND DISTRIBUTION 11.1 1.49 PROPULSION 146.2 2.63 ATTITUDE CONTROL 11.0 1.18 I. A, C/O, AND TEST - 1.18 SYSTEMS ENGINEERING AND INTEGRATION 46.6 INITIAL TOOLING 3.3 SYSTEM 'TEST 122.1 TEST HARDWARE 44.1 TEST OPERATIONS 78.0 AIRBORNE SUPPORT EQUIPMENT 14.3 3.92 GROUND SUPPORT EQUIPMENT 21.2 SOFTWARE 10.2 TRAINING 3.6 PROGRAM MANAGEMENT 22.7

TOTAL 536.6 7-8 i

Table 7-2. Production and operation costs. (Millions of 1979 $)

PRODUCTION OPERATIONS AVG. AVERAGE TOTAL FOR PRODUCTION TOTAL FOR COST PER TOTAL COST ELEMENT 25 UNITS UNIT COST 25 UNITS LAUNCH RECURRING

Flight Hardware 219.64 8.78 8.78 Structures 21.56 8.62 Thermal Control 11.88 .48 Avionics 84.57 3.38 Power Supply and Distribution 29.42 1.18 Propulsion 40.24 1.61 Attitude Control 18.11 .72 I, A, C/O & Test 13.87 .55 Spares 21.96 .88 0.88 Sustaining Tooling 17.57 .70 0.70 Engineering Support 43.93 1.76 I 1.76 Program Management 17.57 .70 0.70 Ground Based Operations .70 123.87 4.95 4.95 Prelaunch 2.00 Command and Control 53.90 Maintenance/Refurbishment 8.89 Replacement Training 3.50 in-Plant Engr's & Technical SPt 47.26 Program Management 8.32 Logistics Support 4.40 .18 0.18 Follow-On Spares 1.17 Propellants/Gases 3.24 Shuttle Based Operations 636.68 25.47 25.47 User Charge 635.00 Crew 1.68

TOTAL 320.67 12.82 764.95 30.60 43.42 (5.13 W ith nf NOTE: Production and Launch Costs Based on 25 Units. Shuttle) 7.4 PROGRAM SC HEDULE/FUNDING REQUIREMENTS

The program schedule was based on the NASA/MSFC nominal OTV mission model, reference I. Twenty -five missions were selected which required a low thrust OTV and whose payload requirements fell within the capacity of the low thrust vehicle. Three missions were selected: the Geostattonary Platform, including the demonstration article (13 units); DoD Class 2 (4 units); and Space Based Radar (8 units). The first flight is in 1987 with the launch of the Geostationary Platform Demonstration Article. The last launch is in 2000. The launches are shown in the tap portion of Figure 7-6.

The program schedule is alsc P-vided in Figure 7-6. The development program occurs o-ver a four-year time span (1983/1986). It includes hardware and software 7-9 development as well as extensive ground and flight system tests. The units are produced at a low rate, approximately 2 per year. A higher production rate would result in the requirement to store finished vehicles for several years before use. Thus, vehicles would have to be subjected to extensive checkout or refurbishment before being made operational (much like the Atlas F at Vandenberg, WTR). The low rate also keeps the production line open for future program extension or to provide spares. The 25 launches occur over a 14-year period during the operations phase. Launch rates vary from 1 to 3 per year with an average of about 2.

The schedule shown in Figure 7-6 also contains the costs for each activity. These were obtained from 'Tables 7-1 and 7-2. Using these costs and the timing of events, funding requirements were determined. These are shown in Figure 7-7 and are provided in detail in Table 7-3. Although the Shuttle user charges are part of the operations phase, they are provided separately due to their large contribution to cost. Shuttle charges comprise about 39"0 of the total program cost.

CY _ 00 81 8^ 83 80 85 88 87 88 89 90 91 92 93 9^ 95 98 97 98 99 00 FLT TEST LAUNCHES p p V 7 W17VW00 V 7 70'QW 0 0 1 1 2 3 1 3 2 1 1 2 3 1 1 DOM 111 DDT&E 1 1) HARDWARE/ASF10SE 1 21 SOFTWARE 1 102 1 3) TOOLING . SYSTEM TEST! 1 11 HARDWARE ( 5) TESTOPERATIONS ( 8) PROGRAM MGMT/ SE&UTRAINING

PRODUCT IO N ( 7) HARDWARE 1 219.8 2/YR --7 ( B) INITIAL SPARES ( 9) SUSTAINING TOOLING 1101 ENGR'G SUPPORT/ PROG MGMT

OPERATIONS 1111 PRELAUNCH/COMMAND . & CONTROL 112) ASE/GSE MAINTENANCE & REFURB 1131 INPLANT ENGRG/ 1 TRAINING/PROGRAM MGMT (111 SPARES/PROPELLANTS 11 & GASES 1151 SHUTTLE USER CHARGES

NOTE NUMBERS IN BARS ARE COSTSOF EACH ACTIVITY IN MILLIONS OF 1979 DOLLARS

Figure 7-6. Low ti.rust OTV program schedule.

7-10 I

ANNUAL FUNDING (MILLIONS 1975 S)

1980 1985 1990 1995 2000 YEAR

Figure 7-7. Low thrust OTV annual funding requirements.

Table 7-3. Low thrust OTV funding requirements.

SHUTTLE USER YEAR DDT&E PRODUCTION OPERATIONS CHARGE

1983 144.8 1984 214.7 1985 94.3 18.7 1986 92.3 29.7 1987 31.1 4.1 25.47 1988 20.1 5.9 25.47 1989 20.1 7.0 0 1990 20.1 8.4 50.94 1991 20.1 9.7 76.41 1992 20.1 10.9 101.38 1993 20.1 11.8 76.41 1994 20.1 12.2 50.94 1995 20.1 12.3 25.47 1996 20.1 11.8 25.47 1997 20.1 10.9 50.94 1998 20.1 9.5 76.41 1999 20.1 7.8 50.94 2000 6.0

TOTAL 536.6 320.7 128.3 636.68

7-11 The maximum funding requirement occurs in 1984 during the development phase. Extending the length of the phase would reduce this somewhat but would slip the production phase and the 1987 and 1988 launch objectives could not be met. The high peak-year funding in 1984 is a result of the time constraints imposed by the OTV mission model. Following the development phase, the production and operations costs are relatively constant (1987 through 2000) at about $30 million per year. The Shuttle user charge, because of its high unit cost, shows a very pronounced variation with the launch rate and causes an erratic expenditure profile.

The engine costs were based on CSR's (hrom the GDC Space Systems Life Cycle Cost :Model).

It was assumed a new 10001 thrust engine would be developed (Fined Nozzle, Con- ventional design).

7-12 8 REFERENCES

1. Saxton, D. R., Prelbninary OTV Mission Models, Revision 2, NASA/IMSFC Memorandum P504 (80-33), February 29, 1980.

2. DoD/STS On-Orbit-Assembly Concept Design Study, Final Report, Vol I, 11 Oct 1978, GDC Report CASD-AFS-77-005-39, General Dynamics Convair Division.

3. Low Thrust Chemical Obit to Orbit Propulsion System Propellant 1lanagement Study, Monthly Proj ess Reports, MCR-79-657, Martin Marietta Corp.

4. OTV Engine Study, Phase A Extension Task 9 Report, Alternate Low Thrust Capability, 19 Mar 1980, Rocketdyne Division Roc:cwell International Corp., RI/RD 80-123.

5. Obit Transfer Vehicl Engine Study, Task 5 Report, Low Thrust Capability FR-12898, 1 Feb 1980, Pratt & Whitney Aircraft Group.

6. Obit Transfer Vehicle Engine Study, Parametric Data Book, FR12253, 1 Oct 1979, Pratt & Whitney.

7. Oft-Orbit Assembly Concept Design Study - Trajectory Design Analysis for Low Thrust Liquid Stage, by L.A. Cowgill, General Dynamics Convair Report 697-0-78-002, January 1978.

8. Performance Advantages of Constant Acceleration vs. Constant Tbrust for Low Acceleration LEO to GEO Transfers, by L.;%. Cowgill, General Dynamics Convair Report 697-0-79-001,, January 1979.

9. Primary Propulsion LSSI Study, Monthly Progress Reports, MCR-79-653, Martin Marietta Corp.

10. Blatt, M. H., and Walter, M.D. , Cen'tr Propellant Acquisition System Study , NAS CR-134811, CASD-NAS-75-023, NAS3--17802, June 1975.

11. Blatt, M. H. , et al, Low Gravity Propellant Control Using Capillary Devices in Large Scale Cryogenic Vehicles, Phase 1 Final Report, GDC-DDB70-007, NAS8-21465, August 1970.

8-1 12. "Pressure Losses in Three-Leg Pipe Junctions: Dividing Flows", Engineering Sciences Data Item Number 73022, Cbtober 1973.

13. "Pressure Losses in Curved D: cts: Single Bends", Engineering Sciences Dat^ Item 1`11-niber 77008, May 1977.

14. Cady, E . C . , Study of Thermodynamic Vent and Screen Baffle Integration for Orbital Storage and Transfer of Liquid Hydrogen, NAS CR-134482, MDC G4798, NAS3-15846, August 1973.

15. "Pressure Losses in Flow Through a Sudden Contraction of Duct Area", Engi- neering Sciences Data Item Number 78007, December 1977.

16. Centaur-in-Shuttle Integration Study, GDC Report No. 670-0-80-83, 28 March 1980.

17. Low Thrust Vehicle Concept Study, LSST/Low Thrust Propulsion Technology Information Exchange Meeting, 20 May 1980, by Bill Ketchum, General Dynamics. Convair Division.

18. OTV Engine Study, Phase "A" extension 1, Alternate Low-Thrust Capability, 30 April, Report 32999E1-T2, Aerojet Liquid Rocket Co.

19. Low Thrust Chemical Rocket Study, Monthly Progress Reports, ASR-80, Rocketdyne Division, Rockwell International Corp.

20. Low Thrust Chemical Rocket Engine Study, Monthly Progress Reports, 21940, Aerojet Liquid Rocket Co.

21. Lcw Thrust Vehicle Concept Study, FYrst Progress Review, 30 Nov 1979, General Dynamics Convair Div.

22. Astrophysics Project Concept Summary: Gravity Wave Interferometer, Mar 78; NASA.

M. Low Thrlrl. Vehicle Concept Study, Progress Review, 24 Jan 1980.

24. Low Thrust Vehicle Concept Study, Final Review, 15 May 1980, General Dynamics, Convair Division.

25. Low Thrust Vehicle Concept Study, Data Dump, 4 April 1930, General Dynamics Convair Division.

8-2 26. CTV Configuration Development, CASD-ERR-77-050, Dec 1977, General Dynamics, Convair Division.

27. OIV Coufigvration Development, CASD-ERR-78-066, Dec 1978, General Dynamics, Convair Division.

28. Obit Transfer Vehicle, GDC-ERR-79-062, Dec 1979, General Dynamics, C onvair Division.

29. AIAA Paper #79-0880. Orbit Transfer Vehicle Propulsion for Transfer of Shuttle-Deployed Large Spacecraft to Geosynchronous Orbit, W. J. Ketchum, General Dynamics Convair Division.

8-3 APPENDIX 1 DEFINITIONS OF ALGORITHMS — SPACE BASED RADAR-A AND GEO PLATFORM ANALYSES

Computer Analytical Symbol Symbol Algorithms, Definition A a .5 2C, cos 150° + (C 1 2 + 2 rs )5j, ' Truss face width I AA a 16Mf ]1 /3' Membrane edge angle causing maximum membrane stress

AC Ac 21Tla 2./144, Circular area of antenna, ft

AL Of 51 2 /2, Term in K.

BO 00 02 - /4914 5 , Term in Ko, damped resonant frequency L 11 BI bl Time at start of thrust cut-off ramp B2 52 13Mf ] 1/3, Membrane edge angle affecting maximum truss bending loads

B5 b5 (1/$0)tan-1 ( (-G4G2 - G5 k/b)/(G1 G2 + G3 k/b) ll , Time at thrust start up at which K o occurs

B9 b9 (1/6 0)tan -1 I - [(s. /b) H5 + S H6/(b2 - bl ) + (k/db) H7 • k H2/d (b2 - bl) / [(s /b) Hl + S H2/ (b2 - bl) J • (k/db)H3 + k H4/d (b2 -b 1 )1Time after completion of JJ thurst cutoff at which K occurs 0 B b Time at completion of thrust start up ramp ) BC be - 1/a JUM -1 r- 6 J3 - (k/d) J4 ] / r JI + (k/d) J2 ] }, ( l Adjustment in b l to maximize K0 at thrust cutoff i 5 BE f g (Wz + Ww) k/WwNVz F , Term in Ko, undamped resonant frequency

Al-1 I 3

Definitions of Algorithms - Space Based Radar and Geo Platform Analyses (cont'd)

Computer Analytical Svmbol Symbol Algorithms, Definition BF + Ww)d/WZWw I' 5 , al rg (WZ Term in Ko, damping parameter

BP bl Thrust cutoff ramp interval

BT b2 b1 + b at completion of thrust cutoff ramp

C1 C1 3d1 + 4d2 , Height of one stowed truss bay

C2 C2 2(d1 + k1 d1 + d2), Length of one stowed truss bay

CD cd Critical damping fraction

Dl d1 Diameter of primary struts

D2 d,) Diameter of diagonal struts r .5 D d ed I k N'i't /g , Damping constant

DD Increment in dl

DIM Minimum dl

DT Q Attachment eccentricity- of membrane

D1 Nlwdmum d1

E E Youngs modulus truss construction material

EE a Eccentricity in longeron

EM E Membrane tensile modulus m F1 01 tan-1 (2a a o/(ao - a 2) ). Term in K0, phasing angle

FA F N 2 sin a sin 30°, A.dal force on truss due to a a a me mb ran

Al-2 Definitions of Algorithms - Space Based Radar and GEO Platform Analvset, (cont'd)

Computer Analytical Symbol Sol Algorithms, Definition

FB F Joint weight factor in truss FC N I sin 2 , End bending load on truss due to Nd F d a 0 FD Fd Nd R cos 15° cos 2 , kxi. load on truss due to 1,.d a 5 FI 0 tan-1 (60/ - a), Term in K ,, phasing angie

FS FS (Lc ; Lp)/Lq, Fraction of total cargo bay length utilized ab G1 G1 - a cos 0 + ^ o sin 0 + a e cos ( po b + 0)

-a0ebsiL(5ob +0) . G2 G2 d/b (1 - d Ni /4k) 5 ab G3 G3 - d o) cos 01 + sin 0 a cos (,Sob + 01) (a / 1 + (a 0) ab - e sin (a ob + 01) ab G4 G4 a sin o (.Q + 0) 0 +E Cos 0-ae sin ab - do a cos (^ 0 b + 0) C111 G5 G5 (a/ d 0) sin 01 + cos 0 1 - (a / 60) a sin(5 ob + 01) ab - e cos (^o b + 01 ), Factors in b5

G g Gravity acceleration

Hl H1 -a cos 0 + d o sin (^ + a b cos (? ob + ^) ab -hoe sin(6ob+0) ab,) 2 ab2 112 112 -ae cos (6ob2 + 0) + d o e sin (?eb,, + 0) abi ab1 + ae cos (sob 1 +0) - o c sin (Bob1 + 0)

Al - 3 R Definitions of Algorithms - Space Based Radar and GEC' Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition a - of cos + 0 sin + H3 H3 01 0 01 a e cos (? ob + 01) a -50 e sin (Sob+0i)

b H4 H4 -ae 2cos(^b2 +01)+doe 2 Sin (Sob2+01) abl Ub, + 01 + a e cos ($ 0 bi ) - 00 a sin (aobl + 01) a H5 H5 a sin 0 + $o cos 0 - of e sin ($0 b + 0) b - So e cos ($ ob + 0) of 2 ab2 H6 H6 a e sin (00 b2 + 0) + 5o a cos (50b2 + 0) ablab - a sin (^ obl + (^) - 50 e I cos (5obl + 0)

H7 H7 of sin 01 + $o cos 01 - a eb sin (50 +0 1)

-8 0 eb cos (So + 01) b2 b2 H8 H8 a e sin($ o b2 + 01 ) + 8 0 e cos(5 o b2 + 01) ab -Cie 1 sin(5 o bl + 0 bl 1) -5 060, bl cos ( 5 o + 01), Factors in b9

H h 2a cos 300 , Height of truss

II I1 n d1 3 t1 /8, Longeron moment of inertia

-I0 IS Isp 458 -5.49 x 10 (2500 - Tt)3.20 Specific impulse 1 b Jl Jl (t -b l -bl ) -0) oce cos (50 1 +5oe b sin (t - b l -bl ) -0) (90 - a cos (5 0 (t-b I ) - 0) - o sin($ 0 (t-bl ) - 0) 5 AI-4 Definitions of Algorithms - Space Basid Radar and GEO Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition b1 J2 J2 ace cos ($0 (t - bl -bl ) - 01) 1 b + 80 e sin($0 (t - bl -b l ) - 01)

- at cos 00 (t - bl ) - 01)

- (^o (t- b $0 sin l) -01) b 1 J3 J3 ace sin(80 (t - b i - bl ) - 0)

-804 bl Cos (g 0 (t- b l ) -0)

asin (t -hl) - (^o -0)

cos 1 + 80 (8o (t -b ) -0) ) -^1) J4J4 ae bl sun (eo (t -b 1 -bl b1 - 80 e cos (P0 (t -bl -b1 ) - 01)

- (x sin (8 0 (t -b 1 ) - 01)

+80 Cos (so (t-b1) Factors in be

K0 K0 K2, if K3 >K 2 then K o = K3 , Maximum thrust start-up or cutoff amplification factor

K1 kl Packaging factor for truss in C,)

K2 KZ 11/k) (d/b) [(; 2t sin(80 t -0) 1/2 (t-b) -e sin(8 (t -b) - 0) / (1-d 812/4k) 0

+ (k/b) Cb + (1/8 0)le at sin( o t-01) (t-b) -e sin( (t-b) - 01)1 J o Thurst startup amplification factor

A l-5 Definitions of Algorithms - Space Based Radar and GEO Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition Ott 4 Q(t -b) sin(5 K3 K3 ( - d/k) ( S /b 50) le sin ( Bo t - 0) 0(t -b) -0)1

+ `^/(b2 -b1) o Jl.e-oc(t -b2) sin(^o (t -b 2) - 0)

-e a(t -bl) sin(so (t -bl ) - 0) J

at + (k/d) (1 /b) C b + ( 1 1so) a sin (8o t -0l) cx(t - e -b) sin(% (t -b) -0l)

+ L 1/(b2 -b1) [-b 2 + bl + (1/80) a (t -b2) sin(8 0( t-b2) -01) J a(t -b -e 1 ) sin(Eo (t -bl ) - 0l ) l Thrust cutoff J amplification factor - after completed cutoff

K4 Programmed part of K3

K5 K5 Dummy variable for Ko, mwdmum value of K0 at thrust startup.

KG i*"6 Dummy variable for Ko, mwdmum value of K0 at thrust cutoff

K7 K7 Nlaadmuni amplification factor at thrust cutoff due to thrust startup transient at a(t-b) K8 K^ (VI) (8 /b E 0) [ e sin( q0 t -0) -e sin(^0(t-b) -0)

a(t-b2) • -e 80 81 [a/(b2 -bl) ^o ][ sin(ao(t-b 1 ) - 0) - of • (k/d) ( 1 /b) [b + (1/8 } a sin (5 0 t 0 -0l)

-e « (t-b sin(S (t-b) - 0

+ [ 1 /(b2 -bl),[-t Ybl + 2a/82

Al-ti

R Definitions of Algorithms - Space Based Radar and GEO Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition K8 K8 (continued) $O) 4,(t-bl) -(1/ sin( $0 (t-b1) - 00

Thrust cutoff amplification factor -after start of thrust cutoff

K k 1.5 n h2 d1 t1 E/1 a3 , Nominal truss cantilever spring rate

KP k Propellant weight loss per hour for power up to first P 12 hours

KQ k Propellant weight loss per hour for power after first q 12 hours

KS k Propellant weight loss per engine start s KT Propellant weight loss per hour due to leakage, boiloff k and attitude control

KU, KV K , K Factors in WA, K = 3, K = 5 for SBR-A and u v u v Ku = 4, Kv = 9 for SBR-R and GEO platform

LA .2 Truss length - deployed a LB Lb L - L , Available cargo bay length for structure and 4 P Its add-ons

LC L L l /.Z , Length of stowed truss structure c b a x LD Increment in LA

L'I iVlinimum LA

LP L (12 + .75 x 10 -4 Pw)12, OTV stage length P LQ L Available cargo bay length for total shuttle payload 9 and OTV

Al-7

Definitions of Algorithms - Space Based Radar and GE O Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition LX IX al.b/C2, Maximum LA due to Shuttle length fit

LZ Dummy variable for LA

M1 M1 4 4- Wt E aT + (1/2) Wa I a2 Tw Ko IFc a

+ F (p+ h) (, Truss bending moment - membrane separating from truss

M2 M2 I Wt w Ko +(1/2) Koat Tw -F + R a T a R a .289 W1 1 a3 T K Truss bending moment - 11 membrane acting on truss

`IF 1If a/te E m2 , Membrane factor W Tw Ko A MH Fraction of membrane occupied by holes M MU u Mass fraction

3IZ NI .85 + 5 x 10 -7 P , 'Mass fraction - dummy variable z w N N Number of burns

NA N W K T ,2 /2 sin a, Membrane edge load intensity - a 1 o w a drooped

ND WI Tw Ko $ a/2 sin 8 , membrane edge load intensity - N 2 non -drooped

M 7\i P p I or 2 = PJ whichever is larger h h Pi Axial load on longeron alternate P

PC Pcr n E II /a` or Per a 1 rdltl /(1 + c d 1 /2 r1 2) = PK, whichever is smaller

I Al-8 Definitions of Algorithms - Space Based Radar and GEO Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition P Critical longeron buckling load alternate Pcr

PL Pt ksN + k tq + Pp, total payload losses

PP P k t for t s 12 hrs. , k 12 + k (t -12) fort >12 hrs. P Pq q P q q q Propellant weight loss for onboard power generation

T'W P W u, Propellant weight w x R1 r1 .354d1, Longeron radius of gyration

RH p Structural material density

RINI Membrane material density PM RS r Usable shuttle cargo bay radius s

S1 Cr 1 Yield stress of truss construction material

SM a N A , Membrane stress M a e SP Qim Allowable membrane stress

TO minimum TA

Tl t1 Wall thickness of primary struts

T2 t2 Wall thickness of diagonal struts

T3, T4 T'w3, T'µ,4 Initial thrust-to-weight ranges in V1 , V2

T t Time

TA to Full membrane thickness

TD Increment in T1

'A1-9 R Definitions of Algorithms - Space Based Radar and GEO Platform Analyses (cont'd

Computer Analytical Symbol Symbol Algorithms, Definition TE to to L (. 907/NTh)' 5 -1 J, Effective membrane thickness i. e. , that reacts loads

TF Increment in TW

TG Increment in TA

TM Minimum fl

TN Minimum TW

TP TW Tw + (1 - u) (W s - W`v)^ /Ws ffw Thrust to weight ratio of OTV - initial (for constant thrust engine performance)

TQ t t + t , Mission time q r s TR t 25 for N = 9 10 for N =5 5 for N=2 Coast time, hours

TS t P I /T* 3600, Burn time, hours s w sp TT T W , OTV engine thrust T p s TY Maximum T1

11W T Thrust to weight ratio of OTV - final w TY Maximum TW

TZ Maximum TA

Al-10 Definitions of Algorithms - Space Based 2kadar and GEO Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition

V1, V2 V1, V2

If Tw >.001 and Tw<. 01 and N = 9 then V1 = 17300, V2 = 14500, Tw = . 001, T ' s 3 4 01 .001 .01 5 17800 15700 .001 .01 .001 .01 2 18400 17100 .001 .01 .01 .1 9 14500 14000 .01 .1 .01 .1 5 15700 14100 .01 . 1 .01 .1 2 17100 14400 .01 .1 .1 1.0 9 14000 13500 .1 1.0 .1 1.0 5 14100 13500 .1 1.0 .1 1.0 2 14400 13500 .1 1.0

Factors in V, velocity ranges for constant thrust curves-

V V 2)/(Tw - T'w3]) V - rl - V lw - T'3 Velocity requirements 4 LT 1J ,

for LEO-GEO transfer

W1 W1 to P m + WL , Membrane plus array unit area weight

WA W F P Fu n d1 t1 (. 357) + K ff d2t2(. 794 , Weight per unit length cf truss 1 ] WB Wb W la , Weight of one truss in SBR-A or entire ring in SBR-R

WC We W1 3 l a cos 30°, Weight of lens. (In GEO platform this is a part of the hub weight)

WH W .46 (Ww - Wh), Weight of hub

WL wL Unit area weight of lens array

WS W Stage total weight s WT Wt Weight on end of truss

WW %%1 G W + W } G W + W Total payload weight for SBR-A ho

Al-11 I

Definitions of Algorithms - Space Based Radar and GE O Platform Analyses (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition WX Ws - Wy, Loaded stage weight W

WY W (Ws/u) [exp(-V/l - + u -1 -P , OTV payload Y^Jsp capability - single Shuttle

IVY W^ IVs ju exp(V/Ispg)/ rexp(V/I spg) -17 -1 -P't, OTV payload capability - L,double shuttle

WZ (1 -u) (Ws -Ww), Weight of orbital transfer stage - empty W

ZH Hub Freight fraction

Al-12 APPENDLY 2 DEFINITIONS OF ALGORITHMS SPACE BASER RADAR-R ANALYSTS (New 'Terms and * Redefinitions of terms in Table 1)

Computer Analytical Symbol Symbol Algorithms, Definition Al Al frd1 t1 , Area of longitudinal member

A1%1 a I/Al .2a2, Hoop stress deformation factor

AT At 3 Al , Total truss cross-section area

BU 8 EI,/GAt lag , Transverse shear deform y dcn factor

DA d 157.38 - 9.23d , Storage envelope for lens - outside a diameter and insidei diameter for packaged ring truss

DB d Inside diarnevnr of stowed lens package

*FS FS (a + Lr + L)/L Fraction of cargo bay occupied by P q pavload and OTV

GE G Shear modulus 2 2 I I 2A.1 x -1.. (a cos 300 - x) Al , Moment of inertia of ring truss

KA kl 1 + a + g

KB k2 1 - a+ a

KC k^ k2/k1, Hoop stress correction factors

*LA a Lens radius u LF L Length of feed structure

' L:o1 Minimum LA

A2-1 Definitions of Algorithms - Annular Phased Array Analysis (cont'd)

Computer Analytical Symbol Symbol Algorithms, Definition 5 *LX $x 1(da- db2) 2a/3tL , ^ / 2, Maximum LA *Ml Ml W A [-1/4 + (1/Tr)(1 - k4/3 + 3 it/8) J, Maximum moment in ring truss

NO N 21T /a, Number of structural truss bays in ring o a circumference

*P P M1/(a cos 30° - x), Axial loan on lcageron

TL Thickness of lens t W W 2vla W1 TWK0, Peak loading intensity or ring truss due to lens

*WA F +rd t 1 (. 857) + 511d2t2(. 794) + W , Weight W 1 J per unit length of truss

WB Wh W 2Tr 2a, Weight of ring truss structure

ITI *WC We W1 a , Weight of lens WN Wn Weight per node

WP Wa 3 Wn/a, Distributed weight of nodes per unit length of truss

*WW + Wh, Total payload weight W We + W XB x (a/3) cos 30', Neutral axis location from face of truss structure.

A2-2 APPENDIX 3. OPTOTV COMPUTER SYMBOLS DEFMITIONS

SBR-A, SBR-R AND GEO PLATFORM ANALYSES

$YH0TV100A - ^,M!01S l% IN OPTOTYAOA

Ai - AREA OF LONGITUDINAL MEMBER A - TRUSS FACE WIDTH, LONGM LENGTH AR - MEMBRANE 'EDGE WANGLE CAUSING MAXIM.M MEMBRANE STRESS AC - CIRCULAR AREA OF ANTENNA IN SQUARE FEET AL - TERM IN K8 AM - NO STRESS DEFORMATION FACTOR HT - TOTAL TRUSS CROSS SECTION AREA B8 - TERM IN K8, DANPED RE-gftT FPEGUENCY Bi - TIME AT START OF THRUST CUT OFF RV BAY. - NEAREST INTEGER NIMU OF CYCLES BEYOND THRUST START IJP PEAK FOR Bi B2 - NEMBF1iNE EDGE ANGLE AFFECTING MFLXIMUM MISS BENDING. LW - 84 - DUMMY VARIABLE FOR Bi 85 - TIME OF THRUST START LAP PEW K8 % - TIME BEYOND START lV THRUST CUT OFF AT PEAK 0 B9 - TIME AFTER COMPLETION OF THRUST CUT OFF AT WHICH PEAK K9 OCCURS B - TIME AT CAMfPLETION OF THRUST START UP RAW BC - ADJU;TMENT IN B1 TO MAXIMIZE K8 AT THRUST CUT OFF BE, BF - TERMS IN K8 BP - THRUST CUT OFF P.W INTERVAL BT - TIME AT L IftETION TW Wn- OJT OFF PAW 9J - TRANSVERSE SHEAR DEFORMATION FACTOR Ci - HEIGHT OF STONED TRUSS BRY Q - LENGTH OF STOWED TRUSS BAY rD - CRITICAL [A&ING FRACTION Di - DIAEETEP OF P91ftY STRUTS 1dX - f4. W YARIA KE FOR INCREIANTING Di D2 - DIAMETER OF DIAGONAL STRUTS D - DAMPING 016TANT DR - STORAGE ENVELOPE FOR LENS - OUTSIDE DIAMETER AND INSIDE DIAMETER FOR PROKAGED RING TRUSS [8 - INSIDE DIAMETER OF STONED LENS PR ROE DC - DUMMY YARIABLE FOR DM C{► - INCREMENT 14 Di N% - IMMENT IN DV DM - MINIMUM Di i MINIMN_M Di.:t DT - ATTAIt_HMEHT EC ENTRICITY OF MEMBRANE N - 1NOPEMENI IN M1 (A - MAX I Ml JM Di K., - MAXIM :: Di% DY - Rft VARIABLE FOR DX E - YOUNG '5 MOKLUS TRUSS CCNSTPIXT MATERIAL EE - ECCENTRICITY IN LONGEPON E9 - MENBPANE TENSILE MOCIULUS Ft - TEPM IN Vh FA - AXIAL FORCE ON TRUSS DUE TO MEMBRANE FB -JOINT WEIGHT FACTOR IN TRUSS FC - END BENDING LOAD ON TRUSS 141E TO ND FD - AXIAL LOK) ON TRUSS DUE TO ND FI - TERM IN K8 FS - FPACi ION OF Uft BAY OCLI PIED BY PRYLOAD AND OTY Gi-(6 - FA CTOPS IN B5 G - (RWITATION WTANT

A3-1 GE - SHEAR WILUS H - HEIGHT OF TRUSS

Hi-W - FACTORS IN B9 I - WNT OF INERTIA OF RING TRUSS 11 - LON RON MOMENT OF INERTIA IS - SPECIFIC IMPULSE Ti-4 - FACTORS IN PC K9 - TMJST CUT OFF AMPLIFICATION FACTOR, LARGEST OF K2 OR 0 Kt - PACkAGING FACTOR FOP TRUSS K2 - 1HPlnT START LP fMLLIFICATION FACTOR Q - TlPl1ST CUT OFF AMPLIFICATION FACTOR AFTER COMPLETED CUT OFF K4 - FACTOP IN Q

KS - C+JfMY YRPIAELE FOR 0, MAXIMUM YALUE OF K8 AT THM--T (VT OFF K6 - DUMMY VARIABLE FOR K& MINIMUM YPLUE OF KQ AT THRUST CUT OFF K? - MAXIMUM AMPLFICATION FACTOR AT THRUST CUT OFF DUE TO THRLIST STA 0 - TF MT CUT OFF AMPLIFICATION FACTOR AFTER START OF CUT OFF K9 - THRUST START LP AMPLIFICATION FACTOR - FIRST PNIMUM AFTER B K - NOMINAL TRUSS CIl ILEYER SPRING RATE KAY KB. KC - HOOP STRESS CORRECTION FACTORS KE - MRiIM.M M POWER KP - PROPELLANT WEIGHT LOSS PER ON FOR UP TO FIRST 12 HOURS KO - PROPELLANT !EIGHT LOSS PER HCr.P FOR POIEP AFTER FIRST 12 HOURS ^S - PRALLANT HIEGHT LOSS PES ENGINE START KT - POVELLANT ►EIGHT L9SS PEP HOLR DLE TO LEAKAGE, BOIL OFF RNI) AL rl'14Y - MNSTANTS IN NA P.R Dl.arklD AND TPINGI.LAP M CROSS SECTIC KX,KY.K2 - DUMMY VARIABLES IN 0)MRITING 0., 6 h K7 LA - TRUSS LENGTH OR LENS PADIUS LB - AY IUKE CARGO BAY LENGTH FOR STONED TPUSS STP IXTLPE a ITS A LC - LENGTH 'IF STONED TPLMS STRLCTl_PE is - TRIMS LENGTH INCPMT - DEPLOYED LF - LENGTH 'f FEED STRUCTURE L_ - [^MMY VARIABLE FOR LZ

LM - MINIMUM TKISS LENGTH - DEPLLW LP - LENGTH OF OTY LU - CARGO BAY LENGTH AYAILABLE FOR IOTRL SHUTTLE PR&M LU - DUMMY YAPIABLE FOP. LX LX - MAXIMUM LA LY - (VW YARIABLE FOR LM

L - DUMMY VARIABLE FOR LA ALSO PEINITIR-IZED MINIMUM LA

Mi - TPL6S MING MOMENT (W OR MXIMUM MOMENT IN PING TPL6S 1APA) M2 - TPUSS BENDING MOMENT - MEMBRANE 9CTING ON TPUSS MF - ME10AFE FACTOP I^ - FRACTION OF MEMB7+ ►i1F !? UIEI! BY HOLES MU - MASS FRACTION MY - MINiM.M M1 M; - DUMMY YAPIABLE FOR MU N - N_MBEP OF BURNS

NA - MEMBRANE EDGE LOAD INTENSITY - DROOPED M - MEMBRANE EDG£ LOAD INTENSITY - '.M *OOPED Ni) - NUMBER OF STPUCTLPAL TPUSS BAYS IN PING CIMMEPENCE P - WAL LCKi CON LONGEPON K - CRITICAL LLVkP+'N 801ING LOAD

A,3-2 PI - ?.1416

P! - ALTERNATE P PK - ALTEPNATE PC PL - TOTAL PAYLOAD LOSSES PP - P40P91MT WEIGHT LOSS FOR POWER PW - PP.OPEIAX WEIGHT I.W. - DUMMY YRP.IABLE FOR Di(N) Pi - LONGEPON PADIUS OF GYRATION

PM - STRUCTURAL MATERIAL DENSITY RM - MEMBP.AME MATERIAL DENSITY PS - USSII LE 9MUTTLE CARGO BAY RADIUS Si - YIELD STRESS OF TRUSS CONSTRICTION MATERIAL till - MEMBRANE STRESS

SP - ALLOW MEMBRANE STPESS TO - MINIMUM TA Ti - WALL THICkTESS OF PRIMARY STRUTS Tip - DUMMY YARIABLE FOR INCREMENTING Ti T2 - WALL THICKNESS OF DIAGONAL STRUTS T'J4 - INITIAL THRUST-TO-EIGHT PfHfMES IN Vi AND V2 RESPECTIVELY T - TIME; IN DYNAMICS ANRLY515 AND IN q-f PPINT OUT, TIME AFTER START OF THRUST CUT OFF AT WHICH K8 IS MAXIMUM

TA - FJJ L MEMBRANE TH IOli S TB - DUMMY VARIABLE FOR TN TC - DUMMY VARIABLE FOR TM TD - INCREMENT IN Ti TD - INCREMENT IN Ti:,.' TE - EFFECTIVE MEMSft THICKTESS I. E THAT R@XTS LW-.) TF - INCREMENT IN TN TFG - INCREMENT IN TNT.

TG - INCPEMENT IN TA TL - LENS THICKNESS TM - MINIMUM Ti TM% - MINIMUM T1;: TN - MINIMUM TN TK: - MINIMUM TIC!

TP - 1HRUST TO HEIGHT PATIO 'IF OTV-INITtgL T4 - MISSION TIME IN Hl PS TP - 0W TIME, HOURS TS - Ek" TIME HOURS TT - OTY ENGINE THRUST TU - DUMMY YARIABLE FOP TX TW - THRUST TO WEIGHT PATIO OF OTY-FW THH/. - D10Y YAPIABLE FOP INCREMENTING TN TX - MAXIMUM Ti TX': - MAXIMM.M TSX TY - MAXIMUM TW TY'i - MAXIMUM TW

TZ - MAXIMUM TA Yi,Y2 - FACTORS IN Y. VELOCITY PRNGES Y - VELOCITY REOUIPEMENTS'FOP LEO-GE0 TPAMtFEP. 14 - MEMMNE PLUS WRAY UNIT AREA WEIGHT W - PEAK LOSING INTENSITY ON PING iMS DUE TO LENS * - WEIGHT PEP OMIT LENGTH Cf TPUwS Q - WEIGHT OF CM TFI6S IN -8P Cf ENTIPE PING IN APA

WC - WEIGHT OF LEN:.

A3-3 I4H - WEIGHT OF WM.IB WL - ARRAY UNIT AREA WEIGHT WN - HEIGHT PER NODE HP - DISTRIBUTED WEIGHT OF NODES PEP UNIT LENGTH OF TRUSS WS - TOTAL PAYLOAD PLUS LOADED STAGE WEIGHT WT - WEIGHT ON END OF ONE TRUSS

WU - INITIAL INPUT WORLD WEIGHT FOR GEO PLATFORM WORD WY - R MY VARIABLE FOR WU WIN - TOTAL PAYLOAD WEIGHT 14X - LOADED STAGE WEIGHT WY - OTV PAYLOAD CAPABILITY W Z - WEIGHT OF ORBITAL TPAEEM, STAGE - EMPTY t - NEUTRAL AXIS LOCATION FROM FACE OF TRUSS STP47URE IN APR ZH - HUB WEIGHT FPACTION OF WE, WT AND REFLECTOR STRICTURE WEIGHT ZI - AVERAGE BOOM WEIGHTS IN GEO PLATFORM ANALYSIS ZO - SBR *2 OPTION IDENTIFIER 11 - DESIGNATION FOR AK. GEO PLATFORM OR SPACE 880 RADAR ANALYSIS ZW - DL"W YARIABLE FOR ALTERNRTIYE APR PRINT OUTS ZX, ZY - PRINT CONTPOL YARIAEL.ES ZZ - INPUT YAP.IRBLE FOR TYPE OF PAYLOAD DESIRED ZZS - PRINT CONTROL VARIABLE ASSOCIATED WITH ZZ

A3-`l APPENDLK 4. OPTOTV PROGRAM LISTING (BASIC) to CLEAR 290- OPTOTWOOR 19 PRINT "ENTER 9IF WORD IS TO BE IN OTY AND SHJTTL V PRINT 'ENTER i IF PAYLOAD IS TO BE IN SHUTTLE ONLY' PRINT 'ENTER 2 IF IN 0 SHUTTLE FLIGHTS ARE TO BE LWIDERED' iWIJT`ZZ*;ZZ 19 PRINT '(1? 3GEO PLATFORM HT KYSIS' PRINT '(2) RINJLRR ARRAY' • PRINT 4(3) SPACE RM PADAR ANALYSIS'' INPUT 'ZU': ZU ?5 PRINT' (1) SBR t)pNSTANT IN. N=2 ANALYSISÌWJT ' ZO'c?0 4A READ RS, EM, T2, D2, KL WL EE, E, FB, DM, DD, TN, TY, IF. TM, TD. SP, SL DT, RH, MH, T9, TZ, TG, RM, B, 81, Pl, LM ' LD, G, WL, WS, CD, BP ' MY, DY ' LQ, KS, KT ' KP ' K Q, W n, M, D8 ' GE. A TL, LF 58 DRTR W. 3E+96, 925,1.8, 2,1 . 187, 48.9E+96, 3.218, 2, . L 9L 41, 94, 95, @95.33'.38, 37998, 3, 963, 6, 982, 92 894, 95:3,1,199, 3 A 1542654,1999' i9, 32.17,3. X4M, 69999, 914 ' L 85, 9L 684 ' t5.4 t' L 5, 18562, 47, 36, 354E+9r ' 19, 125, 40 59 ZZVI) ="i PAYLOAD ' i OTY, 1 SWITTLE' .rcIQWi PRYLOF.a t SHUTTLE' ZZSQ '1 PAYLOHTQ, 1 OTY, 2 SHUTTLES' M LPRINT'OP10TY OOR - ' ZZS(ZZ+i) IF ZLOt THEN LPRINT 110 PLATFORM RKYSIS' ' 0f.s(19) ' CHRS(18)' rlOTO RWELSE IF ZU=2 THEN LPPINT 'R NNULAR PHASED fEJ+flY" ' CH4R ( l8) ' CTiRfci9) GOTO 89- ELSE IF ZIJ=3 THEN LPRINT 'ST3RCED WM RRDAR 9NRLYSIS' ' CHi^fî9) , CHR3(i9) 7 5 IF Mal THEN LPRINT 'SBA CONSTANT TN, N=2 f4^Y5I5' ' CHR^^ 19^, ^ 5^ 19) N LPP.INT neid. '80' "BP'''CD''"D2''"DB''"OD''"DM'''DT'''DY''"E''"EE'''EM''`FB'''G''"GE '''Ki''"KP`'"KQ'' 'KS','KT','LD'''LF'''LM'''LQ' "MH'''MY''"PI'''RH'''RM','RS'''S1"''SP'''T2'' Q9 LPRINT 'T8'''TD'''TF'''TG'' `TL','TM'''TN', -TY-, -TZ` ' 440 ' 'WNg . "WS'''WT'''WU'''ZH'' CHRS(18), CHP$(i@) WO IF Zt#i THEN LQ--372: 1M=197 2^ WT=844- DM=1 DD= t TD= 995- TM= 9i: ZH= 46 WL0=1d5JY=2:LD* ELSE IF At-2 THEN FEM t19 LPRINT Bi, B, BP ' CD, D2, DB, DD, DM, DT ' DY ' E, EE, EM' FB, G, GE, KL KP, K(b KS, KT, LD, LF, LJt' LQ, MH, MY, PI, PJH., RM. RS, Si, SP ' T2, T9, TD, IF, TG, TL, TM, TH, T Y, TZ, WL, W WS, WT, HU, N, M, (19), CH PViB) - MINT STRINGM1''9, 45) - LPRINT STRTNGS(i29, 45) L'8 LPPINT STPINGt(2A38) PRINT "OPTOTYMOR IS RUNNING'- LY=LM'TB=TN-TC=TM-DC=DH-TIkTX-DY=OX LzO FOR WQ TO 2 STEP -3 05 IF ZZOtc-1 THEN N=2 149 TN-TB - TM=T r. - D*W T.:=TU - DXfiY - LX=@- tW WY=WU t 4 LZ^M- Tn=TM+TD- DX=i?M+M 168 IF N=6 THEN N=5 VO TV=(TX*i999+ 5) TW/=(TM*iW 5)- TDY.=(TD*t898+ 5) 1* FOR T1X=T1d, TO TV.. STEP TDY. L49 11=1i';.^1899 209 D): DX*l%+ 5) • DM%=, LM*1(4+ 5) DDY=(DD*109+ 5) 218 FOP fM%*=[Mt! TO D g; STEP K-% i'9 T,":=TY*198+ 5 THb'=TN*iW 5- TF%=TF*i98+ 5 249 FOP TK=TI ' TO TYi STEP TF'r' 29 TW=TWIA198 268 C1=3*Dl+4*02- A= 5*( 2*Ci* , - 866)+t Cif 2+2*RSI2)[ 5) : LX=L+96. IF LZ(8 THEN LZ=29 NO IF Z -2 THEN A= 99 :89 IF 3)k1 THEN LX=LZ 2?9 FOR LA= LZ TO LX STEP LD 7* Ar=.'*PI*lf82r144 ?16 FOP. TA:TB TO TZ STEP TG ,20 IF ZTkt OR Z1^2 THEt ,: k?. KY--5 ELSE W-4 KY=9 ? A WP=?*WHL'A- W►T=FB*AH*(KL^*PI*1K*Ti* M4Y*PI*02*T2* 744 , TE=TA*- f 9:47,. M)[ 5-1) Wi=TE*PM+4t IF ?tJ=1 THEN Zl x+lA*4 M ELSE IF ZU =2 THEN WW3=^WP '49 FT M^ TO 1 STEP DY ?59 IF Zlkt THEN 60 ELSE 379 <69 lIi=11V WZ i-M^)* c WS-481) WX--WS-W F41^1^h MZ R5+5*19[-7*N IF M2?MJ THEN 49B ELSE 4i8 ',N IF 7J^2 THEN 3rd+. ELSE 399 ZfiR W =PI*WS.+±A[2 41-1 ii*2*PI*LA Wl0ZH*twC+W8) Wlot+WB+WH Q= t i-HMJ w WS-M) WX=i6-H81 Pwn)*Wt- MZ= 85+ 4*16[-7*PN IF M:"MJ THE N 4ê ELSE 416 ??A I *00 A WC*i*3A A(2* 8ho W=14 4B+WC+±+*T WWDt 81 1 60= 1+MH WZ= f 1-M l %*4lS-WW) - NX=WS-W , "J*WX MZ= 85+5*19[ -%*PW IF MZ`!MJ THEN 4* ELSE 418 4A8 HEST M! 419 LP-t2* 1 12+ 15*i9(-4*PW) IF ZZ=1 OP ZEN' THEN LP--A 409 IF tlJ=2 THEN LB=LA-LP-LF D4f157 s8-9 D*11 LX- DA[2-DK2,*2*A'.'f:*TL))[ 5/2 U)=LX FS-A+LF+0 1r19 IF A?LB THEN t?19 ELSE 44 Cr A4-1 439 LB=I.A-LP' Ltt=P^6/C2 LK=LU- LC=LB*LP/LX FS= iLC+LP)/LQ IF ZU=1 THEN Lid 440 ON EPROR 130TO 16% 459 H=2* Rte' K=1.5*PI*FI[2*Di*Ti*f/(LF(3)- 8E=(387*(NN+NZ)*K/(F1H*0)(.5' D=CD*(K*NT/G)(. 5: BF=(387*(WHC)*D/(NM*11Z))[ 5 460 W(BE[2-*E4/4)( 5: AL4)F(2/2: FI=RTN(801(-AL)) Fi=RTN(2*RL*W(BA(2-U2)) 479 GOTO 879 489 Gi=-Flt*COS(Fl)+89*SIN(FI)+FTL*EXP(FL*B)*COS(B9*O+Fl)- (FL*B)*SIH(B9*9+F1)- G2=D/tB*(i-W2/(4*K))( 5) , G3=-(AL/B8)*C0S(Fi) +SIN(Fl)+(ALiS3)* EW FL*8)sC% (BB*B+Fi)-fXP(AL*B)*SIN(BN+Fi) 490 64=FL*SIN(FI)+B8*CDS(FI)-FL*EXP(FL*B)*SIN(B0*B+FI)-6'B00P(AL*B)*(Z(BM+FI): G5r--'FL18A)*SIN(Fi)+COS(Fi)-(RUI*)*EXP(FL*B) ►-SIN(B 9* B+Fi)-D?(FL*B)*COS(B8*B+Fi) 589 85=(1180)*ATH((-C,4*62-U5+1( S)/(Gi*G2+G-IWA)): T=B5: IF 85(0 THEN T=BS+PI180 510 KZ=KZ+i 523 K2=(i1K)s((plB)*(EXPt-fL*T)*SIN(BB*T-FI)-E)(-FL*(T-B)t*SIHtBO*(T-B)-FI))/(i-D*BF[21(4*K?)[.5+(Keg)*rg+rug0)*(E}^pr-^*T)*SIH(B0 *T-F1)-e?(-K*(T-8))*SIN(B0*(T-B)-F1) ► )) 539 IF K2(i THEN 549 ELSE 500 540 TRS+PI/88: IF T110*P1188 THEN 18% 559 GOTO 529 509 IF KZ=1 THEN 579 ELSE 58A 579 K?=K2: BiX=B9*Bil(2*PI) T=T+Bi7.*2*PI/89: GOTO 51A 589 K7=1(2: KZ=O- 84=81 598 BT=B4+BP: KY=KY+i 090 Hi=(-AL*COS(FI)+80*SIN(FI)+AL*E*(FL*B)*COS(B%+FI)-BB*W(FL*B)*5IN(BW+FI l )*E:=P(-ALs81^ Q---k*DP(AL*BPp s (.BmT+FI)+BA* EYP(ft*ABP)*SIH(B9*BT+FI)-k*M(88+81+Fl)-B9*SIN(B0*Bi+FI) 019 H4=-k*EWP.AL*BP)*tU5(89*BT+Fi)+BB* Efi FL*BP,*SIN(B%T+Fi)+HL*COS(8B*k+Fi)-WIN(B0+81+F1) y2g tfi=(r^*SIN(FI)+B9^ (FI:-AL#XP(AL*B)*SINtBA*B+Fi) 3+E) (AL*B)*COS(B9*B+FI))#aP(-AL*Bi) F(Z.A*4xP(FL*^BP)*SIH(BNT+FI)+B0#X Pf AL*BP)*r0;(60+ T+FI)-AL*SIN(BB*81+FI)-BBOS(OW,I+Fl) <:9 H7=(^ *SIHtFi)+69+rOStF1?-AL*4«=PtAI*B)*SIN(B0*B+Fl)-B0*E^(FL*B)*(OSi89*B+f1))*E/P(-AL*Bii: HB=L*EXP(AL*BP)*SIH(WAT+F1)+BKY, P(t1*^')**ÔS g9*^T+Fi:;-4L*SIH(69*Bi+fl)-$9tC45(BA*61+F1) 64A B9=(L'88)*ATN(-(BE1BOf5+iBE/(BT-Bi))*N5+(K/(D*B))*H7+(K!(D*(BT-Bi)))*H2)/rBEi!R*d+BE#,2;roT-Bl,+(K./tD*B))*H3+(K/(D*(BT-Bi)))*H4) ) T=B48T 650 IF KY-1 THEN 609 ELSE 718 e09 Jill*EXP(ql*gp)srt}5(B9*T-Bi-BP)-FI)+80#:=P(AL+^P)*S1NtB9*(T-Bi-BP)-FI)-AL*C'OS(B9*(T-Bi}-FI)-ÎH(gg*(T-f31)-FI) 678 J2=L*b?(AL*HP)*COS(B9*(T-Bi-BP ?-Fi)+B&,O?(RL*BP)*51N(89+(T-Bi-BP)-Fi)-k*COS(B6*(T-Bi)4l)-W- IN(B9*(T-Bij41) gg ►:=AL#)P(FL*T3P)*SIH(BA*(: T-Bi-BP)Î)+ #}{P(^.*$P)sCOiS(BB*(T-Bi-BP)-FI )- ►^*SIN(B8*tT-Bi)-FI)-B0^y15(B9*(T-Bl)-FI ) 690 i4^*frF(L*^P)*SIN( B8*tT-Bi-BP )-fi'!+B8*ESP(FL*9P)*COS( B9*(T-Bl-BP)-Fi)-fL*SIN(6^3*(T-Bi)-Fi)-EA*C0^(F33*(T-Bl)-fi) i99 BC=-(L'gA)*ATN((- *.)?-KsJ4Û)i(BE*.11+K*12/D)): B4=B1+B(:- GOTO 598 719 KY=9' KX=KX+i 729 K4=(11(BT-B4))s(-BT+B4+(i;'B9)*(EXP(-FL*(T-BT))*S[N(&(T-BT)-fi)-EKP(-&*tT-84))sSIH(BB*(T-64)-F1)))- 65^?-B1 7'?9 K?=(D,^t*((gElrB*gAj)s(E)^pr-^sT)*SIN(B^1*T-FI>-E^(-ALs(T-g)sSIN(69*^T-B)-FI))+(p,Et'rrBT-B4)*B@))*(E)T"- ►L*(T•-BT*SIN(BA*(T-BT) -FI)-EXP(-FL*(T-B4))*SIH(&'T-B4)-Fl))+(ViD)*(Ii;?)*(B+(i/B8)*(EKP(4t*l11*SIH(&T-Fl)-E` 'A*(T-B)1*5INtB9s(T-B)-Fl!))+K4)) FQ !F 0,'l THEN 759 ELSE 759 N T=T+PI,/89- IF TAT+18* IM THEN f 4 EUE 728 760 IF KX=1 THEN 779 ELSE 7% 779 T=T+P1/89- K'Al GOTO 79 l81 Y6-13 09 A IF K5?K2 THEN VW. ELSE ((0=K2 899 Bi=Bi%*2*PIlR0+85 819 FOR T=Bi TO Bi+BP STEP BPA8 R29 Kf i DAK *( ^ BU( DO))* f EXP(-k*T)*5IN(89*T-Fl)-EXP(-AU(T-B))*SIN(69*(T-B)-FD) CBE!((BT-Bi)*BA);s(-E{Pe-k*iT-BT^'sSIN(E^*(T-Bi 41--BkK-+(K/Dv*r(1/B)*(B+'i,•B9)*(EXP(-AL*T)*SIN(g9*T-F1)-ESP(-FL*(T-B))*5IN(89*(T•-B)-F1^)^ R

A4-2 q78 0--2' IF :1122 THEN. WELSE 488 N=t*PI*lR*Hi*TFMkA'. 4kW'TE 84 0* %613 • Ai=P[*Dl*Tl- I=2*ft*X (2+ r R* RT=3*Ri 8U=E*Ir GE*AT"&1 )- KR=1+RM+IIJ' tai-"slJ KC #BiYA' M14i*LA(2*(- 25+ x'-KC1?+3*PT,B> PI) N0=2*PI*LR A- 966-XB) GOTO q68 V. MF4ti*TW0*LW(2*TE*EM) 410 82=x'?*MF)( r113)- Afz-(W^(^il3). NRotl*MA*TN*tAI(2K(N(ARW 9MWTE- IF SM)SP THEN 428 ELSE 10 M. NEXT TA

Q?8 4)=K*TN+40*lAtt 2*SIN(B2)) • FDc**LA* QbSQ*COS(B2) • FC A*SIN(B2)- FP0fKA*SIN(AA)* 5 • MI=FC*LA+IIT*LA*TIM4'9+ 5*0iA(2*TWO +FD*(DT+Hl• M2=1T*LR*TftO+ 5*4R*LA(2*TN*K8-FA*DT+ 289*Ni*Lfi3*TkW8 448 P=li1H' P )=FI • IF P T)f THEN P=PJ 45A IF r1I=1 THEN P=20 469 Ii=P[+(113*TL, 8- Pi= ?54*Di' PC=PI(2*E*IL%A(2' Pv--S1*PI*D'1*TL'ii+EE*M.,f2*PU211 478 IF PY(PC THEN PC=PK 499 IF P= 991 RC TP"z Ai. WD tOl.; THEN 10x9 ELSE 1188 1930 Y1=17499 • Y2=167%- T3= 891- T4= 9 • GOTO 1279 L99 IF TP)= 881 AND TPA= 892Q AHD N=2 THEN 1118 ELSE 1L''9 1118 V1=1949A • Yt 17789-T3= X301J4= *2Q?-GOTO 12978 WA IF TN'>= A" WD TPA= 01 AND WQ THEN 7= 0024' T4= 91- Yi=15380 'Ql--14599 - PT0 1279 11.9 IF TP)= AO RND TV= 91 AND N=5 THEN T?= 90 Ta= 91 Yi=16799- '?M5708* GOT0 UN,, 1149 IF TP',= ice^? AND TPA= 91 AND N=2 THEN T3= 9829• TO Ai Yi=17798 'Q--17189- GOTO 1239 1153 IF TP)= 01 AND TP,*= 1 AHD H=a TH91 li68 ELSE 1178 1168 Y1=14599 Y2=14889.13= 91:14= i: !30TO 279 1179 IF TP)= 91 AND TV= 1 WO N=5 THEN 1148 ELSE 11Q8 1139 Yi=15790 - Y2=141%- T?= 91 T4= 1: GOTO 1279 1148 IF TP)= 81 AND TP{= 1 AND N=2 THEN 1289 ELSE 218 1299 Yi=171* Y?=i44% T?= ft: T4= 1 00TO 1278 1218 IF TP)= 1 AID TP" =1 RND W9 THEN 1226 ELSE 1238 Q20 A--14@@8l2)--I]-,W T?= i T4=1 9)TO 1278 12219 IF TP>= 1 AHD TPC=1 WID NA THEN 1240 ELSE 1258 249 Y1=1-189 Y 1599 T;= 1 T4=1 GOTO L7pj t2% IF TP'>= 1 AND TP,:=1 AND N=2 TO IV ELSE 1279 19 Yi=14498 Y2=13589 T's= i T4--i 130T0 1218 1'61 IF TP)= 901 AND TP 983 THEN Yi =l4k0 Yd=18499T'-,= A81- T4= 983 WTC 1278 1252 IF TP)= H03 Wd4 TP,'= M THEN Y1--i44%- Y2=19`B9 T?= 093 TO 91 GOT0 1270 1263 IF TP>= 01 AHD TP<= 1 THEN Y1=14599 Y2=14689 T3= 91 T 4= 1 GOTO 279 1264 IF TP"= 1 WO TPA:=1 THEN Y1=14698 Y2=13590 • T3= i T4--1 17p Y=Y1- ,'Y! -Y2)/iT4-T3)-*,TP- T< TT=TPOC IF TT?=6% THEN TT=c'444 1175 IF Z0=1 THEN TT=4KW*TP IF TT)=2563 THEN TT=2449 I," I5=459-5 44* o 1E-10)* 1 2599-TT > I3 2 TT=TP*NS 12'455 IF :1121 THEN TT4@MO*TP 1'YA TS- ' ISPTT*36.39 , IF WO THEN TP=25 ELR IF N=5 THEN TP=113 ELSE IF *2 THEN TF=S 245 IF Yr-i THEN TS--W f TP*115?20

A4-3 08 TWP+TS IF TV.--i2 THEN PP=VP*TR ELSE PP=12*KP4#Ek«(T(M2) 1318 PL2YS*N+KT*TR+PP 1329 IF ZZ=2 THEN WXzl6-WV- GOTO 13$8 09 WY=(HHSiMJ)*(EXP(-W(IS*G))+?Wi)-PL: WX:46-WV • IF 7J=1 TO 1:48 ELSE 1959 1.340 IF 98*WV)441 THEN W =01 81:GOTO 3348 UN IF WY?-II AND MIC THEN 1370 ELSE 1488 L69 IF WS)=WW AND Pe.Pr THEN UN ELSE 1498 L78 LZ=.A- IF (LA+LDKLL THEN 1348 ELSE 1499 1:30 IF WS'-W RHD WV)z* AND HS:>-*r. fir11D MIC THEN 1378 ELSE 1488 ls?0 A,N)=R- 88'N)=ffi- AL(N)zk: 89(N)=80- B2*=82: B4,N)=B4- 89,N)*5 86(N-E'h &N-P- Er. , W=C BEiN)=BE- SF-N-W CIA'- :1 !'.N '-C2 M'N'-ft: D I N"- ' D;n:+N-0'X Ft'N)=F1 FP(N)=FA- FCiN,=F,:- FD , N)=Fi, FI-N-FI FS-4)2PS' ZI,N-ZI 1400 w, N! =H Us N 1=11- IS(N>=I50(w-i0- Qt N)=4(2 k5'N,#5- k0,#6' k7^N)=► 7 0,H)4? ' , N-^ KE, N-4E LA'W=LA' LB f N l=B' L C(N)=LC- LP(N)=LP- LX(N)=LX- M1(N)=11: M2(N) =M2: MF(N)=W- MYN)=NJ: N(N) zN: NR(N)=W- HD(N)=ND: P(N)4 • PCt N)=K,. PI(N)=PI 1418 R(N)=k PW(N)=PW: Pi(N)=R1- c-H(N)=SM: Ti(N)=Ti: TR(N)=TA- TE(N)=TE- TP(N)=TP- TT(N)=TT- TW(N)=TW: TX(N)=TX: V(N)=V . W1(N)--K WR(N'=WR: WB(N)--W: WC(N)=WC: WH(N)=*: WX(N)=WX- WW(N)W WY(N)=WY- WZ(N)=WZ: PL(N) zk TG(N)=W TP(N)=TR• TS(N)=TS 1428 W(N)=W- hM(N)=AH: BU(N)=BU: Ri(N)*i: AT(N)=AT: XB(N)=XB: 1(N-I- DAMN O: NO(N) zm- WP(N)=WP 1470 IF LLzLZ AND TW=TB THEN 1678 1448 IF LL=LZ THEN 1458 ELSE 1479 1.450 IF WV l rAM OR Di(=C THEN L718 1.460 W4-DD- TM=ii+TD , TX=Ti+2*TD: GOTO 178 1470 N0'.T LA 1489 LZ=A - IF 5t-TB AND "(N) AHD LAXA(N) A10 D4-=DVN) RND T1=TVN) AHD DOWN) 130TO 1679 14'43 IF TW=TWiN) 40 H=N(N) AND LR>0M) AID Ti--TI(N) AND Di=1(N) THEN WTO 1718 1599 LZAA-LD: LL=LZ: GOTO 70x8 -1518 NEKT TW/. 1528 IF LX-LR)L.D THEW 1628 15-1 LL=LU- L%LR: TN=TW: IF LX-LR

A4-4 1770 NEXT N 1788 IF ZW@% THEN i889 1799 S1=35088 E=1. 81E+37 - ZIt2N+i : Me 81 1?0. 95 1*2 : RIB t: OOTO 88 M IF DK?i THEN 1829 i8i8 Si--4E+97 E=4F+07-1Lz 6E-84-ZWZW+i TN= 81:T!k FA: W2: R* 063: 1F Zth 3 THEN 1838 1812 OOTO 86 1828 IF ZNO2 THEN 1848 N`-3.3E-04: TL=. 896 : ZWzZN+i : TNs 91: DMs2: TMs 95: I F Zlk3 THEN 1858 1832 OOTO 86 1848 IF 2NO3 THEN 180 19% TL= 125:2* 6520301: TW 91: N 95: DM=2: GDTO 98 1%8 IF ZN04 THEN 1888 1878 3He 47 - R#i5: Z14`ZN+i: TN= 01: TMs. 85 : D*2: OOTO 80 .889 END i" PRINT 'PROGRfII1 PPElWPJRELY FIN190 - K2 CFLCLLRTION IS NOT AID' iQ% PRINT 'PRO" PREMRRIRELY FINISHED - C CPLaUTION IS NOT YfILID' Vi8 PRINT 'R IS TOO LARGE'

A4-5 APPENDIX 5. OPTOTV COMPUTER SYMBOLS AND LINE NUMBER CROSS REFERENCES

9 IN 149/2 268 418 438 589 588 718 7% ON 1548 1579 15% 1698 VW 1768 1788 1 66 78/2 75 189/3 135 268 289 44—A 338/2 348 3`A YA ?% .W 418 438 488 5% 518 529/3 538 568 598 648 a38 788 718 728/2 736/2 746 768 828/2 838/2 89813 918/2 9% 969 1819 1848 1858/2 1852 1216 1.220 1238 1248 US 1268 1264/2 1275 1288 1285 1.295 1328/2 1338/2 16281668 1788 17981888 1818 1838 18581878 2 69 79 i8912 128 139 135 268A 278 X9/2 328 338 3'/9 302 398 418 42815 45812 468/6 488 528 578/2 888 83V2 878/2 M 898/7 988 918 938/3 950 4,W4 1188 1148 iLa9 12% 12x9 220 1468 1568 1688 1798 1816 i$29 1838 19% 1878 3 68 78 138 268 329 338 398 428 458 898/4 918/3 938 968 1818 1832 1848 4 268 329 418 469/2 488 5281818131838 1%8 5 168 328 ?69 388 3% ON 1L^ 1179 Un i^.'^8/2 6 168 269 338 39812 918 7 369 389 3% 1798 1818/2 8 8% 966 9 IN 29 IR660 W18 11`9 129 1298 A 79/6 75/2 9912 .8/2 368 386 399 418 548 756 1," 12'99 168912 4.78912 1748/2 12 410/21,% 17-6 15 1$78 18 818 29 268 25 1299 45 118/2 98 79/717991$12 1832 1859 1878 189 299x?. n238/3 2% 1558 128 11812 Ll 128 144 ?A9 178 1469 1689 299 1568 299 18 IW- 3% 278 ?48 1349 ?68 'h8 379 458 34 im s8r 459/2

498 368 ?IA _99 418 369 ?A "98 448 429 45A Q% 518 578 528 VA 549 539 W. 530 578 569 A5-1 See W 90 700 668 656 718 6% 713 l2B 7% 756 746 760 740 770 760 7% 766 A44 leg 879 470 WA 87A OA6 879 028 418 QsA 418 069 899 999 O9A 18Ae 170!? LQe 049 Me 998 182@ 101e tel 18!@ 1649 088 WA 1@58 1e. 1" to% 1668 Ii@@ 1869 1110 ilAA 1129 11@9 1168 1158 1179 1158 1188 1178 1149 1170 12''@6 it% 1216 11_Q9 in@ 121e 121 i21e 1.140 12'8 259 12_78 12'08 _?A 1261 1852 127@ 181A 10% ill@ WN 11-70 We 1169 1188 1288 lrV 1249 1250 1266 1261 U''f,2 1263, 1.49 1?-78 1_'58 1-7?9 1-769 1848 078 1356 1"she 2W LA9 028 1_% Lib 1459 1449 14N. 1448 1499 1759 U A 1.779 0% 1599 1898 Me ie?A Me 1851 1609 1?% 1748 13"19 1888 M 18-78

A5-2 15R M. 1578 1570 1559 1598 1548 1698 1538 1628 1528 1599 1649 1628 1659 449 1668 16% 1678 1438 1488 1608 1798 1698 1718 1459 1498 1698 1768 163!1 1778 1759 1789 1769 to 1188 1828 1888 1838 1818 i849 1828 18% 1838 to 1849 1899 to 1899 549 LQ99 i59 iQi@ 429 2499 UN 1275 2598 1279 1275 1289 :699 12?8 low 199 131. W 1Z'8 1249 1269 1264 140 1169 128 14188 1189 1248 14499 1298 1269 14599 ;'.1281108 146A9 Qb-11264 15:0 1976 " 157% tal1188 16798 10% ill 17188 1140 12" ir"?99 1878 1r7`!9 1119 1149

18499 1119 18599 1262 1263 (6999 1101 UQ 14999 1261 3 +918 17% 4A999 1275 1285 A 269 278 338 429!? 4?8 459 89914 %@ 131% i3V 1718( At A%t! 1428 1429 :719( AA M129.19 UM 13%( 17191 W ?9A 13* Is%( 1719 AL 469/4 4*18 4%A 5E914 ;`h/9 610.. 62E0 638/9 569/- a?9,4 68914 M4 72k2 739. a 82'8./5 8?9/2 t?99 1?99( 1719,

A5-3 AM 849/3 1428 1428( 1718( AT 848/2 1428 1428( 1718( 8 49 118 489x5 49919 59812 528/7 6W4 618!4 628/4 638/4 64814 778/7 828!1 88 466/4 49819 499/8 W2 Wl 548/2 578!2 689/9 61MO 628118 5-8/18 648 669/6 678/6 689/6 698/6 788 728/3 738/9 7W2 778 898 528/9 839!'2 13% 13%, 1718( BI 48 118 578 51 a 598 69913 619(3 62813 039,3 64914 66914 ROA 68814 698/4 799 728 9% 9991 819/2 828/2 8_7&'4 82 910 9-9/? i-748 1398( 1118( 84 589 549 799 72814 7-9/3 1398 1399( ili8( B5 589!4 549 903 ON 248( 1119( 86 728 1798 is`98( 1758( B9 649/2 728 1_5548 121W 1718( BC 788/2 13-8 L748( :118( BE 458 469 649/4 788/2 138/2 82813 838 2% 1708( !718( OF 458 46W 488 529 1398 099( 1718( BP 49 119 599 0A912 61812 6291"2 63&2 668/4 67814 699/4 699/4 81812 BT `Q9 098/2 iWl- 628/2 6:9/2 64M ?28/4 739/? 759 82!12 R-9 N 899/2 1428 1428( 171.8! Cl 2691? ON 1718( C2 268 4-1298 1`(98( 1718( I'D 49 118 458 D 4591 488/2 5'012 640/4 799/2 7-812 82911 090 1399( 1716( D1 2' 228 VI 206/3 320 428 459 899 96814 298 1, Q9( 1458 1468 1489 1499( 1498 1498( 15%( i5AY/. 1578( 15881"/. 1689/2 1719( D2 -49 119 268/2 -739 DA 42612 1428 1428( 1718( D8 40 118 429 DC UO 140 1450 DD 48 189 118 150 .M 288/'! 2IM4 1408 1549 1569/2 16N DM 48 109 118 128 148 159 298 288x7, 218x% 1468 1560 1009 17-8 1818 18-78 185550 1878 DT 40 118 9:812 * 49 118 140 * Ll 149 150 288 2W1 21VI L9N 1:98( OW" 1568x? 15791 16 +* 1710, DY 128 149 E 48 118 258 459 894 909 1269 17812 1818./4 18-8 EE 40 118 969 En 40 118 Q% F1 460 4%/4 499/4 51812 61818 0:'918 67V4 MA 728/2 779 2 KW2 8_'8 1-739 13?9( V10( FA 4-8*2 1298 1.,%i 1718( FP 48 109 118 ':'0 FC 4 k 2 1_?% 119955 1710, FT) Q?►3r2 139(' i-743, 1719( F1 460 488/4 444 "12 t*-%! QO/8 --&4 s;W'4 719/4 `3 /? 1-733 L7`^'( 1710, A5-4 FS 420 438 1398 1398( i71e( G 49 lib 4581328/2 1338 Gi 488 588 G2 480 50812 G3 489 508 G4 490 588 65 496 580 GE 401199% H 012 930. W 1408 1498( 1716( Hl 660 648 H2 688 64912 H3 610 640 H4 610 640 145 620 648 H6 628 640 H7 639 648 HB 638 I 899131420 i42e( 1718( It 96912 14981490( 1718( 1S 128812901326121338 1498 AM VA( It 668 7% 12 07% 13 6% 798 M 698 700 K 45913 488 5@912 52913 64914 798/2 73012 828121499 1499( 1728( K8 79912 86612 870 888 998 919 93M 148814e8( 1718( Ki 40 116 26e K2 528 539 578 588 79912 1498 1486( 1716( K3 73e 740 779 780 K4 729 739 K5 770 7991214991400( VW K6 7% 1468 14@8( 1718( K7 5% 1489 1488( 1718( K8 820 83812 84812 K9 578 1466 1406( 1718( KO 89912 KB e9812 KC 89912 KE 84012 9M Ae9 1488( 172e( VP 40 il@ 136812 KQ 4e tie ON KS 48 119 1310 KT 48 lie 1318 KU 32812 339 KY 32@12 338 KX 71912 760 786 KY 59812 6% 718 K? 51812 566 5% LA 296 388 38812 39612 439 458 898 899/4 998 9i9 936/i8 137012 1468 1498( 1478 1489x2 1486( 1499 1496( 1598 1526 153812 1728(2 LB 42912 43913 1499 1469! 1728( LC 43613 14961468( 1138(

6 C ^ A5-5 LQ 48 188 116 248 1378 1568 1528 M LF 48 118 420/2 LL MSI 1448 15881538 1678 i71e LM 48 188 118 128 156 ?39 LP 41812 4284 4384 WS 1400( 17.18( LQ 46168 ite 428!2 43812 LU 426 438/2153816761716 LX 266 286 298 42812 43812 1378 1488 140 1528153'0 !738( LY 128 L2 150 26812 288 290 ,78 1438 1448 1488 1502 1538 Ni 89612 938 46 1488 140 17381 M2 938 94814881488( VIM MF %J 91812 1488 il88( 1738( mm 46 lie 3330 M 348 ?6813 :88/3 39813 468 ift I'M 0.4114e010W 1736( IW 48 118 318 M2 368V'2 3882 3982 N 138 US 168V2 M 188811881128 AN U481158 117811.x8121812-18 iz'% !2.4913 13181398x'25 i488127 i488( 1418125 14 MO 14Aith: 140 1498x6 1440( 1" 157@ V00 1718/391728x5 17 x39 11`36( 174815 1778 41 9182 938 149►? 1488 1738 NP 9?0x3 1488 14@8 17 @f 40 89@ 1420 14281 173@ P ?IV 4@x3 9543(21363 13 3 14M ? 1488(',41 ii^ K SO 979/) 04 1:631:3@ 1488 10 , 173@( PE 40 110 A 3x2 M'2 458 588 548x2 5'13x`2 75Ar2 ''a Pax 0 9104SM, P I 90,", 14 3 i4@@- 1 ` 3 f Ph 968 1410 1410, 1770' PL 1319 1.20 1330 140 1410( V?0: PP 1306(21'318 PN ?60x'2 A/2 39@x2 418 12N 1418 1418( i : r +?t 155@x;2 Pi 968x`2 1418 1418( 1738( PH 48 W 3?8 /7x8 i818 PH 4e 118 339 RS 48 u6 x'68 Si 4A 118 %0 I'M i8i8 S1 8% QM2 1418 1418( 1738( SP 48 118 918 T 5801'2 %*tk 5481, 57812 648 66814 678/4 6W4 69814 72814 738112 75813 7782 818 8WA 83813 8% 1738( T8 48 118 318 T1 186x% 198 198P/. 3338 456 898 96812 1418 1418( 146812 1488 i488( 14981498( 15781Y.168812 1618!?; 1738( T2 46118 338 T3 1879 1898 1118 1128 11381148 11681188 12891228 12481268 1261 1262 1263 1264 1278/2 T4 18781898111611:8 "t 1148 41681186 i2e61228 1248 1268 IM 11''62 ' F7 '1264 L?78 TR 318 338 428 1418 1418( 17"38(

A5-6 TB i2A 146 1438 1488 1668 TC 128 146 TD 40 is6 119150118 ilAdY.1891Y. i469/2159616AAl2 TE 338!2 ON 90 918 1418 1418( 1738( TF 48 its 239 21 "% 2W4 TG 46116 ': TL 49 ii0 q '; :63816% TM 40 l991i0 120 148 158 170 17818 IWX !4661571/9: to 090 leis 1830 lo% We TN 40 118 0 149 238 M. 248/8153616001196 IN We 19% i876 7 1866/2 is61186A/2168s121186/2 1120/2 1130/2114612 1158121118/2 i19A/2121912 123812 !256/2 i26i12 1262/2126312 UR/21218l212151286 !28512951418 1418( 1738( TO 13W4 019 1418 1418( i738( TR 1290,31380 We 1416( 1136( TS 1290 1295 1366 1410 1410( 1738( TT 1210/3 1275/3 i28612 1285129914181418( 1738( TU 129148 TW 248/7. 250 25616 886 906 916 93616 1850/214181416( WA(48812 Wk 1499 1498( 15181/.15.101680 1668 11891739( TX 128 148159 iA 178x"/. 19911. 1418 1418( 1468 1606 1736( TY 49 109 116 230 2301/, 24014 TZ 49 ile ?is Y 1270 1295 029/2 038 1418 1418( 1138( Yi 10781990 iiis iiA 10 iiQ 116811801288 1228 1246126612611262 12631264 1219/2 Y2 19161690 We 11261130 ii481166 U90 IM 1228 12491269 1261 126212631264 1278 W 88012 8% We 1428( 1738( A 338 386 399 896 9% 9ie 9381214161419( 1730( WA 338/4 388 3% 939/214191418( Me( WB X8/3 39812 l41s 1418( 1736( WC 307 398/2 14101418( iM WH 38@12 399/21418 AM 1736( WL 48 ii0 338 1818 1838 WN ;8119 ?3e We WP 336/214201428( 1738( WS 48 iie 36812 386/2 398/2 1829 iSM 1858/21279 12801326/2 133012 1360 O912 WT 48 99 118 396 450 93912 WU 40 iAe 118 146 WV 149 30 1346 111 360 38013 ?99/6 45014 1869 1920 1939/2 1858/2 04012 1359 068 089/2 1416 1419( 14581149( HX 302 38612 399/21036 028 038 088 1410 1419( 1739( WY lees WO 132812 03812 040 1350 1386 1418 1419( 1459 1748( WZ 36e 3% 390 45014 1416 1418( 1748( XB 8W4 1428 1420( il40( ZH 0 ie9 119 386 3% !859 1878 A5-7 Z( ?A i398 i38( !. O ZO '1 75 135 10% LOU 1275 1285 12 +5 r) 36 7813 i88t2 276 ^a8 326!2 33612 358 376 128 138 878 9% 1336 i91610 ZN 17661788179812198819161219281838/218181998/2 18681876/2 ZX 118 1668 1698 1788 N 118 1768 ZZ 26 6047 78 794 41812 996 1616 1818 1328 1628

A5-8 0 APPENDEK 6 SUMMARY OF OPTOTV PROGRAM INPUT-OUTPUT PARAMETER CATEGORIES

INPUT OUTPUT " PROGRAM OPTIONS ZU = 1 GEO Platform Analysis ZU = 2 .BR-R Analysis ZU = 3 SBR-A Analysis Z Z= 0 1 Payload - 1 OTV - 1 Shuttle Z Z= 1 1 Payload - 1 Shuttle ZZ = 2 1 Payload - 1 OTV - 2 Shuttles ZO= 1 SBR, Const. 'TW, N = 2

* GENERAL • Payloads D2, G, ID, LM, L'1, DK, Il, LA, LB, LX, T2, Z :J1, P. PC. 1111, Tl, TX. ':-A, WR, WH, WW primary struts TM, TD, DM, DD construction materials E, Sl, RH, RM, SP

a Shuttle LQ, RS, WS

• OTV N, KP, KQ, KS, KT IS, LP, MU, PL, PP, PW, TP, TQ, TR, TS, TT, TW, V, %VX. WU, WZ - final T/W TN, TF, TY - mass fraction NIU, DV

Geoplatform FB, Kl, WT C1, C2, Peculiar H, K, PJ, PK, Z1

a► SBR-A Peculiar DT, EE, EM, FB, LF, A. AC, C1, C2, FA, FC, Kl, MH, TL., WL, WT FD, FS, H, K, LC, 112, ND, PJ, PK, TA, TE, Wl, WC - membrane TO, TG, TZ AA, B2, MF, NA, SM

A6-1 n INPUT OUTPUT

- dynamics B, Bl, BP, CD AL, BO, B2, B4, B5, B6, B9, BC;, BE, BF, D, Fl, FI, KO, K2-K8, KE, T

Q SBR-R Peculiar DB, FB, GE, LF, Al, AC, AM, AT, BU, DA, WL, WN FS, I, KA, KB, KC, NO, W, Wl, WC, WP,:CB

A6-2 APPENDIX 7

OPTOTV COMPUTER PROGRAM FLOW DIAGRAM

A7-1 CIA \ /^^ ^]^ ^ .^

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^ ^ I

^ n | f . ^ ^

n ^ ; ' ^_ , n

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APPENDDC 8 SOLAR ELECTRIC PROPULSION SYSTEM

SEPS Power Module

The following is an outline of the more significant algorithms and input data that could be used to expand the OPT OTV computer program to include structural models of SEPS type Power Modules. Material from AEC-Able Engineering Company, Inc. is the source o: much of this information.

• The Folded-Container-Design Configuration, shown in Figure A8-1 will be used to establish the stowed state configuration of the power module.

• Two automatically deployed ABLE booms are considered: (1) The ACR system which is a lattice boom having articulated longerans, and (2) the CC system which is a lattice boom having continuous coilable longerons. Both systems are sired on the basis of strength or simple G loading considerations.

• Following are sizing algorithms for the ACR system:

a) Bending stiffness EI = 1.5 C 1 E A;, R2 E = Youngs Modulus

C1 = 0. 9, joint flexibility - caused reduction factor

A11 = Longeron cross-section area

R = Boom radius

b) Bending strength MCR = 1.5 PCR R

PCR = Longeron strength based on Euler column buckling and bay length LB = 1.4R = column length

c) Boom weight W = 3 C2 pAjl.

L = Boom length

C2 = 3. 0, empirical coefficient P = Material density

I . A8-1 d) Retracted height of boom H - .76 Ld/R d = Longeron diameter

e) Height of canister H + 3R HCAN

f) Weight of canister WCAN KILd/14 + K2 3R2/5

R 5" 10" 18"

Kl l .40 .30 empirical coefficients K2 1 .75 .50

Following are sizing algorithms for the CC system:

a) EI = 1.0 C 1 E A t R2

b) Same but based on LB = 1.1 R

2 C) WB = 9 e PR e 2 L

e = Fractional elongation allowable in construction material

d) H = (3L/ir) (e + .005)

e) Same

P Same

The following materials and minimum gages should be considered for the ACR system

tmin = Minimum tubular wall thickness E p tmin Pat lbs/in3 in.

Steel 2.8 x 10 7 .280 .010

`Titanium 1.6 x 10 7 .160 .010

Alum 1. 0 x 107 .100 .010

G/E 4.0 x 107 .055 .025

I A8-2 • The following material are to be considered for the CC system:

S - glass epoxy e = .03 E 7.Sx106Pei p = .072 4t/in3

The c = .03 is based on prestressing the coiled longeroni to minimize the coiling radius. The longerous have a rectangular cross-section to maximize packaging efficiency.

• The unit area weights of the SEPS blanket as well as the weights of the blanket canisters will be based on extrapolation of state-of-the-art SEPS weights.

• This study will be like the SBR-A. SBR-R and GEO Platform trade studies generated by the OPT OTV computer program. The objective of the SEPS power module study will be to select the optimum payload and OTV parameters that maximize the area (cr in this case power generating capability) of the power module.

A8-3 t J

.-i 00 Q s, J ^ t

L' I

4 ,^

t `

AN-4 APPENDDC 9-1. SPACE BASED RADAR-A ANALYSIS (One Payload Using Entire Shuttle Cargo Rayr

1"W/M - 1 Fft n, 1 SNI m

M B 6P 0 D2 w M OT Dv E EE EM F9 G Ki Lb LN LI MN My PI IN Im RS Si SP T2 is TD TF TG TM TN Tv TZ K I!S WT

166 i i Si4 1 S 2 3 M *40 167 386666 3.218 3217 2 46 4386 681 6 85 114159 863 8513 94 371M 3336 .625 2E-93 61 04 9E-83 85 81 .41 62 3.3E-N 6M88 1

A AA AC AL 96 92 94 85 96 89 9C BE 9F a C2 s D DX Ft FA FC FD F1 FS N it IS K8 K2 h'S N K7 K9 K KE to LA(FT) L9 LC LP LY M! M2 IF Nu N Nk ND P K Pi PK Pw R1 SM Ti T TA TE TP TT Tw TX v wi wA we wt 44 wx ww wv wz

53 766 69369912 i 24423E+86 i SWK-4 4 356793 87436% 8 8 1 8 1 .356793 8574149 18 6.8 2 2 2739E43 2 5 8. 5M12E-N 59.488 186.816 !386.97 -1.57837 9e7-n 93.122? ism 1 2 / 1 8 e a 1.17347 1 5348 445 664 675 372 8 5022 751798 405131 i 37116E-% 85 9 8 254895 8148.38 9579.13 4359.12 9" U 5212 65 788 439.62 85 / 2E-63 4 58997E-84 8 6 81 86 8 3. S-S47E-N+ 332116 1773.5 26192 16945.9 6215.12 5.7784 9 8 ?474 %

52 5315 15843► W36 4.81666E-14 8896x95 119481 8 4 8 1 8 898695 8263431 11, 5 8 2 5 5.98176E-83 3 914556E-14 172 676 345.352 2789.44 -1.57834 984894 98 9846 747671 / 2 8 8 8 8 8 58729 / 4429 AS 333 684 671.12 a 449144 2 59249E46 ! 68649E+16 5 67424E-44 85 9 8 65."'.7 264?3 7 x267 4 18536 32267 4 1625.55 885 1135 83 13 6 2E-83 4 W"E-84 1 6 85 14 8 3 SME-N 823362 3639 26 17944 5 !8311 S 1912 41 39897 6 8 31321

A9-1 Appendix 9-1 (cont'd)

51386 .173885 6!8324 5.36694E-94 124731 .137378 8 9 9 8 9 124731 8327599 13 9.2 3 8.8321:-03 3.5 8.6842E-94 222 537 441874 3189.59 -i. 57837 .988473 98.8619 159843 8 2 9 e 9 e 9 iz 9835 9 3748 iii 667 684 679 643 9 3814.49 3.33999E+96 2 34222E+96 8.6423E-94 .85 9 9 .868982 37485 45791 i 26357.9 45791 i 4519.25 1962 1585.43 .15 9 2E-93 4.58997E-94 8 9 .99 .16 9 3.53547E-94 199628 4188.98 12847.8 !7229 5316.77 54683.2 8 3381.86

58.8893 .186132 452487 6.75262E-94 163372 .147733 8 8 8 8 8 163372 .8367495 14.5 A.4 3.5 8128488 4 8.26657E-94 238.271 476.543 3893 -1.57838 .977435 86.154? 2 69392 8 2 9 9 9 8 8 23.7962 9 3228 268.333 684 668.566 8 3294.34 3.6143E+86 2 68843E+86 187477E-93 .85 9 8 188442 416611 58917 38998.8 588!7 713116 1239 1742 A .16 6 2E-93 4.58997E-94 8 9 13 .17 9 3 53547E-414 133644 436134 9523.53 16268.8 8389.59 51619.4 8 3697 56

58.9893 283543 452487 6.75262E-94 163372 .161553 9 8 8 8 6 163372 .9367495 14.5 18.4 3.5 8329486 4 8.26657E-94 311586 623.17! 3693.38 -157838 .977435 96.7547 2 69392 8 2 8 9 8 8 9 23.7962 9 3228 268.333 684 668.566 8 3294.34 4.69486E+96 3.51564E+86 148547E-93 AS 9 8 129317 54116.4 58817 49523.8 58917 ?13116 1.239 2885.87 .16 8 X-C 4.58997E-94 8 8 .17 .17 8 3.53547E-94 13:.644 4393.34 9523.53 16268.8 8389.59 51618.4 8 3697.56

59.9893 .217489 441237 9.58862E-84 2 877 .172621 8 9 9 8 8 2 677 .8437736 14.5 18.4 3.5 .9137168 4 9.22545E-94 375.397 758.793 4159.24 A. 57934 .965293 86.7547 3.3674 8 2 8 8 9 9 9 .X318 9 3189 265 684 _;a 26 9 ?2,'%4 :4 E. 2493c 4 3e 4967??E+d6 1714 -93 ?5 ? 1.:7454 29:d 3 ` A, _ 1 2 72521. A 4 5 1997E-94 A 9 71 3.53547E-►34 1.64211 5221.9 9288.39 186879 686.332 59313.7 9 M. i3

48.88!5 2A455 346989 119874E-93 2 48836 .175769 9 9 9 8 9 2 48836 .9469199 16 116 4 917L63 4 8.8471E-94 351443 782 886 3822 7 -157935 978377 84.00 S. !}265.5 9 2 9 8 9 8 9 48.198 @ 2828 235 684 669.21 9 2882 32 6.91739E+86 4.84477E+96 181911E-83 .85 9 8 142539 71874.8 84923 3 57224.2 84923.3 M83 1.416 2472 27 2 9 2E-93 4.58997E-04 9 8 .25 .21 9 3.53547E-94 186844 5246.44 7384.38 17942 9 3368.94 56632 8 18165 A9-2 n W

L Appendix 8-1 (cont'd)

48.88iS .232W 316969 i 28569E-3 263625 !84603 0 a 8 0 0 2 63625 0491858 16 U. 6 4 IMM 4.5 9.14702E-84 487.674 811348 4215.7 -t. 57034 978377 01.6628 5.27788 0 2 0 0 8 0 8 56.6879 0 2820 235 604 659.21 9 2882.32 7.16241E+06 18=496 2 09973E-03 .85 9 8 157448 04599.3 88224.5 68759 88221.5 1849.86 1416 273!.74 .21 0 2E-03 4.58997E-4 0 0 29 .22 e 1 SWE-44 194777 5492 72 7384.38 18522 7 1218 4 58789.4 0 2960

47.6826 .233367 272787 i 37512E-03 12448 185224 0 8 0 0 0 11.3448 .0524427 V. S t2 8 4.5 .8217651 4.5 8.77453E-04 344.595 729.19 3758.97 -i 57036 981147 82.5862 7 51479 0 2 8 0 a 0 0 77.7551 6 2508 288.333 684 671 105 0 2548.04 6.63644E+06 5 54433E -06 211825E-83 .85 9 0 158376 80357.7 188328 67219 100328 33411 i 593 2747 89 .21 8 2E-03 4.58997E-04 8 8 33 22 a 3 2547E-04 217782 W55 574 n 17661.9 3933.06 56866.9 0 3239.7

47.6826 2424339 272707 131512E-03 10448 L47.424 6 a 6 0 a 12448 .8624427 V, 5 !2 8 4 5 81217654 4.5 8 77413E-84 488.788 811.576 4853.16 -157836 991147 82 5862 7.51479 8 2 a a e a 0 77.7551 0 2580 208.333 684 671 185 0 25404 7.42784E+06 0.216337E+06 2.37496E-03 .85 9 0 171006 89938.7 100328 752713 180328 33411 1593 2967.83 .21 8 2E-03 4.58997E-4 0 8 37 22 8 3.53547E-04 217182 W55 5740.71 17664.9 393106 56666.9 8 3239.7

47.6826 249533 164851 162773E-63 3 53116 .198005 0 8 0 0 0 3.53116 .8578565 V.5 U 8 4.5 8233361 S 9.21922E-04 434 602 877.204 422199 -157834 .965449 82.5862 8.23849 6 2 8 a 0 a 0 89 3821 0 2460 205 681 664 367 8 2548.84 8.41035E+06 7.13785E+06 2 58961E-03 .65 9 0 18122? 101837 189883 86419.3 189083 614 352 1593 3145.89 .23 0 2E-03 4.58997E-04 0 0 41 .24 0 3.535OHN 2.37352 5838.86 5558.48 18674.9 727.473 59272 5 0 2914 36

Aq-3 APPENDDC 9-9. SPACE BASED RADAR -A ANALYSIS (PPeeltne Configuntiar)

OPi0TW00A - 1 ft ft. 1 OTV. 1 SOTLE

r$ a eP m D2 OO OM OT u^ E EE a F6 G Kl KP KO KT LD LM LO X14 My PI RM RM RS Si SP T2 T8 TD TF TG TM TN TV T: WL WT

IN 1 1 414 1 1 2 3 ^^! 4E+d7 187 38d8t18 3. zl8 IFv 2 5 is 41 le 1?* iN 6 83 1 14159 463 4513 se 3r1)ed @ 5E-43 84 9E-d3 Eke di 41 d2 X-N 680 1

A AA Ac R11 AL 8i !+4 ^i tZ 0 tx @E OF L"1 01 9 u F i FR Ff FLT F1 M 11 1; Kd (2 K5 Ki FP ) M K KE LA LAO) t.8 Li LP 0 ml M2 R MU M w t[^ P; Pi PK PL PM 91 w Ti T TA TE TP TO TF T: TT rw TA v RI Ml Ne w 0 4w NV W1

53 d7t452? :.' 4. S21M-04 I. LW 0559183 a a a d d 1.17 d3 0318721 to o 8 9 :2':f-ds" 2 1 72^.%E-N , 105 18 2179 314 :N -1 5(o4 1 t^4dr3 4S 193 2 8 a d d d 2278 illy 167 543 0 11W k* 190 A 3978 aRr?d 5 41x34 6 5 8282,4E-45 1 9 d 143597 ^2 333 ;570 1? 440 a5"Q 13 v.% 612, 611 4dL^ o l^ 248 as a 2E-03 4 "?E-84 3 1333: 4i 4 3562 25 '15K l?i dd4 al dti5 IW. A I 3 5 7E-LM .124,116 +53 9112 47:3 tai 42 4 '2 464N 9 13%1-1 1 13594 9 55? 7 22

53 Ni 12440 7 279242 3 14121E-04 9Q,8 '.Q 49,41385 a d d a a at^"SU W371 to ' 44244,E-43 46 1 7 48045E-84 S6 5754 113.151 lee 81 -1. SZMa '1'8'535 931227 15748 451347 2 d d 8 d A le 0, 3? 8 4 2538 .`id 83' StlS 5311 319 ?y 1^ 4pw 3??e 4%W ,Flm ?.ANN* 97 9 0 451dp, x:52 +R s5^ 13 2,44,,4 4,6 9679 13 516 339 38:29 r t liti aS d 2E-d3 4 58` . 44-84 ft*t;61 29 1323 25 4 43224 1*1 S a3 d5S 14455 2 3 54,534,4,E-04 3 32116 a.w 252 S8y :1 %X 34 4404e 8 15 53 2 1e0d2 3 57.1i iA

A9-41 t r {i Appendix 9-2 (oont'd

SS 5184 15.'54 M455 3. 61051E-14 S.1% 1.11671 0 6 d a 6 96W W?1.4 16 3 T A• 2.1 r 1. i T 4a2.'SE-A4 1A/. 265 200.51 56'55. u -1 S76K2 732661 ^ 6'916 181839 457 ?51 2 6 a a @ A 16 ix 0 2566 213.333 566.912 336.752 STB. ae^ 3046 u 895849 U104 5 91559E-64 .87 9 6 674452 9^5 9r 161131 5513.14 161411 567.166 37833.1 ?434 lb? 94 d 2E-a3 4 5697E-64 6132195 1? 4149 25 Z 41406 1993.17 a9 .065 143h 3 L54FE-04 143x31 ON let 6019 57 515 49 43510 16498 1651? . w 3

53 014: WK11 a1=4 4 41333E-a4 1 02k 134'17 6 0 d a a 11.x'86 62:1;7T5 16.9 752 2 3 9.7i1c1E-93 2 4 T 40144E-04 134 583 269 166 1*23 97 -15w( 2ki $ 458 2 a a a d a 15.4883 6 x'426 0'61. W Sae, N6 343.200 177 934 :.^Oa V 115849E+@6 W39 d 87145E-6/ 8? 9 0 OWN 1252% + 13,529 i 7543 d6 135 a 503.613 37784 3 8142 1433 y do a 2E-03 4 58997E-04 @403074 1e 6s$ 15 165496 45 13 @e 14287 1 i 9W(E-@4 415002 1084 5 VNI lu 524; i 43: iK i 16661. ^ ibex. ^ 5013 ^l

52, 1777 1:1.,40 13002'@ 5 1?170E-@4 141467 144412 0 @ 0 @ d 141407 03.'55'24 11.2 7 N 2 4 k15u 2 5 7 0/47E4% 158 9 r2 ?17 944 2111 7 e -1 5 842 7544@1 ?1.410:0 3007 450 1 @ a d a a , 01,1 0 2300 191 eel 5de lei ?'S 17 3 177 8 7 :441 54 13e3e1E+% 855167 10639E-633 87 9 0 9D6571 14917 4 1eS 955.1 16579 Sal 7 7 375.`x+. a9c 1664 43 07 @ aE-03 4 58997E-N 06347M a'6 , 115 1. 2m 17 07 140\^ .9 : 5:547E-04 4*5a1 11@5 34 4050 94 ^o 43'.14 4 e 10834 6 517 1103

515'115 19177 INN h 12957E-04 1 06444 1522@0 a 0 @ a @ 166444 0352975 u 5 a 2.5 013547? 2 0 7 4855E-04 176 41 351.841 u:'i. y -151142 745239 4c 4W. 4 456 2 @ a a a 3@ 1229 0 2100 181 eel 506 240 331991 177 ?S1 3124 24 i 51827E+06 9943'24 117541E-03 87 9 0 1 *70 16797 08% 9 10020. S 19850 9 Sad 645 37502.5 885 1^'S0 11 00 a 2E-C 4 50997E-04 Wi 411 X W. a'; 1 :tom'? 4714 47 21 08 1411£ 1 3 53547E-@4 550447 1194 90 43iS 15 53aa. 6e 4,iA 3 10 1-1 7 16907 6 5683. 81

52 2857 ,W934 v.Fgi4 6 37171E-144 1 701 " 163869 0 0 a a a 1.711171 a35c9ee u 8 3 24 ± e @1.AC 2 e 7 400E-114 2aa 1@2 416.164 1'6474 27 -1.57@41 595a^ Z5t^ 5S116i 450 2 a a @ @ 31 4664 a 2116 180 833 566 424 341983 177 576 3113 44 1 8043E+1 1.19105E+06 1 3188E-03 8? 9 0 1 0 , 19914 1 26767 6 13151 2 '0?67 6 499 %8 37716 4 92@4 a'@ T: 32 0d @ 1f-63 4 5699?E -04 a94W 2S IKE6 :5 83o5e9 a'S as 140336 1 3 53547E-04 50 914 12:2 37 + 25 19 S`?3 ? 4 9 17119 1 1114i t+ W 4 51 A9-5 Appendix 9-2 (oont'd)

T9,S1 31 $1.'43 194254 a 8311E-0 1.32454 WN44 0 a 0 a F L 8:451 604624 11.4 1 72 Z 0 ai4b'4S3 3 T 1t38rt3E-A1 238.:14 4Se 4eT 2W ?9 -1 STa42 T78953 7a9i Q%Q 458 2 a a a a a ',,i i r22 a 2110 1T3 833 Sae 427 4M 23 177 574 !^8 08 1 ^142+41^ 1.335S8E+ap L 5718TE-a3 Fl 9 0 129c6 2118 3 u5?^ 8 14887 9 2a^S.`a 8 1 ?iT si:1i: 4 NL' x.'17 t +^ a 2E-C 4 ".91-04 HS*I 25 r21i 25 late o '3i 87 08 0". . e 153547E-04 6047 1272.0i 4tV 31 W94 Sa 428 1 171219 Ma 6 $574.15

51 A 18Sip2 7 Ie2rSE-04 1 ?W54 03049 a a a a a 19 :tS1 a331i23 L: g 2 ds77 ? 7 48?a-* 247 551 445.162 2,7 2a 2 -1.57W 8uh 458 2 a 0 0 a 40 *41 0 a'9e8 171 per 5x0 414 : e9 i92 177 Sap a:L'4 S? 2164 w+ls 1483. E+0p 174541E-a3 87 9 0 1 '?104 ma, 9 1604. 244219 498 894 ?7aa 5 1. *2 2412 53 88 a 2E-a3 4.58997E-04 L''S241 25 e^13 25 0:02,77 514 of 33 08 li?dp 104iE-a4 ON L313. 83 8 5421 ?1 42791 ? 17288 r' IMP a 5502 87

?i8< Z*i 174533 7 70 CE #4 2 0M ill a a 0 a a 2 0?38F d394adS i? c 9 :g 3 2 die8r31 3 2 7 48849E-a4 2ei. e24 523. 24; 27}!? 18 -L 57K, 31e:6 x1,72 1 M%4 4% 2 a a 0 e 0 46 t2t3a a .0 leo. 067 W. 455 `x36 "to 1 77, 545 . 8 2 298:'S•c^ odil9f+d6 18"("E-a3 87 9 a 1.4724: a ils 7 2t+.'S1. 5 8 26254 S 498 594 3 ,r18 11:2$ 2534 L 08 a 2E-113 4 ". '(E-M 14x137 25 51+19 415 Se:m x489 4: 37 as is-9T7 7 353547E-04 a?271x' 1343 43 164 06 54x6.21 4&1 2 17158 8 1443 3509 35

S0 ,h9 8 23342E-414 2. 19935 1806 0 0 0 a 2.199i5 OW943 14 2 A W 3 1 a1"ell 14 7 4WE-04 275 594 551188 28TL1e -15,71)42 8350411 87 1,749 L 23477 458 2 a F 0 0 0 54 V r2 a 1 M iQ 5 W 485 ?93 027 17' 516 25tH ON 2 1380+ed i .^S2SEN5D 2. a4r1E-03 8,7 s d 1 5.17 2x74 29N89 9 V704 7 28x89 g 478 335 3,721? 4 12830 2690 46 0 2E-43 4. 5M?E-04 M3 25 5a77 25 50T691 KA K 41 08 L`^ea 2 3 5 47E-01 7Nk 13^ ?1 _49: .417 ?3 4,N,* 1'LS 1a'Sa7 TO 52

' A9-6 APPENDIX 9-3. SPACE BASED RADAR-A ANALYSIS (Continuation of Appendix 9-2 with i\ . 5 burns)

A AA K AL Be 62 94 e 66 e9 ec eE eF Cl U 01 D OX Fi FA Ft FD FI FS M It IS K9 K2 K5 K6 0 K9 K KE LA LA(FT) LB LC LP LX % IQ 1f fU N NA ND P PC Pi PK PL PN Ri 511 Ti T TA TE TP TQ TR TS TT tM TX V A NA NB MC N1 Mk ww MI N2 Y

51 N6 r ,51 178411 5, Ui X-44 1 41688 4557852 d d 4 a a 1410 ODA64 16 o s 3 did4J21 '. i 7 BiIM-64 123366 14.4N6i 254. d21 -1 57041 640355 93.1227 15748 425.345 2 a a a a 6 14 8316 6 20 !68.333 %1415 255.478 id2 525 3465 a4 67611 4 364111 5. i8641E-0 88 5 a 02W 7:6 046 9579 0 m 599 9579.13 542 A 42845 4 784 _Z ?t?i 63 0 A-C 4 58.997E-04 2 900.03 34 4333 is 29.4131 171 .W 01 lle 167Q, 8 1 SUM-04 332116 678 873 3747 9 3578.41 4x!42.5 11357 a 11486 8 5837.09

41:",0 je 51 'co 1219e 446749E-04 L 08497 a: 1. d t^ a a 0 1x4497 W110 !e o 3 8 7 44788E-04 48.397 4b.794 V. -157x42 695676 93 1227 15748 4%. Sat 2 e d d d a 12. 7 575 a 2240 195 564 107 15 949 1 y N *10 85 4160 ZW31 3. &W-44 87 5 0 4Lkft 44F. 18 11-579 13 4M 21 9579 13 ?8? 881 3*A. 2 7* ?42 :14 a5 0 2E-d3 4. *447E-t4 0107057 14 *t8 10 4 Q%83 10 055 35579 ? 3, 53547E-44 3::11c '77 .5 50,2 44' 4461 23 45840.5 14159 5 14159 7 5959 6

53 1160 1499%3 2571% 183439E-44 1 U349 II&Qe6 a a 0 a 6 182349 a2767ft 16 0 8 8 331KE-+13 ) 1 7 480444 93.944h 187 889 1513.04 -i 5142 711:59 93.1:27 iS7d8 457 517 2 e 0 a 0 0 U. 39:8 a 2430 ms 5N 763 :a7 332 179 ):' 3991 04 779425 449652 5 63539E-04 87 S 0 651267 836.0 8d 1,674 13 4738.89 9579 13 379 199 39152 5 1% 11)7 41 a5 a A-63 4 "FE-a4 0312712 12 1a 2 651:6 ibis )r a9 0 153218 1.53547E-N ,?).le 4117 441 54L f: 4725 11 45x42 9 14997 1 35x07 5856 38

53 5104 167581 246-66 4.49627E-04 1.24194 133m 0 a e a 0 1 20194 &14%75 16 3 7 04 2 1 79735E-0's 1 2 7 4WH4 126 91 253. &2 1832 33 -157042 713494 ^2 6M 218,'07 458 2 0 0 a 0 a 15 7)4) 0 2N 05 833 W. 494 Alw 129 378 9a* 74 1 61e.' 4 Q4182 7 8438E-04 87 S 0 814436 11345.1 1211717 6733.8 12171. 7 Pe 4as 38784 5 7434 1418 77 % 0 2E-63 4.58997E-04 04%9 u. 7" 1e L ?M' 2754 a2 13 !16 15464 6 3.53547E-04 Im 933 87) M 49 4858 y 44579 9 0" 1 15478 41 5795 38 A9-7 Appendix 9-3 (cont'-'.)_:)

2711 131154 2248:8 5.'8Y?5E-a4 1.19295 14', 7k2 a 8 d 6 a i 39295 a3u799 18 1 7 2$ 2 2 dii34S3 1 3 7 48644E-44 154.831 W. 7N 2kb 13 -1 SAW 71458 92 . Not a 2782 458 2 6 8 a a 8 21 WZ 6 2179 189 167 W 444 318 217 178 5% 3098 57 i L44e%+k% 380617 9.9801E-a4 87 S 8 QU171 148333.9 Aft 6 ati8a a3 149916 359 1182 38395 2 7798 1149 34 87 a A-63 4.58997E-84 INW6 it 33 !a i. U-Q 317182 it 87 14789 5 3.53547E-81 456815 162153 4733.01 0" 7 8 4412 4 15W o ON 4 5-131' 21 53 13142 !J42.55 111135 5 9434 t-i+4 i. 5^^870 152172 d a a a a 1. S Sd7t 0344775 18 9 7 52 Q9F I 2. 4 7 WA-04 179172 359.3345 2255.75 -1.57842 '1 SS .1.3379 182233 459 2 8 8 a r, @ 27 4723 4 2290 183.333 585 :4 3 312. 1`3 112 35'7 dti 1 52933E+8ti 981 ti 11812E-a3 87 S d i 87403 1053 1 1!39 5 16335 i 13839.5 M. 11i31 37182.1 1142 1 x 1.43 t^8 8 2E-03 4.58997E-a4 9773421 .11 KA is i 13 391 4648.55 21 08 14502.9 3.53547E-84 515514 11'4 1 4445 it 51" 34 4.`115. 7 114:4. 3 16454 4 56;3 54

:+1:.4 isles 1 11E-aa 1. ?1217 151 22 N y '3 ; a 1.31oY' 3?N?31 11. '0 .+ ? ; ^l^ aa4 4lt564E-a4 193. 122 a? 245 " INK ?1 +1+9 +::.5 ^1 45S a a a H y 15 't?k 8 2128 171; `ir 5Y 11. '? 177'42 :^43 19 1 "^'Srr¢'r30 : 1449`t"* i 687:3t-4'1 $7 5 •3 1 17 41 18878 3 21:15 12525 ; ?isis 9 147 2'S '4 a! $ ?4* :841.1: '3" a 2E-i33 4. 589.7E-84 0437383 18. 3471 18 k,a5 5:x'7: as 25 14216.4 3.5.547E-@4 535:81 1241 At 4123 IF `: ^ 1 41 2 9 10 d7 i 10975 7 SQ. N k X115 1."41 La4259 5.,440E-04 1.45471 117etA 9 a 8 a a i. '.4547'3 8'372682 115 ? . 5 X1591 2. 7 4MIE-94 2S.234 4% 467 2irA 11 -1 57842 '^1a0 9►31300 55113 453 2 a a a y 6 37 -742 0 2110 175 Qs's 45x 121 1.1 i' t' 1s:'S 0s 1. 99d30c+i+v I "A4E +06 1.57107E-0s K7 ^ y 1.:+•:: 21875 8 24139 14710 1 M419 344 054 '371'18 8 ~ `241 a a. 2E-83 ,539. E-84 ii*07 18 7193 18 '11:*4 1592 6 29 00 1469: 4 3 5354 7E-84 1l^3s1 1 7^ 71 4rw .i 544( 3 46if 7 17101: 17133 0 550j 9 r )2 @482 AWN 131584 7 51?23E-84 2. OWN 172MM a a y 8 8 2 06839 838-1'795 12.1 8 4+? 2, 810:417 2.7 7 48732E-84 242. '18 485 535 :182.41 -1 57842 '45025 43 13x7 193655 458 1 8 a a 8 8 41 3319 8 Me 178 5R• 412 _12 42 177 It'Q :1^?/ ^!• 2 12054E +Nt 1 451E+80 1 12846E*@2 87 5 y 1`3138 __S^s.. 2412Sa 0 101.71 922 1'319 5 9558 2 e 70 89 8 x-+33 4 589.7E-84 124743 10 6344 to 0-,:54 grWE '484 57 b9 14483 S 1: 1 4 1: v:' 5183 " 4 42 1'184 171% 5 55% 48

A9-8 Appeadix 9-3 (oait'd)

X 5513 224318 thow 3 $1342£-44 2 ISO 178133 1 8 4 1 1 215635 Moll 12 7 8, 96 2 9 1175451 2 9 7 48885£-44 2%419 512 M 2752 dd -1.57812 162745

I 89.2832 8619x'9 4% 2 8 1 1 0 6 x.5247 a lab$ 155 Sd6. 4•^i 344145 171372 2943. a7 2 26257E+k 1 579a")6 1 "7E-W . 81 S 8 1 4c264 12409 LN44 1 11681 7 .'E445. 3 A 184 ;7:132 S 1. a2 .'S3a a2 89 8 2E-W 4.58947E-04 139968 It %U i8 56W 8398. 88 37 $9 ton 4 1 S3541E-44 061 12M 34 NO ?5 53931 42876.1 17121 a 17228, 5 5Sn 94

51 386 :3$9:8 IaQI5 a 1AI M44 223337 1832% a a a $ a 2 23331 8448945 13 a. 3 ai81'1 1 7 48813E-,N 275 544 551188 :874 a6 -i S ? ad2 8~` SrL4 ?!s4.'S.q 458 2 a a a a Q 34 0? 2 a 058 162 5 A 4"i :49 ;r7 17 574 ,,.'4. 2 i /;^ Este+ i ?1045E+tf + : W4H3 $7 NS a i . 042 ? x'44 7 L:3^t / ;7414 ? 338 IV :'^4 s 1 842 N.Q6 4 0 d ZE-43 4, .?" E-114 155881 16.51 ld 51m 9386 22 41 0 14x63 3 1.5:i547E-d4 110328 1:714 34:2 i5 3y4 43 42879 S 17111.5 V248 1 $574 21

A9-9 APPENDDC 9-4. SPACE BASED RADAR-A ANALYSIS (Continuation of Appeneix 9-2 - with N = 2 burns)

A ►ip Ai k 8a 62 84 BS 86 B9 8C ec 9F cl U Ul D Dn Ft FA FC FD F1 F: M li IS K9 K2 K5 K6 K7 K9 K KE LA LRkFT! L8 LC LP Ui M1 M2 MF MU M 141 P PC Pi PK PL PN Ri SM T1 T TA TE TP TO TR TS TT TW TK V A IN M8 MC

53.760 0656893 147725 6.54917E-a1 1 64762 0521377 a 4 9 6 9 1.64762 0361076 10 6 8 2 0!26139 2.1 7.93928E-04 5.96491 113697 221549 -1.5704 W17 93.1:17 15798 424 767 2 b a 4 a 0 26.2396 a 1944 153.3333 550.322 23Z ',12 183.678 3955.9? 54691 23993 4.72425E-85 88 2 0 124$47 506.3334 9579 13 257. 09 x579 13 449.36 440.1 . MAI 215 41 05 a 2E-93 4509.7E-84 2 652.1E-83 37 6869 5 32.6869 1591.3 91 055 177913 3.53547E-44 lla16 611893 ?104 72 ?119.35 58097.9 99e U 9913.74 6911 74

118494 203575 4 95839E-44 126593 .9940489 9 0 4 4 a 1.26593 43134x1 1a 0.8 2 9 w4a95E-a3 2 1 7 'WK-* 41.23377 82.4754 844 5?9 -1.57041 i^o40d1 931227 15704 448.403 2 0 0 a a 4 16 22 9 AR 188 502.414 273 184 181581 347 51 337018 1845?'r 2. "MME-44 88 2 ^? 4Q0591 iL % 4579 13 LK8 r' 953 13 M W 417t 3 7 ? ^3 i^s8 95 0 2E-43 4.58997E-44 01529'31 1@ 3017 5 5. 701 911184 NS 055 109419 3.53547E-a4 .115 717.37 42A 42 3453.44 474`4 4 !1549. 0 11593.0 5644 95

51 . 70i 146977 228817 4 4577E-94 115437 110655 6 M 4 N 4 115437 tlE%"7 W r 8 2 9 1INW-43 2.1 7 72317E-04 ?3.4315 166.863 1375 34 -157W 0i^7501 93 1227 15144 457 doe 2 0 y 0 13 0115 0 2290 L44 833 5as '333 2$4 Ali 18+? M,7 347? 73 671478 .4 . 5 MA-* 98 1219.69 ?579 L 4056.73 9579 L 40 +63 4074A 8 783 1^+H4 a5 0 2E-03 4 59997E-+4 kli;8^3 77 :-`:41? 5 2 :418 171: 33 It,4 ? ,5 ii5:.. .. 53547E-04 3_416 "i 544 4:17 58 46291 4 137@'1. i 137'33 9 5558 57

53 5184 165658 224838 4.60603E-94 1.19L4 1311483 4 4 a a g I M9 9383314 18 3 7 44 1 9 41152E-43 2.1 772238E-91 118 416 2336 833 17N 74 -1.57841 00046 44 0939 181839 4% 2 a 0 0 4 0 14 5383 9 We 189 107 54'3.362 k9x 064 108 038 _Ki 5e 933521 5407K 'r 57RW_-a4 8@ 2 e 745'42 1*4 1814333 1 1433 86 7014? 1 207 334 WIN i ?4:4 1.1 78 42 e 9 2E-33 4 58437E-94 441798 ' 06511 5 I. !6511 2587 88 L; ! 55 16146 1 1 5'_54'E-94 ^4_ 7'8 083 473; ft 4320 11 409 8 13749 2 14481 1 M 17

A9-10 [! Appendix 9-4 (cont'd)

53 Q42 1'9543 21:109 5.12344E-04 L 36866 142564 x 6 6 a 0 13" 032fte7 it, 9 7 52 13 aiu&l 2.4 7.49a0SE-04 146.774 293.549 1976 21 -1.57042 12x73 x 8379 .296678 458 2 0 a e 0 6 20 32% a 221e 194.167 %4.447 Al 427 179.553 3556.9 L 1915SE+d6 ?M. 9.64621E-01 aT 2 a 93%1 12965 7 13529 6 7839.9e 13529 6 191101 3.9Se3 .8442 1626. 48 as 8 1E-03 4.59997E-04 .85N17S 6.44229 5 1. 442229 3494.5 17 a6 M. 57 8 3.53547E-64 415062 417 331 44% u 4598.2 454x5.7 14594 3 14671.7 5901.74

52.777"7 L41476 205464 5 ?097 -x4 154141 Isul 5 0 d 6 a 0 1.54241 03:`M99 U. 2 7 76 2. I 01;'%606 2 5 7 48149E-04 174 *6 349 611 2:'x4 67 -157x+2 ?"^'S ?1.1109 ? 7 458 2 A a a 9 9 6'3..57 0 diix 173+3 83? Say 124 A9. e9 173.676 :415 41 L 44W- ..+a6 914121 i. I"; "*A- 3 87 2 9 L X411 15785.5 1657 9 10000 1 1657? 183 Z 3 `* 7 34 ^ IN4 41 x L b7 a 2LZ`3 4 5^?9N-a4 Ni 41.27 Q IIS7? 5 S, 1N17? 44Z 6 21 6i 1.1tt• S :. 53547E-!14 tit 1142. CZ 1? 487.0 04 44% 15457 15404 2 5790 33

52. x:15 21110$ Ls; So 0. 44.41E-04 1 72248 1011+1: a a N 0 1. '2248 63559x'2 u. g _ 5 @14112'1 2.0 1 4131x6-64 IN 501 401601 23`33.3:+ -1.57642 7'A714 ?0 9810 4974 45.3 2 0 0 9 0 • 2* y 2138 177.5 SOS. M :14 :77 1 `3 107 332152 1. 7ast+al, 1 u348E+06 L 36721E-03 .87 ? 0 118115 187':9 4 Lab% 9 114'93 i I` e%. 9 177, 56 3IN." .1 MI 59 0 -d3 4 58991E-04 0918179 5.8767 5 3707N3 SS09 a8 25 08 14645.5 .1 53547E-04 5%"F 112.45 4107 11 5155 r5 1;0:0 4 16A3.6 103;0 i 56'4 A

52.2857 .269981 181584 741 E-a4 197919 165M 6 a 0 0 9 197319 9384871 11. 8 8.24 Z e 9101u9 17 1 48418E-04 W. 341 126 083 2463.08 -1.57 042 7:LV55 1598 021187 458 2 x d 0 a 0 42.6877 x 291® i i W 7.:'3 '121. 41.25 177 :'211 6[ 1.8559+60 L 25756E+06 151895E-03 87 C. 9 1 27'9 294W 0 23363 © 13886 6 2336: 6 174 011 .3623% 9104 ItQ 7 5 09 6 2E-03 4.58997E-94 106146 5 '377 S i 370 ^8 04W 75 29 09 14331.9 7. 53547E3 -04 K468 1274 35 3812 4^ 5278 3 13247 i 10751 ? 16827 3 5022.13

52.2857 218074 181%4 7 41629E -04 L 97918 17199% 6 x 6 a a 19%(88 0384871 118 S. 24 1. e !iS=11 Z 7 7 48416E-44 142 768 485 535 2682 41 -1.57042 7 EK155 9Q 5588 021187 458 2 0 0 0 8 0 42.0977 0 2948 179 5% lit :11 495 7.77 %3 3^1. 0t 2 106^^^1* i. 43101E+06 1 r184tiE-03 87 1 9 1 fsi^ 2=197 13X a IS &I 233363 0 ii 1. 472 :7015 .02x4 1396 '• e9 _ a 2E-03 4. W.("E44 12-Wes 5 t*AM 5 04819 :R 5 .. NV 1^411N 5,1543E-•.14 `•.'41.16 12'4 5 :x.11. 49 ^ Yi p : I:&.4. 1 SNI .4 SWJJ ^ 500

A9-11 F

Apprndix 9-4 (cunt,d)

s2 2442v19 loies5a a 67267E-a1 2 31752 17621e 8 e a 0 a 2.31752 0416477 12.1 8.46 2 7 2488656 2 9 7 481 A44 241113 462.226 2615 02 -157211 717407 98.1337 77296 456 2 0 8 8 8 8 58.4164 8 1928 168 Sa6.16 312 966 W 241 3106.21 2 LAVE#% 14951E+8b L 82397E-03 87 2 8 1 43:69 23613.4 27899.6 !e'S87 6 27099 6 169.487 37088 6 9SS8 2184.99 1 a 2E-03 4, W. 13813 5.57719 S 57719 8287 79 67 .1 14'6619 1 S3547E-0 78328 1356.3 6'68+6. ai 5247 15 432LQ 1 167x8.9 16804.2 5618 48

52 a1K . 230X46 162530 8.68782E-04 2. 182667 a 0 0 0 8 2 x986 0414917 12.1 8,48 2 7 al$7LQi 2.8 7.4RM-04 269.97 539 44 2823.25 -1571+12 7L%6 48.1337 .77295 458 2 0 N a y a S7 SU 8 1930 lox 833 `ia 218 495 L'7 752 31* 75 238866 +06 i 67715E+6b 2 Ohil.(-03 87 2 6412.6 278:°9 6 107.4 27099. i 168 483 375021 4558 267137 1 a 2E-K 4.5 7E-04 1537: 517-4 5 ?44 15 41 .1 14346.3 3. S3547E-a1 7KZi I:,F 33 4Z. 3$ 13185.8 16&.2 iti9L. 3 S6a3 r-6

I , A9-32 OPTOTY/M - i PAV nl l OTY. 1 9utTLE

81 B BP CD D2 DD DR DT DY E EE EM F9 0 K1 KP KQ KS KT LD LM LQ MI NV P1 RH RM RS Si SP T2 T6 TD TF TO TM TN TY TZ WL W5 WT

198 1 1 814 1 1 2 61 1OIE+O 187 0" 3.218 .417 .2 21 S 15 4.1 18 1888 584 d 95 114159 .898 852 88 35888 3338 .825 -a3 SE-43 .94 9E-03 85 .81 41 68888 1

A AA AC AL Be 92 B4 B$ E06 N K BE OF ti C2 Df D DX Ft FA FC FD F1 FS H 11 15 K9 K2 K5 K6 r.7 0 K KE LA LA(FT) LB LL ! LX Ml Mf IF }u N NA ND P K Pi PK PL PW P.i sm Ti T TA TE TP TO TP TS TT TW TX V Wi Wh WB W[ WN WX WW WY WZ

51766 *bW 159178 276 9-0 738&3 ^ '4!6 a 0 0 8 0 73861? 8235287 18 6.8 2 4 41 ? 49.9SE-94 6.44887 U "i7 L'I 768 -157842 W462 93.1227 1588 426.281 2 8 8 0 8 9 5.92343 8 1918 159.167 583.298 141%6 198 792 3978.74 66182.5 33184.7 4.98398E-85 98 9 0 10.5 '18.'83 5416.59 35'3 495 986134 6k 442 40.0 S 7% x. :46 aS 9 2E-63 4.58997E-94 3.18 9E-93 58. 3243 25 1.:1:4 1."1-195 ai 055 15267 7 3.53541E-84 516624 9% 752 Kle. A 4267.58 4045.1 13549 136.1 6886.72

53. 66 118311 2904 21361E-84 593982 .8939835 0 d 8 8 a .593982 OM93 18 6.8 2 5.83816E-83 2.1 7. i9346E-04 48.8567 817115 8"38.84 -L 57844 e58589 93.117 15788 451337 2 0 0 8 6 a 4 1517 8 2558 179.167 585.499 271919 178 58i X96.66 MAO 221842 2 76C+19E-04 87 9 9 4036. id 5416.59 2-173 67 9061. N 516.488 38334.1 7% 781.51 05 8 2E-93 4. WE-04 8188549 29.4365 25 4 43648 1883.2!9 85 055 11155. 3 3.53547E-04 5!6624 1110 74 4245 t; 5021.49 44862.2 15937 8 wa 6 6388.88 A9-13 Appendix 9-5 (cont'd)

S3. Si84 141651 183369 2.55285E-04 711245 IL429 1 0 d a 0 711245 .8225958 16.3 7 04 2.1 6 6365%-di 2 : ?.1.853£-64 66. 8681 133.72 1143.97 -157x44 .654% 92.6938 .21&'67 457.754 2 0 6 8 0 6 198!96 6 2958 !76.633 544.946 269.664 178.054 3846.22 612765 367412 4 73799E-04 .87 9 0 .581488 6618.64 7541.24 4179.49 115!3.7 547.06 37831.3 7434 11446 62 16 6 2E-63 4.56997E-61 81332421 27.4122 25 2 4m 1994 '% 09 .06 14371.9 153547E-84 694927 3'48.1 3568.06 5281. Li 434913 16567.7 !6517.6 6338.13

53.2711 .156392 159174 3.12156E-84 .869992 124128 8 9 0 / 8 .069992 1249662 16.6 7.28 12 7.38642E-3 2 3 7.17666E144 03.8353 167.67! 18.81 -1.57844 6418 42 2656 .292M 458 2 0 0 0 0 0 s 95194 6 1911 Is. 16? 586.828 26102 177.972 3702.84 ?M% 517338 6.37318E-64 .8? 9 8 784936 8114.75 18281.7 5667.06 141812 583.636 37746.3 77% !227 92 87 8 2E-83 4. 58997E-04 .04064 26 is" 25 1. ;54'93 2@92 98 .13 97 14287 7 153547E-d1 ?8092? 12.11 a1 ?350.83 x:4 :8 C3-96. S 16613.5 166819 6328 72

53 0242 IM22 1"W 3.2241E-64 899042 2151*4 0 0 a a a 899942 6253933 18.4 752 2 3 7.6:?ObE-43 Z 3 7 17231E-91 187 347 214.694 1523 73 -157%4 654.916 91.8374 334458 456 2 8 8 8 0 4 9.5729 a i89/ 157.5 586.161 268 044 177.839 x68. A 41.9452 674187 8.24947E-44 37 9 9 842'_45 INK $ 11868.1 7343.23 149314 501.778 37599 2 8142 1454 24 0T a 2E-83 4.58997E-AI 4634689 26.2541 25 125611 38E+8.14 17 07 14283 3.5:347E-04 ?2:797 1367 40 x'411.42 52'17 61.1 43217 S 1678?. 5 168.46 6-'85.67

52 7777 .182212 Ism 3.32772E-84 .928373 1440 0 ^+ 8 8 8 .928373 1257981 112 776 2.4 s 7 P"K-83 2. 7 16892E-84 129.013 259.62? 17 09 -157%4 X61757 .14109 Im 450 1 0 0 0 d 8 iS. 204 6 1871 05. W 506.287 274 5 11*7 713 :443.38 1.1o70sE 5.'w95 1.49627E-03 87 0 463363 1:103.1 Ll*," 2 : W^ l 43 !5602 9 5 00.6:1 :7458.4 +34% S49 27 47 0 -63 4. 56997E-94 010% 00 ^"^ i. ^•u'< 4733.73 21 07 14117 2 ?. SISM-44 747571 1',417 Al It 5;.8.62 43855.7 16.044.3 16999.2 629126

52.5315 W24 149335 3.43256E-EM 950827 152727 a e 0 8 8 958627 0262614 11.5 8 2 5 814541E-03 2. S 7.16509E-% 151252 362.58.'3 !848.24 -157644 6"71531 4a w?k 4ms 438 2 / 0 0 0 a 18.0898 8 16% 154 167 W. 468 281 :6 17 7 !&1 33325.3 1 MA496 1839 i 18740E-93 87 9 0 Y 1 97401 15 6 iS515. 3 1*14 1 16435.6 49r b53 s7:^^ ^^ ,y^ d3 07 0 A-03 450947E-M 8944033 25.8376 •93'616 `.! s`-' ` a7 1 s^1 a S 5CE-44 71, L45 14,* :+Q 62 %97 31 429912 17048 9 1,7145 9 Q77 5 A9-14 Appendix 9-5 (cont'd)

516462 . W.474 139916 3.71296E-64 i 83617 156719 6 0 6 0 6 163617 AMC 12.1 8.48 2.1 9 81693E-83 17 7,16499E-64 164.2% in 512 1982.37 -157944 WWII 99.1337 .541965 45e 2 6 9 6 0 9 11742 6 17"9 149.167 566.445 191682 177 565 UP' 96 153461H96 1 iftm4s 13326E-93 .87 9 6 1 IU16 17625 9 17944.3 !?236.9 17944.3 499.312 37287.9 9558 20L2 % 87 9 2E V 4.58997E-94 1847:+9 25. .12 25 71.w? 81 6567.97 .29 .97 0", 6 3. 5.154M-04 018843 140; tic 641 in %%2 42653.9 17146.1 112317 627129

515x93 296415 138599 4.01672E-64 L JIM 163832 9 6 6 6 6 i ltm I M= 11 7 9.96 11.9 9.52611E-83 2.9 7.16486E.-64 174.5% 349.183 2646.22 -L 5,7944 649184 89 M2 .6784^6 4% 2 6 6 0 a 9 14.9756 9 1736 144.16' 596.451 1p it 177 5+9 X13. $t 1649M+K i MON i 4M-93 .87 9 a :. 2:7`. 184718 1%% 7 Om 19456 7 499 °09 31276.6 `e 07 a 1-03 OWE-% 124477 25.6:4 15 c ++ 'x448.65 33 .07 13996.1 3.S'5^.47E-04 6Ee441 + :!8.44 X749 @Q ;484 ?4 42846.9 17153.1 17227.4 621167

51 6 ^3L°? I'''bi66 4.11814E-94 1 %346 1c92114 ) 0 e a a 116346 0M726 C ..1 ;. ibY98E-4M 18.9. 024 179.047 2!37.:5 -L 37$44 -:8 L° 742& 458 2 9 0 a 8 d 10 N44 0 ill % 141.667 566.442 3$1 638 177 558 x'824 L 1. '9219466 i 32407Hd6 161440E-03 87 9 9 1 :2+451 201'14.1 14003 21Q'14.1 496.597 31:87 4 07 6 2E-03 4.58997E-44 14W 15.5645 25 !-071 8403.82 37 87 13917.7 3.53517E-44 890214 15L:. 36 2654 51 `,466.72 42856.6 171414 I7K 9 6273.71

51 *19 214.134 166196 5.12672E-64 14314 170156 a 8 a 8 8 14312'5 93 VI, 1s, 4 44 1 ft2l1a 3.2 7 16397E-94 176.36 3,52 76 1983.1 -1 51044 : liei 08.4392 935912 456 2 0 0 a 0 24 3191 0 1". IA 566 4Eh M. 403 1" 511 ti:? 05 L 119?4E^66 L ?35046 1. E4WE-03 87 9 d i :"5 8 1978x. 7 23968.2 15104 3 ?69.2 490 '?5 :7234 9 L 0:'74 -22:15 ee a 2E-Q 4.58997E-61 I'nQ 25.5p75 25 5873%"e 9 : i 9 41 08 13969 1 :. L 01927 ll!* 07 K: 5 `.4V 42'48. 7 172613 17:,58 7 (IV 36

A9-15 APPENDIX 9-6. SPACE BASED RADAR-A ANALYSIS (Baseline with Truss End Weight WT Changed to 1000 Pounds)

OPTOTY/OGA - i PAYLOAD, i OTY, i SHUTTLE

81 B BP CD D2 DD DM DT DY E EE EM FB G Ki LD LM LO !H MV P1 FM FM RS Si SP T2 T8 TD TF TG TM TN TY TZ WL WS WT

i88 i i .814 1 .5 2 3 81 4E+87 .187 388888 3.218 32.17 .2 46 388 684 .6 .85 3.14159 .863 .8513 88 37888 3338 .825 2E-83 .81 .84 9E-63 .85 84 At .82 13E-84 68888 1888

A AA AC f4. N B2 84 Cfi 16 89 BC BE BF 1:i C2 Di D DX Fi FR FC FD Fl FS H li IS K8 K2 K5 K.6 K7 K9 K KE LA LRTT) LB LC LP LX rd M2 ff mu N NR it) P PC Pi PK PW Ri SM Ti T TA TE TP TT TW TX Y Wi WR W8 WC WH WX WW WY WZ

51766 .8568883 56785.8 .8376414 118266 .8444475 8 8 8 8 8 1 18288 .2?4377 18 6.8 2 .8m 2 5 .8236529 2 29735 4.59469 99.7826 -155897 .474789 93.1227 .15788 426.388 2 8 8 8 8 8 118.331 8 1146 95 583.425 144.18 188.575 3988.46 41945.5 3". 8 2 92698E-85 .88 9 6 .8987883 458.433 9579. i3 323.625 9579.13 WON 9 .788 156.884 .85 8 2E-83 4.58997E-84 122685E-83 193.611 .81 .86 15263 2 '. 53547E-84 332116 378.612 1193.7 4354.87 46188.6 13819.4 14222 2 6864.16

53.766 .185857 163491 .8228962 194686 .8848187 8 8 8 8 8 194619 .21822 18 6.8 2 522188 2.5 .8227876 28.9618 41.9236 488.829 -155%4 .545165 93.1227 .15788 451648 2 8 6 8 9 8 44.7558 8 1548 128.333 585.87? 194.77 178.123 3999.85 384163 239636 1977E-84 .87 9 8 .324394 3266.26 9579.13 2476.7 9579.0 ' 37913.9 788 561332 .05 8 2E-83 4.58997E-84 .8184851 1184. '^ .85 .86 14453.3 153547E-84 .332116 511456 2178.35 5173.66 43579.2 16428.8 16538.9 6337.88

0- . A9-16 Appendix 9-6 (cont'd)

51766 .138961 114511 .9284676 low .183944 8 6 6 9 8 1 98964 .282828 16 6 8 2 . 4S3989 2 5 .8226675 41.7531 83.5863 7710 -155946 .559383 93.07 .15798 457.889 2 9 9 8 6 9 38.4473 6 ir^4 135 586.325 284.889 177.675 4983.39 579646 449111 ? 146E-94 .87 9 8 .496888 6224.54 9579.13 4812 95 „79..3 37416.5 .768 859.993 .85 9 2E-63 4.59997E-% .8338752 2832.51 .99 .66 14367.4 3 53547E-84 . M6 530. 027 ANA 65 535?. 21 436th'. it9R't. 5 0=9 6296.97

51766 .148839 114511 .8284876 189652 .117498 8 8 6 9 8 188864 .282928 i9 6.8 2 A83989 2 5 .8226675 68.318! 12F_ 62 986.9% -L 55946 .559393 93. !227 .15799 458 2 6 9 8 9 9 38.4473 9 1628 135 586.325 294.888 177.675 4993.39 8247% 647272 5.49722E-94 .87 9 6 .635146 8856.99 9579.13 69%74 9579. i3 374% 5 .788 1999.78 .95 9 2E-93 4.58997E-94 .9489399 2499 .13 .96 14283.7 15354it-94 .332116 5X 927 2419.55 5353 81 43887.5 i6992 5 17192 6 6296.97

52 5315 .159176 193481 .9248387 212112 .126338 8 9 8 8 8 2.12125 .219229 it 5 8 2 5 .5M2 2 5 .806575 712781 142 54 1883.96 -155947 .662444 99.9846 .386796 458 2 6 6 8 8 9 53.4853 9 1549 128.333 596:454 234.526 177.546 3325.6 i 98A89E4% 896178 6.72179E-84 .87 9 9 .734576 11889.7 12419.5 8869.52 12419.5 37273.3 .885 1272 22 .95 9 1-03 4.58997E-84 .8643923 2499 . V .86 14197.8 3.53547E-94 .386698 595,5% 2178.35 5485.66 42842 9 17157.1 M45.5 6272 31

51386 .167782 93988 6 .827837 Z 45775 !33169 8 8 9 9 8 Z 45791 .255954 13 9.2 3 .661127 3.5 .8226514 79.138! 158 26 11411 -155947 .642281 88.8619 .5344 458 2 8 8 9 8 8 717486 9 1969 121667 586.56 261882 V7.464 3.82 114662E+96 944849 7.87295E-94 .87 9 8 .216395 12983.4 15263.7 19623.8 15263.7 371818 1862 1414.17 .85 9 1-6 4.56897E-94 .8798639 2499 .21 .96 14111.9 153547E-94 .441282 644.271 1957.91 5438.82 42737.6 17262.4 17449.5 6262.94

58.9893 .176183 87982 .9396513 Z 6997 .139837 8 9 8 8 8 Z 69987 .247189 14.5 18.4 3.5 .7M3 3.5 .9226321 89.1114 178.223 1223.91 -155948 .699969 96.7547 .841$49 458 2 8 8 8 8 8 86.7882 8 1428 118.333 586.826 294.833 177.174 244192 132281E+86 118598E+86 9.11473E-94 .87 9 6 .999471 15247.7 18139.3 12748.3 18138.3 36868 1239 1568.99 .85 8 1-03 4.58997E-84 .8964168 2499 .25 .86 14819.9 153547E-84 .495964 794.128 i$52.89 5555.35 42367.8 176322 17665. i 6238.81 A9-17 Appendix 9-6 ( coat' d)

58 8893 .185119 87982 .8386513 2 6997 . !4M 8 8 8 8 6 Z6998? .247189 14.5 it 4 15 .72683 3.5 .8226321 181369 286.738 049.29 -155948 .698869 86.7547 .841819 458 2 8 8 6 8 6 86.7882 8 1428 118.333 586.826 294.833 171.174 244182 152823E+86 128293E+86 186731E-83 .87 9 A .994464 17616 6 18!38 3 14788 !8!3813 36868 1239 1723.28 .85 8 2E-83 4.58997E-M .111843 2499 .29 .86 139914 153547E-84 .496864 784.128 1852 89 555.35 42367.8 MM 2 17713.1 6238.61!

48.8815 189667 78347.8 .83477 181248 .158459 8 8 8 8 8 3.87267 .263784 16 116 4 .827186 4.5 .8226322 !84.747 289.493 1334.12 -155948 .723932 84.6628 i 2x.664 458 6 8 8 8 8 112 W 8 1348 iii 667 5K 824 317.993 177.176 2!35.72 168WE+86 137251E+86 i 136E-83 .87 9 8 184381 18x96.5 2im 8 162115 21885.8 36862 5 1416 1887.58 .85 8 2E-83 4.58997E-14 A2 2M 2499 .33 1 k* 13984.9 3.53547E-84 5`8447 737.6 1649.28 5554.45 427%. 176=x.? 1772.: 6238.27

47.63;'6 i?36 69??? 1 .9294652x. 4969 .15303 8 8 +4 8 3.48713 . M146 17.5 +? 9 4.5 .938634 4.5 .822EM, 183.839 287.678 292. 69 -1 55908 'S_Sri^i 82 W 1 78924 458 2 8 8 8 d 8 !44.686 8 1268 185 586. , , 338.07 V7. s" 5 1881. P4 L 66871E+86 143874E+86 i 1698E-83 .87 9 8 188856 X818> . 8 23887.5 17421 23887.5 36916.6 1593 1872.81 .85 8 M-83 4.58997E-84 .142348 2499 .37 .86 13"976.5 153547E-84 .685831 762138 1458.23 5534.84 42432.9 17567.1 17743.8 035.81

47.6826 iw9652 692721 .8394652 14869 .158464 8 8 8 8 8 3.48713 .288946 17.5 12.8 4.5 .938634 4.5 .8226353 !15.865 238.129 13%.97 -1.55.48 .75,599 8? 5862 178924 458 2 8 8 8 8 8 144.686 8 11''68 185 `96.775 328.237 177.125 1887.84 1833683E+86 159428E+86 i 32639E-83 .87 9 8 115742 222317 23887.5 19384.4 23887.5 36916.6 i 593 2886.35 .85 8 1-83 4.58997E-84 .157736 2499 .41 .86 13967.9 3.53547E-84 .686831 762 338 1023 5534.84 42432 9 17567.1 17759.3 6235.81

A9-18 .APPENDIX 9-7. SPACE BASED RADAR-A ANALYSIS (Baseline with IS = 450 Seconds - Held Constant)

OPTOTYM - i PfKm 1 OTY, 19l u

81 8 w CD D2 m DM DT DY E EE EM F8 6 Kt KP KO KS KT LD LM LC MH N PI FM FM RS S1 SP T2 T8 TD TF m TM TM TV T2 IL WS WT

188 1 1 .814 1 .5 2 3 .81 4E+07 .187 388898 1218 3217 .2 21 .5 15 4.1 40 198 684 .6 .85 114159 .863 .8513 Be 37888 3339 .0 2E-63 .81 .84 9E-83 .85 .81 .41 .92 13E-04 68898 1

A AA AC AL Be 82 84 B5 86 89 BC BE BF Ci C2 Di D DX Fi FA FC FD F1 FS H It IS K8 K2 K5 K6 K7 K9 K KE LA LATT) LB LC LP LX Mi M2 MF 191 H NA tD P PC P3 PK PL PH RI SM Ti T TA TE TP To TR TS TT TH TX V Ili 141 18 10

51766 .8711696 21A918 18136E-84 185866 .8564873 8 0 8 8 d 185856 .8276174 19 6.8 2 8.81627E-03 2 5 7.2681E-84 9.9% 19.3588 3381672 -157843 .6956(6 911227 .ONO 458 2 6 8 0 8 6 i2 7575 6 2346 0 WAR X5.949 179. $93 3905.85 95316.8 X4386.2 6 88802E-85 .87 9 8 .146536 i02156 9579.13 M 643 9579.13 689 clot 39881.2 .788 253.469 05 8 2E-03 4.50997E-04 3.35313E-633 49.77A6 25 24.7786 281.188 .01 .06 15248.9 153547E-04 .332116 777.15 5829.42 446123 45M 5 14159.5 14477.2 6539.22

51766 .124411 2x787 139095E-04 .943684 . NN45i 0 6 0 0 6 .9~3684 .8268693 18 6.8 2 7 9036E-03 2 5 7.28i67E-04 55.2417 118.463 1077.2! -157844 .723539 911227 . 15780 450 2 0 8 8 0 0 1A. 4615 0 2500 281333 585.285 316.185 178.715 X95.16 483780 263518 128941E-04 .87 9 0 .448278 5195.17 9579.13 2829.8 9579.13 556.702 38572.7 700 7?1986 85 8 2E-83 4.50997E-EM 1 0178561 29.. W 25 4.5004'3 Wt. 36 .85 06 14456.4 3.53547E-04 332116 838.299 5740.71 4435.86 44336.5 15663.5 15924 1 W 3 S A9-19 Appendix 9-7 (cont'd)

51166 .152141 28lSO4 3.38762E-04 .91''8825 .128755 a 8 8 0 a 928825 .025128! l8 618 2 1.79575E-03 2 5 1.19829E-84 !82 642 205 265 1634.05 -157044 .730193 91 iw .15188 4% 2 0 0 0 0 0 9.97496 8 2540 211.667 585.587 321244 118.413 3997.56 811791 494372 5 86937E-84 .87 9 8 .678926 936175 9519.13 5388.83 9 9.13 587.225 38235.3 .7% U6183 .05 6 2E-83 4.58 MAN . 8326456 21.4181 3 2 44811 !968` 73 .09 .06 14374.2 3.53547E-84 .332116 841 M 5925.88 5066, 93 43%9.8 16859.2 16828 2 6374 9?

S2 5315 .169229 253534 3 88488E-04 188963 .134317 8 8 8 8 9 1 am .0278741 115 8 2 5 9.15699E-83 25 7. OW-4K 134.583 269.166 1923.97 -157844 .79x628 90.9846 .306796 459 2 8 6 8 8 8 13.7626 a 2426 281667 586.595 368.541 178 405 3319.96 112724E+06 67" 8.87745E-84 .87 9 6 .838576 12389.4 124% 5 7376.31 !2418.5 %1 ?66 38227.7 .885 143& 78 .05 8 1-6 4.58997E-84 .8471732 26.6883 25 168827 28N 39 .13 .06 !4293.5 3.53547E-84 .386698 935 81 5379.19 5" 02 43939.9 16868.1 16224.5 6369.99

51 x16 .181948 239828 4.4881.39E-84 12480 144412 8 0 0 8 8 124813 .8299686 13 9.2 3 0185736 15 7.19187E-84 158.972 317.944 76 -157844 .863867 68.8619 .538144 458 2 6 0 6 a 8 18.3582 8 2380 191667 586.543 412 428 178.457 :819.28 1.322M06 819825 i 0039E-83 87 9 8 96A572 14887 15263.7 9M 83 15263.7 SK 964 38285 4 1062 1664.43 . 85 9 M-63 4.58997E-84 .8615''46 2& L%4 25 129641 36?i.48 .17 .06 142118 3.53547E-81 44LV 10!4.95 4059.94 5839.13 44*6.3 159918 16368.2 6373.9

5a. x893 .192936 215842 4.98895E-84 138655 .153133 8 8 8 8 a 138655 ,8315625 14.5 18.4 15 . 811758! 15 7.18465E-84 182 954 363.988 M. 92 -157844 .93448 86.7547 .841849 458 2 6 a 8 6 8 22.69!7 0 2228 185 585.752 468.937 178.248 2435 85 153190E+86 983183 119698E-03 .87 9 0 1.08066 17658.8 18138.3 11332 MX 3 588.747 3W9 1239 !872.81 05 8 2E-03 4.58997E-04 .876815 26.0321 25 Lem 4688 9 .21 .06 14U& 8 3. S3547E-84 .495864 1108 82 45N 8 5035 43739 9 162611 165215 63521

18.8815 . !99446 185162 5 84262E-84 162236 .1583 8 0 6 6 8 162236 .0341837 16 it 6 4 . 8137233 4.5 7.28263E-81 187 539 375.078 2269.46 -157844 976016 84.6628 125664 459 2 8 8 8 8 6 38.9!11 8 2068 171667 506.26 488 855 178.74 2129.13 1.55663E+86 183183+06 132228E-83 .87 9 8 1 I 18386 21968 12178.1 21885 8 5x8 W 38599.5 1416 288218 05 8 2E-83 4.58997E-81 .0891686 a 9818 25 .901839 5359. it 25 86 14868.2 3.53547E-81 S58447 !133.92 38978 4925.37 44367 3 15632 7 16645 6 6488.85

A9-20 1-.--

Appendix 9-7 (cont'd)

47.6826 .281885 114531 7.48283E-84 2 83826 . 159681 f 8 1 f 8 2 83826 .838416! V. 5 i2 8 4.5 . 80898 4.5 7.26388E-84 169.888 339.615 2831.86 -1. 57843 .977348 82 5862 178924 458 2 8 8 8 a 8 47.96266 8 if28 151667 584.868 489 564 179.942 1871.72 131R6E+06 938999 135554E-86 .87 9 9 117416 16646.3 23887.5 !!369.9 23887.5 499.946 39935.7 1593 2835.44 .85 0 2E-83 4.58991E-84 .8969715 25.9379 3 .857924 5818: 65 .29 .06 14016,8 3.53547E-84 .684831 li8i.16 3812 49 444149 45981 i 14896, 9 i6724.1 6544.8

47.6826 .289935 144531 7.4VM-% 2 83826 .166625 8 8 8 8 8 2 83826 .8384161 17.5 12 8 4.5 .8!1898 4.5 1.26388E-84 191229 386.459 2219.47 -157843 .977348 82.5862 t 78924 456 2 6 9 8 6 8 47.9826 6 1818 i5i 667 584.848 488.564 179.942 1871.71 155587EW 186852E+06 154286E-83 .81 9 8 128827 AW.3 23887.5 !2938.2 23887.5 499.469 39935.7 1593 2219.9 .05 8 A-03 4.58997E-94 .1!8354 25.7539 25 .153933 662123 .33 .86 13994.1 3.53547E-94 .684831 lift S 3842 49 444149 45983. i [email protected] i6765.4 6544.8

47.6826 218895 144531 7.48203E-94 2 83&,6 .17Y82 a 8 6 0 6 2 83826 .8304761 17.5 i2 8 4.5 . 817898 4.5 1.26388E-84 216.951 433.382 23916 -157843 .977348 82 5862 178924 458 2 8 6 8 8 8 47.9026 8 1828 151667 584.0 488.564 179.942 1877.72 t 73634E+06 119883E+96 172897E-83 .87 9 8 13522'5 21824.5 23887.5 14586.5 23887.5 499.893 39935.7 1593 2397.24 .05 8 2E-93 4.58997E-84 .12373 25.6724 25 .672427 74218 .37 .86 !3986, 8 153547E-91 .6@5831 lift 16 3042 49 444149 459811 11896.9 16779.3 6544.8

47.6826 .225687 14531 7.97199E-84 2"3 .179128 8 8 8 f 8 21!393 .0399299 17.5 V. 8 4.5 ai81299 5 1.2668E-04 248.873 488.146 256131 -157944 975555 82 5862 214798 458 2 6 9 9 a 8 57.5792 8 1820 151667 585.284 488 564 178.716 1882 's 2 84968E+96 146898E+86 191589E-93 .87 9 8 1.48869 24818. 7 28665 17698.4 28665 498.53 38572 9 1593 2568.49 .06 a 2E-83 4.58991E-94 .146418 25.50 25 .5488:9 87M 11 .41 .07 13974.1 3.53547E-04 .78328 1279.97 "842.493 4935.02 44336.7 15663.3 16882 7 6484.32

A9-21 APPENDIX 10-1. SPACE BASED RADAR-R ANALYSIS (Baseline Configuration Including N = 5 and 2 Burns) 3

OPTOTYlOOA - i MM, 1 OTY, i SHUTTLE AlNLlIBt PHASED ARRAY

91 B 8P Ca D2 DB Co DM OT DY E EE EM F9 O OE Ki KP KQ KS KT LD LF LM LO MH N PI RH FM RS Si SP T2 T8 TO TF TO TL TM TM TV R WL WW WS WT WU ZH

188 i 1 814 1 36 i 2 3 .91 4E487 .187 388888 1 3217 3.54E+86 .2 21 .5 15 4.1 18 48 1788 684 .6 .85 114159 .863 AM 88 37888 3336 .825 2E-83 SE-03 .04 9E-83 .0 .85 .8! At .02 3.3E-94 18 68888 1 18582 .47

Ai A AA AC AL AM AT Be 82 84 85 86 89 BC BE BF BU R C2 Di D DA A Fi FA FC FD FI FS H Ii I IS K8 K2 K5 K6 K7 K9 K KE LA LA(FT) L8 LC LP LX MI 82 MF MU H HA HD NO P PC Pi PK PL PW Ri SM Ti T TA TE TP TQ TR TS TT TW TX Y wi W WA W8 WC WK IP WK WN WY WZ x8 ZI

392699 386 8 253426 2 57794E-83 7.74735E-83 i 1781 6.67386 8 8 8 6 8 8 6.67385 .8718844 oi882 ii 5 8 2 5 .8526281 134.385 2 5 7.72641E-84 9 6 8 -1.57841 .761373 519.6 .386796 17678.4 426.288 2 6 6 8 8 8 454.463 6 2418 288.833 463.221 8 188.779 2587.8 218688 6 8 .88 9 9 8 56.4749 1216.45 1345.76 8 l2O 5 612 389 4888.4 .885 233.272 .85 8 2E-63 4.58997E-84 3.18911E-83 58.2845 3 25.2845 191.346 at .865 15267.4 1 WE-% .187871 .18325 2774.86 6451.85 4336.18 .1 46437.9 !3562. i OP.1 5572.55 86.6 9

.934624 388 8 247156 3.88125E-83 7.94389E-83 2 88387 18.1596 8 6 6 6 8 8 18.1596 .8871423 8299285 14.5 18.4 3.5 .8827181 125.875 3.6 7.4838AE-84 8 8 6 -1.57842 .758879 519.6 1_ 4"l- 4 42851.6 456.846 2 8 8 8 8 8 1123.64 8 2388 196.333 464.927 8 179.874 2395.64 t 81471E+86 8 8 .87 9 8 8 49.8466 5858.58 6277.7 8 38821.6 517351 38976.6 1.239 115184 .885 8 2E-83 4.58997E-64 .8175245 29.6416 25 4.64157 1851.47 85 89 14458.2 3.53547E-84 .528693 .271927 4852.94 629144 5186 i 44793.8 15286.2 15967.9 5823.15 86.6 8 A10-1 Appendix 10-1 (cont'-- )

i 39487 s'd1 8 238418 4.76174E-13 8.21779E-13 4.1816 i2 7289 1 8 8 8 1 8 iz 7289 .1975883 8389521 15.1 i8.88 3.7 .1836 i23.229 3.8 7.48179E-14 1 8 1 -157112 .757583 519.6 2 38097 62765.3 451.742 2 8 1 1 1 1 1763.5 1 230 195 465.808 1 178.132 2357.87 i.7362E416 1 1 .8r 9 1 1 49.1188 1124.2 18178.4 8 462711 Se7.164 37924.1 Lis 2838.47 .12 1 2E-03 4.58497E-14 .8331138 27.427 2S 2 42782 1916 83 .09 .123 143716 1 S %?E-11 .935653 34W4 5086 616175 5246.43 .1 43% 16419 16515. S 566683 86.6

i 63363 381 8 226823 S. 35943E-13 8 65599E-83 4.98888 14.3246 8 8 8 8 8 8 14.3246 .183532 .832M 16 ii 6 4 .116633 128.46 4 7.48283E-N 1 8 8 -157842 .757388 539.6 3.26726 73588.9 456 2 8 8 8 8 8 2232 ?4 6 2281 M 466.882 8 177.998 2299.1 2 32812E+86 8 8 .87 9 8 8 47.7522 13397.3 143318 8 54615.2 583.652 37776 1416 2861.96 .13 8 2E-63 4.58997E-94 .848!52 26.60 25 166346 2889.12 .13 .135 14288 153S47EM i 31685 .384247 5484.6 5773.86 5388.88 1 43428.7 16579.3 166812 5644.69 86.6 8

178328 388 6 228893 5.7187E-83 8.88837E-93 5.34385 112688 8 8 8 8 8 8 15.2688 .18687! . 8334778 16.6 1288 4.2 .124234 118.614 4.2 7.48414E-44 8 8 8 -157812 .7SMS 519.6 3.92773 89153 458 2 6 8 8 8 8 2533.23 8 22% 187.5 466.134 8 177.866 2268.38 2 9i658E+86 8 0 .87 9 8 a 47.1239 16839.4 17229 8 59826.3 SK 79 37629.3 148Q 3782.35 135 8 2E-93 4.58997E-84 .8633838 26.2588 25 1.25881 X83.83 17 N 14283.4 3. SM47E-N 160937 .488163 "( .27 15622. A Ss54. 8 1 43252 16748 16833.7 56& 76 96. i d

193572 338 0 213189 6.11565E-83 9 21383E-A2 5 88 6 it. 3417 8 9 0 8 8 8 16 3417 118595 0347887 17.2 12.56 4.4 .133822 116.768 4.4 7 48474E-84 8 8 8 -157842 757838 R9.6 4 Q324 87879.8 458 2 8 8 8 6 8 2984.31 8 2218 184.167 466.186 6 177.814 22216 3.4147iE+86 8 8 87 9 8 8 46.2861 19715.4 28543 0 65278.2 588.668 37571.1 15576 449219 14 8 2E-63 4.58997E-01 87858!3 26.8!48 25 101482 4718.88 21 .145 14119.4 3.53547E-81 2 8619 .433897 68!3.92 5424 77 5376.18 .1 43185.1 16814.9 !6985.2 %S. 87 86.6 8

2 89544 368 8 285464 6.54179E-83 9. SM-0 6.28633 17 4787 8 8 8 8 8 6 17.4787 114384 83599!8 17 8 13.84 4.6 142264 114.922 4.6 7 48545E-84 8 8 8 -157812 .756953 519.6 5.54245 94289.3 458 2 8 8 8 8 8 33219 8 2178 188.833 466. d.44 8 171.756 2182 75 3.84913E+16 6 8 .87 9 8 8 45.4484 22223.6 2431i 9 8 78916.9 499.983 31586.6 16284 525184 145 8 2E-83 4 58997E-14 .8937224 25.8485 25 848548 5623.34 25 iS 14831.9 1 S WE-84 2 41822 459849 6258.91 5231.17 5399.86 1 431111 16888.9 17138.2 5681.44 86.6 8 A10-2 Appendix 10-1 (cont'd)

2141 388 8 281694 61e364E-83 9.73443E-83 6.42299 17.9329 1 1 e 9 e 9 17.9129 .11579 .8366645 18.1 1128 4.1 .145014 313.999 4.8 7.49469E-84 1 8 9 -157942 .757944 W.6 191182 %M.1 458 2 e 8 9 1 6 3489.72 8 2151 179.167 466.182 8 177.838 2463.31 1.34333E+16 e 8 .87 9 8 9 4105 251717 25932 2 6 72634 499.382 37575.7 16638 6833.88 .145 1 1-13 4.58991E-84 .188384 25.1351 25 .735189 6513.86 .29 iS 43995.3 3.533547E-84 2.71818 .466427 6388.89 5134.2 5374.5 1 43196.4 i68e9.6 17219.2 5614.75 86.6 8

2 26395 381 8 197968 6. 99864E-83 9.91881E-83 6.78584 38.6759 8 9 8 e 8 8 18.6758 .118242 .937M 38 4 13.52 4.8 .151991 1!3.876 4.9 7.48629E-84 8 8 8 -151142 .756862 539.6 6.51441 181782 458 2 e e 1 8 8 319E 69 1 2131 M. S 466.361 8 !77.693 2143.85 4.88682E+86 8 8 .87 9 8 8 44.6186 27748.4 28575.4 8 76856.3 4%942 37436.9 16192 6883.62 .15 1 1-13 4.58997E-04 .124197 216397 25 .639661 7445.83 .33 !S5 33986.6 3.53547E-EN 3. !7214 .486818 6584.47 5839. 0 5425.49 .1 43838 9 16969.1 17226. S 5594.82 86.6 6

2 3%w 388 6 194259 7.28582E-83 .918587 7.15812 49.4596 8 8 8 6 8 8 UAW .328713 OMM i8.7 13.76 4.9 15833 112153 S 7.48617E-84 8 6 8 -151842 756674 519.6 7.1633 117365 458 2 8 6 8 8 6 4114.53 1 2118 175.833 466.435 a 177.565 2324.36 5.23876E+86 8 8 .87 9 4 8 44.19!7 38216.9 314121 8 832881 49P.598 37294.1 17346 7556.67 .155 8 2E-6 4.58991E-04 .1469% 2S. 5648 25 . 564181 8118 97 31 .36 13977.8 3.53547E-0I 146849 .586118 6789.81 4944.94 5477.76 .1 42867.4 17132.6 17242.8 5572 77 86.6 8

2.43473 3% 8 M94 7.46382E-13 .8383814 7.3842 19.9375 8 e 8 8 8 1 19.9375 .322179 1381998 19 14 5 .162238 13123 5 7.487233E-14 1 8 9 -151942 .756764 519.6 7.6m 189 7 456 2 8 8 8 8 9 43328.18 8 2891 174.167 466.373 8 177.621 2184.86 5.64223E+96 8 e .81 9 8 e 43.7729 32576.4 33374.9 a 82998 4 498 3% 17362.9 177 8294.24 . M 8 2E-e3 4.58997E-4 .151687 M50 25 .512!51 9281.23 43 % 13%9.6 3. S3547E44 3.86783 5140 6749 84 4851.64 5452 7 1 429458 17851.2 i7257.8 m. 96 86.6 8

A10-3 r. Appendix 10-1 (con"

R! A AA AC AL All AT M 82 04 BS 86 8D BC BE BF BU ci C2 Di O DR OX Fl FA FC FO Fl FS M 11 1 i5 K0 K2 K5 K6 K7 K9 K KE LA LA(FT) LB LC LP lit Ml M2 w NU N NA ND ND P PC PJ PK PL FN Ri SM Ti T TA TE TP TQ TR TS TT *ii TX V iii N NO 0 NC NN WP NX w w iQ 0 21

345575 380 8 289269 2 89368E-83 9.38288E-83 L 83673 7.3L2i 0 0 8 8 0 8 7.3126 .8156133 8353374 i8.6 7.28 12 .896984 131.874 2 3 7.81851E-64 0 8 8 -L 5704! .763896 519.6 .289873 15558 425.355 2 8 8 0 0 8 532.961 8 2198 182 5 46L 495 8 182 585 264124 158246 6 8 .88 5 8 8 418673 913.639 917.097 8 !8788.3 1020% 42782 9 .7788 21L 978 .85 8 2E-83 4.58997E-84 2 86951E-83 39.3603 A 29.3683 M. M .41 .0% 16717.7 3.53544E-04 .0972973 .175617 2416.53 5327.82 3639.47 .1 48617 11383 11410.8 5834.84 86.6 8

581217 388 6 228817 3.92731E-3 8.5$85'1 E-03 2 64365 18.4744 0 8 8 8 8 8 it 4744 .8886262 8323185 13.9 9.92 3.3 .88310!2 126.921 3.4 7.49889E-84 8 0 8 -L 57842 .768132 519.6 L 19956 '39652.4 458.866 2 0 8 0 0 0 1188.68 8 22% 198.833 464.87 0 179.931 2434.17 984224 8 8 87 5 8 8 47.9617 5228.69 526L 85 8 28878.1 389.871 39922 8 L 1682 1188.29 .885 8 2E-03 4.58997E-84 .Oi6m 14.9719 16 4.97189 1883.86 05 89 15508 3 3.53547E-04 .5887 .20777 M 21 5824.62 451L 92 .1 45888.3 14M 7 !4156.7 5965.47 86.6 8

L 19381 388 8 224838 4.61471E-83 8.73243E-83 3.58142 12 3383 6 8 8 8 6 8 12.3363 .8968698 8328985 15.4 iL 12 3.8 188363 M 386 3.9 7.48517E-64 8 9 8 -L 57842 .759143 519 6 2 15482 53718.1 457 513 2 8 8 0 0 0 03.28 8 2279 189.167 464.746 8 179.254 2337.76 L W'E+% 8 8 .87 5 8 0 47 5428 9is. % 9452 ii 0 39789 5 37. 216 Mi L 3452 1.977 49 1 0 2E-93 458991E-04 0'312436 12M iP 2.05554 1374 61 125 1522 3 3 533547E-04 4876e4 3L•3 44493 572= '.2 4788 22 1 4%24 1 149-75 9 15896 585 1. 86 h 8

I. 41686 389 0 218934 512396E-+3 8 9.Nf IE-03 4 25858 13 6913 0 0 0 8 N 0 13 69r 181232 0337774 16.3 it 84 4.1 111542 119.537 4 1 7 48171E-N 0 0 6 -L 57842 75864 519.6 2 97117 63754.9 458 2 0 8 0 0 0 2042 08 8 2248 186.667 465.891 0 178 91 2279.75 2 20882E+66 0 3 87 5 8 8 46.9145 12706.8 13059 ' 8 474791 378.9!8 38188.3 L 4514 2818.63 i1 0 2E-83 4 58997E-64 0459581 117896 i8 L 7"55 2757 52 m . 0 115 15868.7 153544E-A4 1.29374 349L"s6 4913.86 W3.04 4928.85 1 44584.3 15415 7 15177 9 5195.95 86.6 8 A10-4 Appendix 10-1 (cent' .)

ism 381 A 2!!185 5 51214E-03 9.29698E-A3 4. 605 14.7353 A 8 1 1 A A 14, 7M .184997 0351!68 169 12 32 4.3 .1!9997 117.691 4.4 7 40156E-04 1 1 0 -157842 .75" 519.6 3.59167 69904.2 458 2 0 1 8 0 0 2363.4 1 2201 i83.333 4651!5 1 179 885 2241 2 72787E+86 A / .87 5 0 / 46.8767 15745.2 15751 8 s. 7 34282 3m 5 i 5222 3629 01 .115 1 2E-03 4.50997E-04 .8684756 113630 1/ 13M 3618.55 17 .12 14818 3.53547E-M 16616! . 371272 5432 e9 5375.78 49X 7 . i 4455 4 15444.6 15917.9 5791.95 K 6 1

176715 381 A 209269 5.91764E-03 9.38288E-03 5.38144 i5.8214 8 0 8 8 8 8 15.8214 .18879 . 0353374 17.5 12 8 4.5 12804 !15.845 4.6 7.48051E-04 0 0 0 -157842 .757924 519.6 4.47389 79516.9 450 2 0 0 8 0 8 2725 4 8 2198 182.5 465.58 8 !78.42 220219 3.32316E+06 8 8 .87 5 0 8 45.8473 19!86.8 !96212 8 59718.8 352 732 38244.8 i 593 445154 in A 2E-83 4.58997E-84 .0761434 ii 845 A 1.8658! 4568, 6 21 .13 14524.1 3.53547E-84 2 84324 405874 5 89 5327.02 5128.4 .1 43969.5 16840.5 16416.5 5714.73 86.6 8

195693 380 0 206164 6.33861E-03 9. %5-03 5.85279 16.8571 8 8 0 6 8 8 16.8571 ULN 0359918 V.8 13.84 4.6 .137271 114.922 4.7 718185E-84 8 8 8 -157842 .75749 519.6 5.1602! 97786.6 458 1 8 A 8 0 8 W. 8 8 2178 !81.833 463.877 8 179123 216275 184913E+86 8 8 .87 5 8 0 45.4484 22223.6 M. A 6e054 347.928 379!4.8 i 6204 525185 135 8 1-03 4.58997E-94 .0928218 18.8736 A . 8w m 31 25 .14 142418 3.53547E-84 2 410 .4335641 5939.76 5239 it 5249.87 .1 43588.2 16419.8 169!9.1 5665.43 86.6 8

2141 30 6 20M 6 70364E-43 9.73443E-83 6.42299 17.91; 9 8 8 8 8 8 8 17.90 .1!579 8366645 18.1 0.28 4.7 .143814 !!3.999 4.8 7.0+69E-84 8 A 8 -157842 757844 519 6 5.91182 ?"1339.1 458 2 8 0 8 8 8 '489.72 6 2158 179.167 466.181 8 177 818 2163.31 4.34311E+04 8 0 .87 5 8 8 45.0295 25875.7 25932 2 0 72628.4 344.451 37575.7 160 6835. P181 145 0 1.63 4.58997E-04 188384 18.735! to 735109 643.86 29 .15 14894.4 3.53547E-81 2 77808 .466427 6';48.89 5134.2 5374.5 1 13199.4 16889.6 !7106.4 5614.75 d6.6 8

2 26195 3% 8 197968 6.99044E-13 9 91808E-83 6 7856/ 18.6759 8 0 8 A 8 8 119756 K8242 8373562 18.4 13.52 4.8 151994 1!3.876 4.9 7 48629E-04 8 8 8 -157842 756862 Si9. 6 6 51441 18!762 456 2 8 0 8 8 8 374169 0 2530 177 5 461387 8 1"? 03 1143.85 4 80682E+04 6 8 .87 5 0 8 44 6!04 27748 4 28515.4 8 76856.3 312 8% 37436.9 1.6991 6883.62 i5 0 2E-03 4 58997E-84 124897 Ito? to 639661 7415 83 33 155 14083 9 3.53547E-64 3 !2201 486818 M 47 583913 5425.49 1 43A31 9 16969.1 17287 6 551.82 86 6 8 ' A10-5 Appendix 10-1 (con,'::)

2 38614 388 6 194259 7 2OW-83 .8181!7 ?. 15812 19.4596 1 8 1 1 1 1 19.496 . Law .03M It? 1176 4.9 .15M 112153 5 ?. 46117E-44 9 1 1 -1.57112 .756674 W. 6 T. ml i17m 458 2 1 / 1 1 1 4114.53 1 2119 175.833 461435 1 177.565 2124.36 123876EN6 9 1 br 5 4 1 44.1917 39246.9 3142.1 1 8l2AB.1 341.176 37291.7 ON M6.9 .15 / 2E_v 4.59997E-/4 .149616 11.5648 19 .564711 9111. 9? .3? 16 14973` 3 3.53547E-11 146944 .516111 6799.87 4944.94 547?. 76 .1 42667.4 I?M 6 !7228.6 X72 7r 86, 6 1

2 43473 368 1 196 1 7.16382E-13 .9!831!4 7.3642 Ulm 8 / 8 1 1 1 !9.9375 =?9 .am 19 14 5 .162236 Iii 23 5 1.48723E-14 .8 / 4 -157142 .756764 Si9.6 199557 451 2 8 A A / / 4328.18 6 2899 174.167 466.373 / 177.62? 2114.86 S. 64223E4* 1 8 .87 5 1 9 42 TO 32576.4 33374.9 9 82999.4 338.95 37362 9 177 8294.24 An / 2E-0 4.5M-04 . L%W 11.5622 18 .5!2151 928123 41 .16 14863.5 3 53547E-94 3.88783 .511885 6749.81 485164 5452 7 .1 42%18 11854.2 ON. 5 mm 96 86.6 8

Al R AR AC AL AA RT Be 82 64 65 96 99 8C BE OF 8U Cl C2 Di D DA DX F1 FA FC FD Fl FS N I1 I Is KS K2 K5 K6 K7 K9 K Ml LA LR( T) L8 LC LP LX Ml M2 MF M! N NR ND NO P PC Pi PK PL PN Ri SM Ti T TA TE TP TQ TR TS TT TN TX V Nl N NA WB WC NN NP WX NN w NZ X8 21

329867 388 9 178841 3. am-03 M277 .989692 8.19882 8 8 8 8 8 8 8. !9182 8886464 9415M i8.3 7.64 21 80478 137.997 21 7 9481E-N 8 9 8 -15794 7656 %96 181839 11843.2 424.762 2 9 i 8 8 8 448 299 4 2828 168.333 468.314 1 183.6% 266.37 124297 8 9 98 2 8 8 42. ;868 717 65 797.447 9 18!43 1 449 481 44895.7 7434 195.523 95 8 x-83 4 %94?E-94 2. N72E-83 37 7133 5 u 71;3 159 94'6 K e5 177419 3 535+1E-64 *17445 172,973 2196 0'5 4n' i i;2 51 1 `!8188.7 mix 9912 57 6913.85 86.6 6

?04248 m 9 281694 4.41679E-43 9 7.443E-93 2.41274 ii 383 0 8 9 1 8 1 it 383 99:9971 9366445 13.6 9.68 3 2 8893696 127 814 3.3 7 7683E-94 8 8 8 -157841 76M 519 6 1 82944 W9 448.369 2 8 8 1 8 8 Me, 89 8 2158 179 167 462 388 9 181612 2453.41 748817 8 8 88 2 1 8 45 9295 4:4.39 4515.62 8 25254 5 298.9% 417912 11326 18N. 53 8 1-6 4 58997E-M 815174 11 717 5 5 71698 411.44 85 8BS 16944 8 3 Ss ?E-64 477611 24991 3376 5134 2 3999 79 1 47498 125ie 12596 1 5696 8 96.6 6 A10-6 Appendix 10-1 (con-6.1.)

113897 388 0 203575 S.ISM. -03 9.6445E-03 3.3918 13.3425 i 0 1 0 1 1 13.3425 101518 14.8 1164 3 6 .016143 124. is2 3.6 7.12415E-M 1 / 0 -L 57440 .76126 519.6 L 832!8 51898, 8 457.852 2 0 / / 0 1 1007.95 2161 181 463.298 1 188.782 237616 136669E+06 0 8 .0 2 6 1 45.2361 769a 0 8036.94 1 37489.6 238.625 41779.9 12744 180167 . i 1 2E-83 4.50997E-M 1281312 7.9!313 S 2 99313 !129.81 M 116 16535 0 3. S WE-0 .06368 .312 0 4119 92 5182 4 4367 24 ± 434 A 1369.2 13732 3 5661.89 16.6 1

13023 311 6 191259 5.89144E-13 P10017 4.1469 15.26% 0 0 1 1 A 8 15.2658 .188545 0381678 16 116 4 i28511 129.46 4.1 7.71793E-64 1 8 A -157141 .764917 Sig. 6 V646 62!99.9 458 2 1 0 1 8 a 2383.67 4 1111 175 833 463 54 8 111.46 2299. i 18"49 4 6 .0 2 1 1 44.1917 11627.3 5226.9 1 46212.8 216, 562 486116 1416 2005 .11 0 2E-13 4, 58997E-1 . M2?2S 4' 83433 5 2.13433 2333.5 i3 .005 16133.3 3 S3S47E-0 & 21066 .343539 4554.41 4944.94 4464.72 .1 44 ►?5 i 13964.0 14421.6 5324.3 86.6 A

15173° 301 A lm S Mal ? 9.4181!9:-03 4.527 15 M 0 4 A 1 0 1 1:344 1M664 037"'61 16 6 1208 4.2 .124408 518.614 4.1 7.49082E-44 a 0 0 -L Sa42 159578 519.6 13495 61278. S 458 2 8 0 A A A 2343.59 A 238 177 S 464.44 1 1119.%1 226938 2 475M+% 8 8 87 1 6 8 44 6106 142%. 6 04676. S A 54963. i VOL 6% 39505 14066 3384 9 115 4 2E-83 4.2997E-ili . ASAIba, 6.41119 S 144208 3404 83 it 12 iSd3t 6 3. 47E-44 i 60074 3655: 4890.48 50s3. i3 4666.91 i 45113. S 14596.5 140 M. 45 06.6 A

L 69646 301 4 283575 a 9206E-03 9.041-A3 5.00 05 VT.6 1 0 0 1 6 1 is. 6176 181617 0363258 17. S lie 4 S .32M 113 945 4.5 7 40137E-44 A 0 8 -157111 758 9 W.6 4.29417 76336.1 138 2 4 8 9 0 1 2726.92 A 264 188 465.146 0 170.054 220219 3.58A94E406 4 0 87 2 1 A 45 2389 184119 18836.3 1 57338.1 183.181 3&726.7 1593 4198 56 12 A 1-63 4 58991E-4 8741166 6.14106 S 0. i1216 4467.4 2 325 15366.3 3.2%?E-11 2 SIM 3944;4 5m 5182.08 4951.49 1 41503 4 1586.6 1505 5786.15 96.6 0

L 77186 381 A 199023 61wry 9.82561E-3 S. 31557 161825 4 1 8 1 4 4 16.40 118775 4370079 18.1 1328 4 7 13331 113.999 4.7 7 48024E-M A 0 4 -157142 750516 509.6 4.09154 79128.9 4! 2 1 4 4 8 A 2928.12

0 2141 078.333 165.168 1 178.032 2M 31 'A. OM*06 A i 87 1 1 8 44 0201 2UV 9 214611 4 ti8099.7 178 701 36M 1 16638 5170.46 11 4 1-6 4 'i^3,'Y)F-44 880747 592475 S 924?% 5324 44 25 125 04rs10 ? 53347E-M 2 3769 1!6637 54613 5056 4924 48 1 440 3 15514 7 16128 SM. 19 86 6 4 A 10-7 Appendix 10-1 (cont'd)

213575 311 a 1!7961 t 62267E-13 9.91111fi-3 6.11126 it 7851 1 1 1 1 1 1 11.71"1 15m 1373562 18.4 1152 4.1 .144392 313.176 4.9 7.4 17E-14 / 1 -1. 57142 757687 5194 5.86297 93613.4 451 1 1 1 1 1 341252 1 2131 177.5 461 742 1 171.251 214185 4 22341E+1d 1 1 17 2 1 1 44 6516 24315 25717.9 / 695717 574. 617 38161 1 6992 "n. 94 13S 2E-13 4.51997E-11 .5161Q4 S. 76524 5 761139 M 45 .29 .14 '14M 1047E-/1 2 74432 44M 6114.14 5139. i3 Si96.14 5 4373!. 7 16218.3 !6621.9 5687 72 86.6 1

2 2321 3101 4 194259 7.13176E-13 .111117 6.69631 58.819'1 o 1 1 1 4 1 18.8092 11164 1381611 11.7 13.76 4.9 153137 112. M 4.9 7 46332E-14 1 1 1 -i 572 751231 539.6 6.Or 18"39 4% 2 1 1 6 1 1 3849 a1 1 2511 115.833 466.154 1 177 W. 2324 X 4 67245E+86 1 1 .11 2 4 8 44.1917 269r7 29385.5 8 75968.9 Sri 819 37717.4 17346 6739.74 145 1 2E-13 4.58997E-01 522534 5 ,tw 5 bum i,S3 25 33 0 14377.4 3.53547E-11 3.19352 .481584 6379.3 4944.94 sm. ?9 1 43353.4 16646.6 16854 8 5635 94 86 6 1

2 27765 31A 8 191694 1.21834E-93 8183154 6.83296 19.272 8 1 1 1 8 6 19 272 128166 6381998 19 14 5 156938 iii 23 5.1 7 48271E-84 8 1 9 -157142 .7573''6 519.6 7,11768 112488 458 2 8 1 1 1 1 4041.46 1 2891 114.56: 465.0 0 178.811 2514 86 5 89177E+96 1 1 87 2 6 1 43.7129 29358.2 aim. 6 6 77643 7 171115 37M. 2 177 746515 14$ 9 2E-63 4.51997E-14 136961 5. WA 5 585862 8217 64 37 i5 1436 1 S3S47E-% 3.43362 461562 6115.73 *ft 64 05. e7 1 43437 i6563 1688L 9 5646.81 86.6 6

2 323 is 8 1%%4 7309-43 9395114 ;. *962 19 7451 1 1 1 1 1 1 L° 7451 111547 139^'2s 19 3 14 24 5.1 16179! 511317 5 1 7 66218E-44 8 6 8 -157112 757424 519 i 7 W433 164"8 456 2 9 8 1 8 1 4242 44 1 2878 172 5 4e 922 9 178 878 2 35 A 5 4f,'.24d+* A a 37 2 8 8 4: !A :4 .:L2 6 g '? if ° 1 3X22 ?7 4 1 1 2^+ y- 4 6 2E-93 4 589?7E-84 5 5978 5 5:811 :97 '1 41 iS 14348 7 3 5-,547E-* :7'16 49`441 4% : 4754 22 5w 46 1 43522 16478 tim 6 507% W

A1O-8 S

APPENDIX 10-2. SPACE BASED RADAR-R ANALYSIS (Baseline with Unit Area Weight of Lens WL = 6.6 x 10-4)

91. 8 8P CD 02 DB DD DH DT w E EE EM FB G GE Ki Kr KO KS KT LD LF LM LA M My PI RH RM RS Si SP T2 T8 TD TF TG TL TH TH TV TZ HL HN HS iR 111 ZH

188 1 1 .814 1 36 .1 2 3 .91 IF+87 .187 388888 1 3217 3.54E+96 .2 2 5 4.1 18 48 1788 684 .6 .85 3.14159 .863 .8643 88 4E+97 3338 .825 F-B 5E-83 .84 9E-03 .25 85 .Ot .4', 82 6.6E-84 18 68898 1 18882 .47

Ai A AA AC AL AM AT Be 82 84 85 B6 89 BC BE BF BU Ci C2 Di D DA DX Fi FA FC FD FI FS H R I IS K8 K2 K5 K6 K7 K9 K KE LR LRTT) LB LC LP LX MS M2 MF MU N NA 1D NO P PC Pi PK PL PH Ri SM Ti T TR TE TP TQ TR TS TT TH A v wl H HA 113 we ill HP wx HH w HZ A ZI

:Z47412 388 8 144531 3.95198E-93 .8!35845 i 19223 18.2294 8 6 8 8 8 8 A.',1% .8889842 85+.1657 16.9 7.52 2.3 .OBV03 136.151 2 4 7.72671E-84 9 8 8 -15" n+i 761387 519.6 .262788 17882.5 426.283 2 9 8 8 8 8 1867.86 8 i828 151.667 463.211 8 186.789 2626. ii 176828 8 8 .88 9 6 8 338.118 1816.28 115272 6 134877E+97 612 486 48876.4 AM 348.5% .0 8 2E-83 4.58997E-4 3.18727E-93 58.3866 25 25.3855 191 236 .8i .86 15267.6 6.83547E-94 .156333 .184814 2184.27 7113.14 4332 B .1 46459 4 13549.6 135%.4 5574.85 86.6 8

96227 388 8 154217 5.48739E-83 .8127313 2.88398 14.6713 8 8 8 6 9 8 A.6713 .184761 847952 14.1 19. i6 3.4 .119494 125.998 3.5 7.48844E-84 8 8 6 -157842 .758641 519.6 1.38912 43257.2 451337 2 8 8 8 8 8 2343.62 8 1888 156.667 465.5 8 178.5 2414.91 %9585 8 8 .87 9 8 9 39.3746 5598.86 6893.36 8 3.41634E+87 516 467 38333.6 t 2836 1759.12 .69 8 2E-83 4.58997E-84 .9183553 29.4363 25 4.43631 1883.32 85 .895 114455.2 6.83547E-94 867432 275353 3252 57 7589.86 0 d 5895.94 1 449616 15938.4 16881.6 5728.81 86.6 8 A10-9 Appendix 10-2 (cont'd)

i 2535 388 8 158951 37817E-83 8138869 3.76819 V. 8278 8 8 8 6 8 8 17 8278 .1!2873 6489888 15.4 1i 12 le 138656 122 386 3.8 7.4821E-84 8 8 8 -1 57812 757439 519.6 2 26256 56481 457.748 2 8 8 8 8 0 3155.54 6 1868 195 46'1912 8 178.886 2337.76 16984E+86 8 8 .87 9 8 0 38.98 9759.82 9924.71 8 4.58757E+87 567.126 37876 13452 3L22.73 183 8 2E-83 4.58997E-81 .833186 27.4187 25 2 4187 V9116 .09 ii 143712 6. 8331 E-61 1 43792 .3.77676 377183 7429.23 5264.12 .1 43535.6 16161.4 165!6.7 5659.63 86.6 8

148126 388 0 146123 7.89811E-83 .804364 4.44378 18.9784 6 8 8 4 8 6 18.9784 .1!9!48 . 85A686i 16:3 it 84 4. t . tw 119.537 4.2 7.48334E-84 8 8 8 -i 57812 .757228 519.6 1 io 666'52.8 458 2 6 8 8 8 6 3915.32 0 1838 152 5 466.856 0 177.944 2279.3 2 332549E+86 8 8 .87 9 8 4 38.3274 13429.5

13653 6 5.33W2'(E487 583.619 37715.3 i 4514 4452 07 1 115 d 2E-C 4.58 X-04 .84$105 26. m 25 1 N27 2897 8i L 12 14287.3 0.83547E-14 2.94349 35956n 41'x4. w 1LO1 52 5±2118 .1 433% 9 16649. i 16682.6 50.62 86.6 8

108861 384 d 14!372 7 77884E-03 d1_,%i 5.46582 41-769S 9 14 0 d 8 8 20.165$ 124666 8.52389! 16.9 12.32 4.3 .169845 117. 691 4.4 7.484L4E-04 6 d .6 3.9E-79 75982.8 458 2 d d d d 8 4690.29 0 1888 158 466.146 6 177.854 Z241 ` 89525E+86 0 d .87 9 0 0 37 6991 16716.2 17119.6 d 6.144-s9E447 581.785 37615.3 15222 5.726.5 0 8 2E-63 4.58997E-44 .06342-34 26...7576 25 L 25755 3"[email protected], 4 '.7 13 14283.2 6.83547E-44 2.62845 393152 4446.45 6°57 67 09,93 .1 43236 16764 H834. i 5628.67 86.6 8

i 83783 388 8 138248 8.24597E-03 .8142819 5.5025 r. 0327 8 0 8 8 8 0 2.0327 128421 85"349!2 V.5 12 8 4.5 .179337 i15. 845 4.5 7.4852E-d4 0 d 0 -157842 756983 519.6 4.652''81 82697.6 458 2 8 6 0 8 8 5278.8 0 17% 148.333 466.224 8 177.776 228219 1454i8E+* 8 0 8i 9 8 0 37.2882 L9972 2 28406 8 6.71431407 508.654 37529.4 i 593 6995.32 0 6 2E-03 4.56997E-84 8786471 26 KA 25 1.8!!81 4718.83 21 135 14118 6 6.83547E-84 3.21083 417323" 4667.37 688? 9 5391.5 .1 43137.2 16862 8 16986.7 5607.84 86.6 8

199334 308 8 136699 8.66725E-83 0143628 a 98082 23.1495 8 8 0 6 8 8 23.1495 21661 8548973 H. i L. 28 4.7 .199355 1!3.999 4.8 7 48806E-04 6 8 8 -157842 .756684 519.6 5.50411 89695 458 2 8 6 0 0 0 5823.06 6 1778 147.5 466.428 6 177.572 2163.31 4.84942E+06 8 0 .87 9 0 8 37.8788 23380 24143.7 0 7.38942E+87 499.847 37302.2 1.6638 8288.97 135 8 1-83 4.58997E-01 8945743 25.8363 25 836321 5674.46 n s 25 14 148.7°.1 6.83547E-04 3.80094 442511 4921.27 6727.67 5475 1 42876.1 17123.9 17146.8 5573.89 86.6 6 A10-10 Appendix 10-2 (cont'd)

2 !ills 381 A 133627 9. 87317E-13 914693 6.33345 24.20 1 1 1 1 1 6 24.2335 BON 8653/89 A 4 1152 4.8 .197174 WAN 4.9 7.48811E-14 1 1 1 -157842 .756679 W. 6 6.89911 94994 2 458 2 1 1 1 e 0 6381.88 0 1758 145.0 466.432 1 177. SO 214185 4. 5407E406 9 9 .9t 9 6 a 36. 0"519 2tv^i6.5 26678.4 0 7 775487E+07 499.316 37290 3 i. M2 0497 :8 14 a 2E-03 4.58997E-84 .189725 25 7288 25 72970'8 65Rs'. 5 .29 .145 139% 6 6 83547E-04 4. W . 461593 5875.48 6576.49 5476.43 i A?Pi 5 17'!28.4 172116 5573.31 86.6 0

2 230 380 8 138598 9.4907E-2 9158347 6.04631 25.3518 8 9 8 9 8 e 25 3518 L7'782 9w Na is 7 13.76 4.9 `9627' 112153 5 7.48925E-04 a 0 8 -157842 .7560'7 Sia.6 6.6m 188439 459 2 a e a a e X99137 a 1"r 1 144 ie7 466 4338 6 17 562 2124 36 4.0?i8E+at+ a 0 .87 9 0 0 336.233 288!16. 8 29386.5 0 8 &''85E+07 498.0!2 -NIn. 3 1-cW 1*819 SOS 0 2E-a3 4. 58997E-04 .1248'+0 11 fi 31 25 03087 7441. N 33 15 13496, 2 6.83547<-°X 4.0938 481184 SM 43 6427 031 5479 1 42963.5 iw& 5 17'11.7.3 557 .M 26 86.0 a

2 356: 3w 0 127588 9.9242' -03 8153904 7. e68S8 26.5057 a 0 0 0 0 a 26.5867 .119884 0579682 L 14 5 2150 W. 23 5.1 ; 4SM-04 0 0 e -157842 .756657 51.4.6 7.3'b 1668:3 458 2 a e e e e 7633.3 0 the 142 5 466.447 0 177.553 2104.96 5.40711E+06 8 e 87 9 e 0 35 8142 M&9 32M2 8 a 68M+e7 4905% 37"281.4 177 118* 4 15 9 2E-63 4.58997E-04 148898 25%42 25 56425 M89 s7 155 139777 b. 83547E-04 5.4347 581.284 520, A 60'9 29 5482 64 1 42852 2 17!47 8 17243, 9 5579. r8 86.6 0

440", 390 a 1261e8 0101104 0!557 7 2M 27 0645 7 0 9 9 0 9 0 27 *45 A22 05"1 19.3 14 24 5.1 204725 lie 367 5.1 7.4870-1-a4 9 0 8 -157942 7566% 519 6 7 8!375 198143 458 2 e e e e e 792416 a !798 141667 4% 42 0 177.58 2985 34 WAD* a a 87 9 a 9 ^. 6047 13"19 342751 a a 8797IE407 499 347 37311 i. 8864 L^43 7 15 a 2E-03 4 58997E-04 155041 25 5163 25 516272 93'82.48 41 155 13469 4 6 03547E-04 S. 0877W 598916 5435.95 6&X 06 M. 75 1 4M 2 17113.8 172 2 5575.21 6 0

A10-11 APPENDIX 11. GEO PLATFORM ANALYSIS (Baseline Configuration)

9 w CD D2 I'D DT DV D" E EE EM ri CE K1 KP KQ KS KT LD LF LM LO MH MY PI PH RM SP F6 cj T2 W TD TF TG TL TM TN TY TZ, WL WN 6 WT bi wu

lee 1 1 .014 1 36 1 3 2 .61 4E+07 W 300 .1 1 3.218 1 4E,* 15 4. 19 aN 19 .85 3. 06-^ 14159 r2 W4 ;.1-03 5E-03 A4 E --A'.3 01 .41 LK, -7E-N 10 0M000 12^; . 47 105w2

Hi 8 FIR FIC AIL AM AT L EA 65 66 B9 • su EE CI C2 t i D F1 Fr. R FD F1 Fwc I is 8 i, Q K5 K6 K7 K4 K .,E LH LH(FT) LB LC N, N Pi PK IL PW T Ti TR TE TP Tg TP - T T TX w 1,1B WP ww idV wz XB

766 0413 9 25292 6 *@3211 K7

2 i 64". 1. -004 2. 84:99 59.7736 SrM 404791 93. 1227 (08 426. 219 2 0 0 C4 .'69 63. -3333 5@3. 2C 'A. 12@- 1,4U. 7, 57 '.32 2VK'Z', 2 3flA22s_(f5 9 a 4 ", !4 226- F1 9579, 0 612 092 4WO. 7 4.53 FE-44 3. M. 21 77 25. Z"721 91 1+55 1 fti6% 7 3. 5?,,4?E-04 0 *52. 4% 4 52, 0. 542 `57974 464ft. 3 0190 1 1:,598. 7 5569. 19 t

53.766 097E 61789 056724 . 9 0 2. 1?1:45^ 2. 'w,478 6.8 0 i0 2 842129 9 2.1 12. 5164 ej. 0329 -12 99 -1. 55784 4W,51 91 1227 451 2, 66 2

K, 4 1% 5@4 IN. 466 :.:x±7. 13 6 _j ';. 'A 1. 87 9 0 2731E @ I-PR 48 9579, 13 516.__1193 472. 544 2E-03 4. WWE-04 1a ;9 2'9. 4_45 25 4 42449 21 e5 tf-5 14455. 1 53547E-0 0 21 x- 5P9, f 557974 C?+ 4 la A11-1 n 0- Appendix 11 (cor tl

.'.'dl A

d id 6.8 0 25.66* 51. 515.627 155794 49501 UY. 457. 7 51 2 0 075 4000. 2 437271 0 1274 105.833 595. 925 164. 6.2"1 17E, 46K. 360126 7, ?S. 19 2 507. 114 37M. 7 NP,4 1.4 4. 410j+7 K5 14'^71. 1 's, 7 41 483 16481. 9 OV 5657. 35 0

8 ObFi 53. 756 i-7!► 61514 YA I.' i&M H @ 4 9 0 .1.59244

77 6742 Im W182 1 1217 FK @ 460142 045 0 US. ON 14 416 177. 924 4. 13M-44 .87 a 548345 I? tip ?7 .708 :"49.24 95 9.13 5764. 59 U `933. 687 s "( 693. 7 C!) 4.58997E-04 0403203 26.6537 25 1 45372 w..C. .- 431 151Z 29 J55 WK I 3.5354?E-^4 B'dS 75 .557974 4320 16073. 9 16683 8 1?

2. =5ii" ^E 15114+4 .5 57 447

',M39

52 Hllt 04 7 22 4 14 4 -1 45^ 2 II 4 506 177 100.946 177 'rZ, 44.112 a 0 1129 ^j . 4. 71 6677 5. 76154E-r 4 37 9 1147 759~.90..".137 ^511q. 0 Sh M 37812 IVN 4. ",. 7tE-A4 M-35,2 X. 2!45 25 i. 545 74" 1 :Xl 41 c,;--,!A' -C44 g C155 i<,.42. 7 E 7* 52`12. 46 151474 42164. 9 1033. 2 i 5615. -'A

A 53. 5184 1!+1+583 71488. F 127455455 @ -c C^'; 7. CA 21 -I. K 69.1ja . 22+ +'Z 0 . 3213 12164-3 MI S. IS59 (4 458 A 0 0 0 3848. 7 2 I? lrfk 106.667 506.274 168.376 177. 726 Ct^44 87 9 0 747t 7434 tl`! i1157.4 9286.99 11157.4 `(+0.636 37472, 4734. 4. ". , 7E-@4 07*W X. 4078 its 466.425 14117. 5 153547E-04 0 4 0 At;AM 434114 16927 7 SIM. 7 5599.

d 3. -1" 53.21 U 1052 68176. 8 8 3.15922 ZIN267 0 0 0 0 0 V..Q W 7 AR 2, 6;, ^A W4 1:189 3* -1 440 458 IN. K4 177 3719. n 1 17 0 506. 46 826866 8 1270.2 I. i+^') KH4 87 9 9 1432.34 12W 9 10909. 8 12849.9 499. 847 8 `.t7 4 -'E-K 4. 3. 5--,5kl-04 401.*2 14;5 15 25 065 14030. 1 4 m *I i%4;, 68 All-2 Appendix 11 (cont, d) 4

53. !442 117%.8 Q?l 9 04 04

7.52 2 3 100 a 2.4 0259429 73.8205 147,641 1W. 99 -155783 !438: 1. K" 9 310%9 456 2 0 0 0 0 0 i_7 !A4 0 UM i@@ W 479 122 1 2? ??E+* 93f&44 147 9 r W.57

!^N 5 14457 1 499.* 4.4, k44" 15:9. 4 1 *5 4.58997E-04 109972 215.11 1,42 3 %VIS *'598.:5 5 44819 Is 132Z 5495.2'S SK 8 Q1788. 1 17187 3 17_119 5545. 65

52. 7777 0 56M. 8 W. 10 0 0 4.6424 11 WS7 0 0 4. .14." 421 7 11. 2 7.76 2. 4 11427: A 75. A-24 151.615 i&%I -1. 55783 5046K '.'4'1 416? 11 8 a a 178.484 `496. 465 107. 616 17l- K-15 -4" 1. a E 45:"4-4 87 4 13:22. 4 165N, 448. 906 7201 2 As 1421. 93

d-43 4 :a ;KH4 LIr.*'3 25. 6317 1.5 1"A6

3. "-411-04 0 c47 ".4.'3 5490. Lql W.4'"22 4277::. 5 171,70.9 17227.5 567 '8 4

K 5=15 Im 547313.5 OW! 0 it 4. 147 8 0 a 0 0 0 4, 22975 11 5 8 4 Ok K 9454 164 10.67 -1557'83 5009!2 90. qw 42.^515 3 458 2 @ 0 0 @ a V4. -., 66 11 M-2 3- 13 W 469 VO. S'4 V7. SM 3133. 7 1 W.,2E+"' t I K06 -A6 1 06390E-03 87 9 O84 4 17371 4 8 140.4 111 N4. 8 498.591 37256. 8 885 17 0. 2E-83 4. %9^41-44 1406 iL5 5433 25 4. s7 er'S t1977 6 I 53c47FE-A4 49tt^t+4 M. '106 11!2. 5491. 60 42756.9 17176.1 17243. 1 5507.11 a

. I K 1 7C.Le ii 4 4V5.4 01451 0 4141A, C.- !. 24:14 0 2. 7 87. 6972 175. 394 1110. 04 -1.55783 542?01 1588 4Kl-'6 .5

458 211. -A .7 N 111+0 w 710 7 SC16 47 L ! 3 I IPNE-Ci 47 %) I 18i71. 7 44. 338 37254.1 6-; 4, %!..E47 -N 2E-03 15,A417 ;.5. c 25'Ir 508 9'. . th

. 7 21*9 2 1 3N27E-04 a 511168 21 4S2 1111 4 ri-i71 5491 QC. 741 4 1'1N.i I q. 172C8 6 5156" 0 1)

All-3 m 0 APPENDIX 12

DISTRIBUTED THRUST ANALYSIS

The effect of distributed thrust on dynamic loading was examined. Distributed thrust results in decreased dynamic loading over the center thrusted case, as long as the thrusters are in phase. If they are slightly out of phase, however, dynamic loading is increased. Since exact phasing of the thrusters is not probable, this situation was examined.

The computer program examines the dynamic factor for a typical OTV-LSS system. Parameters that can be varied are: frequencies of the system response; weights of the OTV, propellant, and payload; size of the payload; engine rise time; and time lag between thrusters. Due to the method in which the problem was modeled, the location of the thruster must stay in the region of 20% of the radius of the LSS. Therefore, this program can only be used for the purpose of varying the lag time and determining its effect on FD. It can not be used to compare the effect of different thruster locations. The problem was modeled as follows:

At any given instant we have

rS r2^

This results from the thruster on the right firing an instant before the left thruster (therefore, it has a higher thrust level). This situation occurs only at times 0 to - the rise time of the engines. After that time both are at equal thrust levels of one- half the total thrust. This situation can be divided into two cases:

SYMMETRIC ANTISYMMETRIC

FL FL * FR FL t FR FR - FL FR - FL FR 2 2 2 2

Al2-1 Superimposing above cases gives:

left side: FL + FR _ FR - FL R F 2 2 L

right side: FL+ FR + FR - FL F 2 2 R

For each case (symmetric and antisymmetric), two modes are modeled: rigid-body mode and one elastic mode.

Modes:

symmetric, rigid-body

antisymmetric, rigid-body

symmetric, elastic

antisymmetric, elastic

FR + FL symmetrical force: F _ 2

antisymmetric force: F = FR - FL 2

Al2-2 n 0 FORC]

SYMMETI

THR

ANTISYI! METRICAL: F = FR - FL

THRUST

t - lag time

TIME

Al2-3 ANTISYMMETRIC CASE:

f I JJ^

Mode "1" Mode "2" q l - ref. dell. due to Mode 1 q2 - ref. defl. due to Mode 2 X i = total defl. of M

Xi Oil q l + 0 12 q2 F i - external force applied to M

Equations of note as in matrix form are: 0 jX1 ^ rM11 j X1 + r1lk1?. F1' Ix 2 IF 0M22Vi 2 21 k22 1 2 1

In shorthand this is written: [M] f X1 + [k] f X} = f F} (1) Let = [^] r1l 0 12 Model matrix (1A)

21 022 where Q ij - modal displacement of M i in mode j w 2j r w 11 0 frequency matrix

0 cv 222 Let 1XI _ [0) {q j (2) substitute (2) into (1)

[M] [¢] fq} + [k] 101 tq} _ {F}

Al2-4 Premultiply thru by [Q]T

where [¢] T _ transpose of X11 021

1012 022 So that

[o] T [M] [9] jqj + [o]T [k] [9];qj _ [9] T {F} It can he proved that

[Q] T [M] [q] _ [ rq ] a diagonal matrix (2A) [9] T [k] [¢] _ ^ a diagonal matrix l n rqj ^I { q } {^} where + {F} _ [O] T {F} (3) -1 Premultiply by 71)

61 + [ 7, J -1 {q} _ ^ l 1 { } [k] It can be proved that L ^1 1 [I] [W] = ^wj w diagonal matrix

whence = ^ 1 0 W2y 0 W2

W i Nat-freq in mode i where + [W ] jqj _ [7j -1 141 1.1 (4)

Eq. 4 can be written out in detail as follows: 3 Mode "1" Eq: y1 + W2 q 1 1 <0 (5) 711 „2„ 2 Mode Eq : + w2 q2 (6)

= 7122

3 i "generalized force" in Mode i (find from Eq (3) q i = "generalized coordinate" in Mode i

Solve Eqs (5) and (6) for q i , q 2 <6

Use Eq (2) to get X 1 and X,, ^,t>.

Al2-5 n • Foregoing has assumed and [0] are known. tw2j ( 1) For mode 0i1 is e. linear function of r.

'For mode ( 2) appl y the boundary condition that d'Alembert moment about pivot point (i.e., CL ) is zero. i.e., r i 0 (7) i M 012 Consider a two-mass beam, as follows:

Eq. (7) gives:

M1 r l r 2 = 0 012 + M-4

M r 012 2 2 T _ M— r (8) 1 1 i.e., and MUST be of of opposite sign for this 012 022 particular case.

Similarly for Mode 1

42 1 _ mtt .

-^t

Al2-6

From geometry of rigid-body motion

0 11 rl (9) 3 r2_1 r2

Let 021 = 1 i.e., ref, values at tip: 022 = 1 Then, from Eqs ( 1A), (8) and ( 9) we get

[0] = r1 - M2r2 (10) girl r21 1 For chosen values of r /r 2 , M /M2 Eq (10) gives [0]. This is exact for a massless biam cariyi g g two concentrated masses as shown.

An eigen value solution is needed for more complicated structures. rw 2j Now consider W2 0 0 W2

wl u rigid body mode = 0 w2 a elastic mode = any chosen value.

Note: Period = T = LW seconds. So we may say in summary:

W 21 = 0 0 ^ y (1;_)2 0 1

[ 0] r i'l 1.2 - M2 r2 M 1 r (12)

1 1

Al2-7

I I {X^ = CO] {y} (12A)

1 1 I Ff 31 = to] T ri/r2 1 F 3 2 , -M 2 r2 ^ 1 M1 r1 1 t F2 Let F1

F2 (t) 0

T; an r 3 1 (t) = r1 Fl (t) + t ) (13) 2 0F 2 (

+ F! 2

Mass matrix is

[7t] = M, 0 ` (15) 0 M2

Substitute ( 1.5) and (12) into (2A) to get generalized mass matrix, thus: _ [ = r l M1 0 -M,,r2 711 r1 M r r 2 1 1.2 1 1

-M2r2 1 0 M2 1 1 ^i r L 11

l M` rl M r2 Ml r F2 r2 Mlrl

-M2 r 2 M2 1 1 rl J L

Al2-8 ..

bi r12 1 ?d2 r22 \-M2 + M2/

M22 r22 + bi2 \-M2 + M2 l M1r1 2

2 M2 IM 1 ^ r12 + 1 ; 0 M2 r2 nl ^ Ai2r22 (16) Ai2 2 + 1 0 M1rl

Comparing (12) and (16) we can see that (16) can be written

2 O il + M2 021 2 ,; 0 (Ml (17) 0 (M1 012 2 + M2 022 2

Eqs (11), (12), (13), (14) and (171 give values of [W 2] ^^, , [3] and [7] in term of Ai l , M2 , F1 & w2

They can be found numerically.

Substituting in (5) and (6) we have two linear differential equations in q 1

t = 0 ql t\ = 0 ql ,;t) = 0 q2 0 q2 fit) = 0

Solve for ql ^^t; and q 2 (tj Use Eq (12A) to getX1 fit;) and X2

Al2-9

SYMMETRIC CASE:

m^ m2

t 021 - - 71 mT T

rigid mode elastic mode

For rigid-body mode (by inspection)

011 - 1 021 = 1

for elastic mode there must be no resultant force .'

B.C. is M1 0 12 + 112 022 = 0

M2 X12 022 M1

1 [^] = -M2 M1

1 1

C^] T [M] C ^]

[^] = 1 _M2 M1

1 1

[ M] = M1 0

0 M2

Al2-10 1 + M2 0

11`7 a 2 0 M _'2 + M2 M1

[0] T 3 1L I Fj^ :' 2

-M2 T F F1 M Force on M 1

M1 l 1

I j F2 Force on 612 F2 = = 0

F, + F/ 0

-612 F 1 + V

0 2 IM 1

FREQUENCIES Mode I (Rigid) Mode 2 (Elastic)

SYM w1 = 0 wn

Anti Sym w 1.5 X 2 = 0 W9

0 11 q R accel mass 1 + 0 12 q e = q = On R + 032 q e accel mass 2

y ,` Compute for both s mmetric and anti symme tric case

add N I (Sym) + X1 (antisym) for accel of 611 X,) (sym) + t 2 (antisym) for accel of 612 Differential eq's solved by Laplace Transforms. a '. Al2-11 Distributed Thrust Computer Program 10 IMAGE D. D,'::, D. DD.7,•t. D. DD 20 INTEGER I,J 30 SHORT Friel? 1 F1i:61), Fit %61),FaCol),F.; k_.o1),Fgsr%i 1.,Fplalli1:,F^ar..e•1: 40 SHORT Fgat l.61),Axi ,'61),Ax2(iI).' 4' 4? asi 4 ?,E•a+:4),Bst4:,S(A),B<4.' 50 SHORT A ;csl^: ol'.Axi2c61 ) , Aaal <61?.r.;;a2 .Qi:,Mg11s , Ma2s , M^11a,My2+^ e0 SHORT P11s,P12s,Ws,P11a,P12a,Wa 70 COM Laq, Tau. INTEGER I1, I2, I3 80 L=200 -- 90 Motu=6000'3 2.2 1\ ^^ C Try 11 100 Mprop =0 --- 110 Mp1=16000,32.2 J 1 120 R1=.20*L 13 0 R2=.75*L 140 M1-.5*(M0tv+Mpr0p7'+.25*Mp1 150 M2=.25*Mpl 160 Wi=.5*2*3.14159 170 W•a=1.5+Ws 180 Pils=1 190 P125=-M2/M1 200 P213=1 211 P223=1 220 PII&=R1•R2 '? 3 1) Plaa= - M2*R2-' < M1*R1) 240 F21a=1 250 P22&=1 26 3 Fm&X =1600 270 REM DIVIDE BY TWO SINCE MA., PER THPUSTER IS HALF THAT: 280 Fina>f.-Fft,ax /_ 240 Mg11siM1 +M2 310 M922s =M2^2: M1 +M2 310 M9I1a=M14P1I1-2+M2-P21a^2 • 320 Mg22a=M1*P12a,2+M2•P'22a^_ 330 REM 340 _ _FOP. Tau=. 3 TO .7 STEP .20 350 - REM 360 FOR Laq-0 TO .20 STEP .1 370 REM 330 FOR T-0 TO 6 STEP .1 390 I=10*T 400 F=Finax 410 IF T =T iu THEN F-Fm3;:`T.au*T 420 Fr^Ii=F 430 J-I-104Laq 440 FI(I)=0 450 IF T Laq THEN GOTO 470 460 Fi(I)=Fr(J) 470 Fs(I)=CFr(I)+F1(I)?,2 4+^ tiJ F3C1)^CFr;s1(I)):: te id^:;<:A;::az ::I%*A:^s2

Al2-14 APPENDIX 13

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