Communications Satellite Systems Operations with the Space Station

NASA CR179527

COMMUNICATIONS SATELLITE SYSTEMS OPERATIONS WITH THE SPACE STATION

VOLUME II- TECHNICAL REPORT FEBRUARY 1987

NASA Lewis Research Center Contract No.NAS3-24253

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(_ASA-C_- 179527) CCM_UNICA_IC_Z 5 AT_LLITE _87- 174_3 SYSTEMS OP£RA216N_ _I_H _SE S_C£ S IATICN, _£LUME 2 P inal Eeport, Jul. 1965 - Sep. I&65 (Ford Aezo_Face and Communications Unclas Corp.) 150 p CSCL 12B G3/6b q375; NASA Contract No. NAS3-24253

COMMUNICATIONS SATELLITE SYSTEMS OPERATIONS WITH THE SPACE STATION

Volume II: TECHNICAL REPORT

February 1987

Prepared by

FORD AEROSPACE & COMMUNICATIONS CORPORATION Western Development Laboratory Division Spacecraft Systems Operation, Advanced Systems Department 3939 Fabian Way, Palo Alto CA 94303

COOPERS & LYBRAND 1800 M St. NW, Washington DC 20036

Program Manager: Kent M. Price

Major Contributors: John Dixson John Egan (C&L) Michael Kanfer (C&L) F. L. Limburg Larry Spanier Chuck Weyandt Steve Wright

Prepared for NASA, Lewis Research Center Cleveland, OH 44135 ABSTRACT

The NASA Space Station has the potential to provide significant economic benefits to commercial communications satemte operators. This report gives the results of a study to quantify the benefits of new space-based activities and to assess the impacts on the satellite design and the Space Station. The following study results are described:

• A financial model is developed which describes quantitatively the economics of the space segment of communication satellite systems. The model describes the economic status of the system throughout the lifetime of the satellite. The economic performance is output in terms of total capital cost and rate of return on investment.

• The expected state-of-the-art status of communications satellite systems and operations beginning service in 1995 is assessed and described. The results of the assessment are utilized to postulate and describe representative satellite systems.

• New or enhanced space-based activities and associated satellite system designs that have the potential to achieve future communications satellite operations in geostationary orbit with improved economic performance are postulated and defined. These activities include retrieval, orbital transfer vehicle (OTV) launch, deployment of appendages, checkout, fueling, assembly, and servicing of satellites. The financial model is used to determine the economic performance of these different activities and combinations of activities. The use of the space-based OTV to transport satellites from low earth orbit to geostationary orbit offers the greatest economic benefit.

• Three scenarios using combinations of space-based activities are analyzed: (1) a spin stabilized satellite, (2) a three axis satellite, and (3) assembly at the Space Station and GEO servicing. The economic performance of the scenarios is analyzed.

• Functional and technical requirements placed on the Space Station by the scenarios are detailed. Requirements on the satellites are also listed.

The major study results are as follows:

1. Economic benefits are realizable for the commercial communications satellite industry with use of the Space Station.

/ 2. A space-based OTV is necessary to carry out APOs in a timely and cost-effective manner.

3. A study of the economics of retrieval missions and the influence of retrieval on the insurance industry is required in order to ac_urately demonstrate the value of retrievability for the satellite.

. Further NASA-sponsored study of a modular satellite design capable of being assembled in LEO (at the Space Station) and serviced in GEO is required.

. Sp (_+._+'_ 1_._1 ...... "._.-1 -¢'.... +^11:+^ " " _1.^..1..1 T.^ :_+_11^_1 .... :1..1^ to demonstrate NASA commitment. TABLE of CONTENTS

SECTION I - INTRODUCTION

Background I- 1

Objectives of Study I-1

Approach I- 1 3.1 Task 1. Develop Communications Satellite Financial Model ...... I- 1 3.1.1 Task 1A: Develop Basic Financial Model ...... I- 1 3.1.2 Task 1B: Assess Impact of System Characteristics on Financial Model Output I- 2 3.2 Task 2. Determine Economic Performance of Business-as-Usual Scenario for 1995 . . I- 2 3.3 Task 3. Assess Economics of New Space-Based Activities for 1995 ...... I - 2 3.3.1 Task 3A: Postulate New Space-Based APOs ...... I- 2 3.3.2 Task 3B: Evaluate Economics of APOs ...... I- 2 3.4 Task 4. Develop Space Station Scenarios and Requirements ...... I- 2

4 Organization of Report I - 3

SECTION II - FINANCIAL MODEL

Introduction II- 1

Model Assumptions II - 1 2.1 System Characteristics ...... II - 1 2.1.1 C-Band Satellite ...... II- 1 2.1.2 Ku-Band Satellite ...... II-4 2.1.3 Hybrid Satellite ...... II- 4 2.2 Capital Expenditures ...... II - 4 2.2.1 Satellite Costs ...... II - 7 2.2.2 Perigee Stage Costs ...... II- 7 2.2.3 STS Launch Costs ...... II-8 2.2.4 Launch and Mission Operations ...... II- 8 2.2.5 Launch Insurance ...... II- 8 2.2.6 Payment Schedules ...... II- 8 2.3 Revenues ...... II - 9 2.3.1 Characteristics of Transponder ...... II - 9 2.3.2 Degradation of Satellite ...... II - 10 2.3.3 Market Factors ...... II- 10 2.3.4 Price of Transponder ...... II - 11 2.3.5 Utilization of Transponder ...... II - 11 2.4 Operating Expenditures ...... II - 11 2.4.1 Life Insurance ...... II- 12 2.4.2 Other Expenditures ...... II- 12 2.4.3 Rate of Inflation ...... II- 12 2.4.4 TT&C ...... II- 12 2.5 Financing Assumptions ...... II- 12 2.5.1 Balance Sheet Financing ...... II- 12 2.5.2 Period of Financing ...... If- 12 2.5.3 InterestRates ...... II-12

2,6 Tax Assumptions ...... II- 13 2.6.1 Investment Tax Credit ...... II- 13 2.0.2 Depredation ...... if- i3 2.6.3 Marginal Tax Rate ...... II- 13

3 Financial Analysis II - 13 3.1 Net Present Value ...... II- 13 3.2 InternalKate of Return ...... II- 14 3.3 Dual Terminal Rate of Return ...... If- 14 3.4 Other FinancialEvaluations ...... If- 15

4 Model User Manual II - 15 4.1 Introduction ...... II- 15 4.2 Accessing the Model ...... II- 15 4.2.1 Main Menu ...... II-15 4.3 Model Inputs ...... II- 15 4.3.1 General Assumptions Menu ...... II- 16 4.3.2 Financing and Tax Menu ...... II- 16 4.3.3 Capital Expenditures and Operating Assumptions Menu ...... If- 17 4.4 RecalculateSpreadsheet ...... II- 17 4.5 Printthe Spreadsheet ...... II- 17 4.5.1 Model Output ...... II-17 4.5.2 Print Options ...... II- 18 4.6 Save and Exit Spreadsheet ...... II- 18 4.7 Use of Model ...... II- 18

Model Outputs for 1985 Satellites II - 19 5.1 C-band Satellite ...... II- 19 5.2 Ku-band Satellite...... If-19 5.3 Hybrid Satellite...... _...... II- 19 5.4 Discussion...... II- 20

SECTION III - IMPACT OF SYSTEM CHARACTERISTICS ON ECONOMICS

1 Introduction III- 1

2 Choice of Characteristics III- 1 2.1 Channel Capacity ...... Ill- 1 2.2 SatelliteLink Equation ...... Ill- 1 2.2.1 Downlink Communications ...... Ill- 2

°.* 111 2.2.2 Uplink Communications ...... III- 2 2.2.3 Technical Characteristics ...... III-2 2.3 Atmospheric Attenuation ...... III- 3 2.4 Regulatory Considerations ...... III- 3 2.4.1 Interference ...... III- 3 2.4.2 Ground Antenna Sidelobes ...... III-3 2.4.3 Power Flux Density Limits ...... III-3

Primary Characteristics III- 4 3.1 Communications Capacity ...... III- 5 3.2 Bandwidth of Channel ...... III- 5 3.2..1 C-Band Channel Bandwidth ...... III- 5 3.2.2 Ku-Band Channel Bandwidth ...... III- 6 3.3 Frequency Band of Operation ...... III- 6 3.4 Power Transmitted in Channel ...... III- 6 3.4.1 Impact on Satellite ...... III- 7 3.4.2 EIttP ...... IH- 7 3.5 Utilization ...... III- 7

4 Secondary Characteristics III - 8 4.1 Attitude Control and Pointing ...... III- 8 4.2 Coverage Area ...... III- 8 4.3 Ground Terminal Requirements ...... III- 8 4.3.1 C-Band Ground Terminals ...... III-9 4.3.2 Ku-Band Ground Terminals ...... : . . .III- 9 4.4 Lifetime ...... III - 9 4.5 Mass and Volume ...... III- 10 4.6 Maturity of Technology ...... III- 10 4.7 Noise Temperature of Receiver ...... III- 10 4.8 On-board Switching ...... III-11 4.9 Primary Power ...... III- 11 4.10 Reliability ...... III- 11 4.11 Time Delay at Launch ...... III- 12 4.12 Transponder Type ...... III- 12

Financial Characteristics III- 13 5.1 Debt Ratio ...... III- 13 5.2 Insurance ...... III - 13 5.3 Interest Rate ...... III- 14 5.4 Period of Financing ...... III- 14 5.5 Taxes ...... III- 15 5.5.1 Depreciation ...... III- 15 5.5.2 Investment Tax Credit ...... III- 15 5.5.3 Loss Carry Forward ...... III- 15

6 Discussion III - 15

iv SECTION IV - TECHNOLOGY FORECAST FOR 1995

1 Introduction IV- 1

2 Technology Assessment IV - 1 2.1 AttitudeControl Subsystem ...... IV- 1 2.2 Communications Payload ...... IV - 2 2.2.1 Antennas ...... IV - 2 2.2.2 Transponders and Receivers ...... IV- 2 2.2.3 On-Board Processing...... IV- 3 2.3 Primary Power ...... IV- 3 2.3.1 Batteries ...... IV - 3 2.3.2 SolarCells ...... IV - 3 2.4 PropulsionSubsystem ...... IV - 3 2.5 Structureand Mechanisms ...... IV- 4 2.6 TT&C Subsystem ...... IV - 4 2.7 Thermal Control ...... IV - 4 2.8 Space Transportation...... IV - 5 2.9 Summary of Technology Developments ...... IV - 5

Description of Three FSS Systems for 1995 IV - 5 3.1 Spinner Satellite System ...... IV- 5 3.1.1 Ku-Band FSS Payload ...... IV-9 3.1.2 Other Design Features ...... IV-9 3.2 Ku-Band Satellite System ...... IV- 9 3.2.1 Ku-Band FSS Payload ...... IV-9 3.2.2 Other Design Features ...... IV- 9 3.3 Hybrid Satellite System ...... IV- 12 3.3.1 C-Band FSS Payload ...... IV- 12 3.3.2 Ku-Band FSS Payload ...... IV-12 3.3.3 Other Design Features ...... IV- 12 3.4 Comparison of Systems ...... IV - 12

4 Large Satellite System IV - 12 4.1 Satellite Design ...... IV- 12 4.2 Payload ...... IV - 13 4.3 Launch Vehicle ...... )'. IV - 13

SECTION V - BASELINE ECONOMIC PERFORMANCE

1 Introduction V-!

2 Methodology V-1

3 Ku-band Spinner System V-2

v 4 Ku-Band 3-Axis System V - 2

5 Hybrid 3°Axis System V - 2

6 Large Satellite System V- 2

7 Discussion V - 3

SECTION VI - NEW SPACE-BASED ACTIVITIES

1 Introduction VI- 1

2 Postulation of APOs VI - 1 2.1 Emergency Retrieval ...... VI-2 2.1.1 Scenarios ...... _ ...... VI- 2 2.1.2 Requirements on Satellite...... VI-3 2.1.3 Requirements on Shuttle,Space Stationand OTV ...... VI-3 2.2 Ground-Based OTV Launch to GeostationaryTransferOrbit ...... VI-3 2.2.1 Scenario ...... VI- 4 2.2.2 Requirements on Satellite...... VI-4 2.2.3 Requirements on Shuttleand OTV ...... VI-4 2.3 Ground-Based OTV Launch to Geosynchronous Orbit ...... VI-4 2.3.1 Scenario ...... VI- 5 2.3.2 Requirements on Satellite ...... VI-5 2.3.3 Requirements on Shuttle and OTV ...... VI-5 2.4 Deployment of Appendages at Shuttle ...... VI-5 2.4.1 Scenario ...... VI- 5 2.4.2 Requirements on Satellite ...... VI-5 2.4.3 Requirements on Shuttle and OTV ...... VI-6 2.5 Space-Based OTV Launch to Geostationary Transfer Orbit ...... VI-6 2.5.1 Scenario ...... VI- 6 2.5.2 Requirements on Satellite ...... VI-6 2.5.3 Requirements on Space Station ...... VI-6 2.6 Space-Based OTV Launch to Geosynchronous Orbit ...... VI-6 2.7 Deployment of Appendages at Space Station ...... VI-6 2.7.1 Scenario ...... VI-7 2.7.2 Requirements on Satellite...... VI-7 2.7.3 Requirements on Space Station ...... VI- 7 2.8 Checkout at Space Station...... VI-7 2.8.1 Scenario ...... VI-7 2.8.2 Requirements on Satellite...... VI-8 2.8.3 Requirements on Space Station ...... VI-8 2.9 Fuelingat Space Station ...... VI-8 2.9.1 Scenario ...... VI- 8 2.9.2 Requirements on Satellite...... VI-8 2.9.3 Requirements on Space Station ...... VI-8 2.10 Assembly of Satelliteat Space Station ...... VI - 9

vi 2.10.1 Scenario ...... VI- 9 2.10.2 Requirements on Satellite ...... VI-9 2.10.3 Requirements on Space Station ...... VI-9 2.11 Servidng for GEO Satellites ...... VI- 10 2.11.1 Scenarios ...... VI- 10 2.11.2 Requirements on Satellite ...... VI- 10 2.11.3 Requirements on OTV and Space Station - Retrieval to LEO ...... VI- 11 2.11.4 Requirements on Servicer Used in GEO ...... VI- 11

3 Impact on Satellite Design VI - 11 3.1 Retrieval APOs ...... VI- 11 3.2 OTV APOs ...... VI- 13 3.2.1 Spinner Design ...... VI-13 3.2.2 3-Axis Design ...... VI- 13 3.3 Deploy Appendage APOs ...... VI- 13 3.4 Checkout APOs ...... VI- 14 3.5 Fueling APOs ...... VI- 14 3.6 Assembly APOs ...... VI- 14 3.7 Servicing APOs ...... VI- 14

4 Discussion VI- 14 4.1 APO Timelines ...... VI- 14 4.2 Combinations of APOs ...... VI- 15

SECTION VII- ECONOMICS OF APOs

1 Introduction VII- 1

2 Methodology VII- 1 2.1 APO Values ...... VII- 1 2.2 Launch Insurance ...... VII- 1 2.2.1 Launch Insurance Rates in 1995 ...... VII- 1 2.2.2 Incidents of Satellite Failure ...... VII - 2 2.2.3 Influence of APOs on Rates ...... VII- 2 2.3 Launch Costs ...... VII - 3 2.3.1 Shuttle to LEO ...... VII- 3 2.3.2 Integral Upper Stage ...... VII- 3 2.3.3 Orbital Transfer Vehicle ...... VII- 3 2.3.4 Comparison of Launch Costs ...... VII - 3 2.3.5 Time for OTV Launch ...... VII- 4 2.3.6 OTV Performance ...... VII- 4

Spinner APO VRI,,pq VII - 5 3.1 Spinner APOs at Shuttle ...... VII-5 3.1.1 LEO Retrieval Capability ...... VII- 5 3.1.2 GB-OTV Launch to GTO ...... VII-5 3.2 Spinner APOs at Station ...... VII-5

vii 3.2.1 LEO Retrieval Capability ...... VII- 5 3.2.2 SB-OTV Launch to GTO ...... VII-6 3.2.3 Checkout at Station ...... VII- 6 3.2.4 Fueling at Station ...... VII- 7 3.2.5 GEO Retrieval Capability ...... VII- 7 3.2.6 Combination of Spinner APOs ...... VII - 7 3.3 Discussion of Spinner APOs ...... VII- 7

4 3-Axis APO Values VII - 7 4.1 3-Axis APOs at Shuttle ...... VII- 7 4.1.1 LEO RetrievalCapability ...... VII-7 4.1.2 GB-OTV Launch to GEO ...... VII- 7 4.1.3 Deploy Appendages ...... VII- 9 4.1.4 3-Axis Combination at Shuttle ...... VII-9 4.2 3-Axis APOs at Station ...... VII-9 4.2.1 LEO RetrievalCapabilityat Station ...... VII- 9 4.2.2 SB-OTV Launch to GEO ...... VII-10 4.2.3 Deploy Appendages at Station ...... VII- 10 4.2.4 Checkout at Station ...... VII- 10 4.2.5 Fuelingat Station ...... VII- 10 4.2.6 GEO Retrieval ...... VII- 10 4.2.7 Combination of 3-Axis APOs ...... VII - 10 4.2.8 Assembly at Space Station ...... VII- 11 4.2.9 GEO Servicing ...... VII-11 4.3 Discussionof 3-Axis APOs ...... VII- 11

Retrieval Mission Economics VII- 12 5.1 Insurance of RetrievalMissions ...... VII - 12 5.2 ShuttleMissionsin LEO ...... VII - 12 5.2.1 Repair at the Shuttle...... VII- 12 5.2.2 Repair on Earth ...... VII- 13 5.3 Missionswith Space Station ...... VII- 13 5.3.1 LEO Retrieval & Repair at Station ...... VII- 13 5.3.2 LEO Retrieval & Return-to-Earth for Repair ...... VII- 14 5.3.3 GEO Retrieval & Repair at Station ...... VII- 14 5.4 Summary of Retrieval Missions ...... VII- 15

6 Hectosat Economics VII- 15 ,)k 6.1 Launch Costs ...... VII- 15 6.2 Economic Performance ...... VII- 15

7 Conclusion VII- 18

SECTION VIII - SPACE STATION SCENARIOS

1 Introduction VIII- 1

°,. Vlll 2 Spinner SatelliteScenario VIII - 1 2.1 Description ...... VIII- i 2.2 Economic Evaluation ...... VIII- 2

3-Axis SatelliteScenario VIII - 2 3.1 Description ...... VIII- 2 3.2 Economic Evaluation ...... VIII- 3

4 Assembly/Servicing Scenario VIII - 3 4.1 Description ...... VIII - 3 4.2 Economic Evaluation ...... VIII- 4

5 Sensitivity Analysis VIII - 4 5.1 Introduction ...... VIII- 4 5.2 Launch Insurance ...... VIII-4 5.3 Launch Costs ...... VIII-5 5.4 Use of Mass Savings to Increase Payload ...... VIII - 5

6 Conclusions VIII - 6

SECTION IX - SPACE STATION REQUIREMENTS

1 Introduction IX- 1

2 Space Station Requirements IX- 1 2.1 Servicing and Storage Bay ...... IX - 1 2.2 Automated Transfer Facilities ...... IX - 1 2.3 Fueling Facilities ...... IX - 2

3 OMV Requirements IX - 2

4 OTV Requirements IX- 2

5 Operations and Policy IX - 2 ti _ummary of Requirements IX - 2 6.1 Spinner/3-axis Scenarios ...... IX - 3 6.2 Assembly/Servicing Scenario ...... IX - 3 6.3 LEO Retrieval/Repair Scenario Requirements ...... IX- 4

SECTION X- RECOMMENDATIONS

1 Introduction X-1

2 Need for Space-Based OTV X-1

3 Study of Retrieval Missions X-1

ix 4 Study of Modular Satellite Design X-2

5 Study of ELV Use X-2

6 Technology Developments X-2

7 Purpose of Space Station X-2

APPENDIX A - FINANCIAL MODEL OUTPUTS FOR 1985 SYSTEMS

APPENDIX B - FINANCIAL MODEL OUTPUTS FOR 1995 SYSTEMS

X LIST of FIGURES

II-1 Main Menu Of Financial Model ...... II- 15 II-2 General Assumptions Menu 1 ...... II- 16 II-3 General Assumptions Menu 2 ...... II- 16 II-4 Financing and Tax Menu ...... II- 17 II-5 Capital & Operating Menu ...... II- 17

°oooooo°oo°o°°ooooooooooooooooooooooooooooooooooo W-1 Ku-Band FSS Coverage Regions ...... W-9 _-2 Ku-Band FSS Coverage Regions ...... W- 13

°ooo°°oo°oo°oooooo°oo_ooooo°°ooo°°ooo°°oooo°°oo°o VH-1 Business-as-Usual Launch ...... VII- 11 VII-2 One Satellite OTV Launch ...... VII-11 VH-3 Two Satellite OTV Launch ...... VII- 11 VII-4 Schematic for LEO Retrieval Costs ...... VII - 16 VII-5 Schematic for GEO Retrieval Costs ...... VII- 17

xi LIST of TABLES

I-i Organization of Report ...... I - 3

• • , • • • • • • • • • • • • • • , • • ° • • • • • • • , • • ° , • • • • • ° • • • • ° ° • • • . II-1 Comparison of 1985 Satellite Characteristics ...... II- 2 II-2 C-Band Spinner Satellite Characteristics (1985) ...... II- 3 II-3 Ku-Band Satellite Characteristics (1985) ...... II-5 II-4 Hybrid Satellite Characteristics (1985) ...... II- 6 II-5 1985 Satellite Costs ...... II- 7 II-6 Payment Schedule, %/yr ...... II- 9 II-7 Satellite Degradation Curves ...... II- 10 II-8 C-Band Transponder Prices ...... II-11 II-9 Prime Interest Rates ...... II- 13

• • , * • • • . • • • . • • , • • • • • • . • ° • • • • ° • • • . ° • ° • • • • ° • • ° ° • • • ° III-1 Ku-Band Weather Margins ...... III- 3 III-2 Ground Antenna Standards ...... III- 4 III-3 Capacity Versus Cost ...... III- 5 III-4 Transponder Price versus Power ...... III- 7 III-5 Effect of Power Reduction on Mass ...... III - 7 III-6 Utilization Versus Rate-of-Return ...... III- 8 HI-7 Satellite Lifetime Versus Mass ...... III- 9 III-8 Rate-of-Return Versus Lifetime ...... III- 10 III-9 Price of Interconnectivity ...... III- 11 III-10 Payload Power Requirements ...... III- 11 III-11 Transponder Price ...... III- 12 III-12 Effect of Launch Delay ...... Ill- 12 III-13 SSPA Transponder Capacity ...... III- 12 III-14 Return Versus Debt Ratio ...... III-13 III-15 Return Versus Insurance Cost ...... III- 14 IH-16 Return Versus Interest Rate ...... III- 14 III-17 Return Versus Financing Period ...... III- 14

• • • • • • • • • • • • • • • • • , • • • • ° ° ° ° • • • • ° ° ° . ° • , ° • • ° * ° ° • • ° ° • IV-I TWTA Efficiency ...... IV- 2 IV-2 Battery Comparison ...... IV- 3 IV-3 FSS Systems ...... IV-5 IV-4 Available Orbital Transfer Rockets ...... IV- 6 IV-5 Satellite Technology Developments ...... IV- 6 IV-6 Summary of 1995 Satellite Characteristics ...... IV- 7 IV-7 Spin-Stabilized Satellite Characteristics (1995) ...... IV - 8 IV-8 Ku-Band Predicted Antenna Gain ...... IV- 9 IV-9 Ku-Band Satellite Characteristics (1995) ...... IV - 10 IV-10 Hybrid Satellite Characteristics (1995) ...... IV-11 IV-ll C-Band Antenna Predicted Gain ...... IV- 12 IV-12 Power Requirements for Beams ...... IV- 13 IV-13 Large Satellite Characteristics (1995) ...... IV- 14

xii V-I DTRR for SatelliteSystems ...... V- 2 V-2 CapitalCosts forSatellites...... V- 2

• • • • • , • • • • • • , • • • • • • • • • • • • • • , • • • • • • • • • • • • • • .... • • • VI-1 Impact of APOs on Spinner SatelliteDesign ...... VI- 12 VI-2 Impact of APOs on 3-Axis SatelliteDesign ...... VI- 12 VI-3 TimelinesforAPOs ...... VI- 15

• • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • , .... • • • • • • VII-1 Incidentsof SatelliteFailure ...... VII - 2 VII-2 Influenceof APOs on Insurance ...... VII- 2 VII-3 Launch Cost Comparison ...... VII-4 VII-4 OTV Launch Time versusThrust ...... VII- 4 VII-5 APO Values for Spinner Satellite...... VII- 6 VII-6 Launch Costs with SB-OTV ...... VII-7 VII-7 APO Values for3-Axis Hybrid Satellite...... VII-8 VII-8 Launch Costs with GB-OTV ...... VII-9 VII-9 Launch Costs with SB-OTV ...... VII-10 VII-10 Repair-at-ShuttleMission Cost ...... VII- 13 VII-11 Shuttle Retrieval, Earth Repair ...... VII- 13 VII-12 LEO Retrieval/Repair at Station ...... VII - 14 VII-13 LEO Retrieval to Space Station, Return to Earth for Repair ...... VII- 14 VII-14 GEO Retrieval, Repair at Space Station Mission Costs ...... VII- 15 VII-15 Itectosat Launch Costs ...... VII-15 VII-16 Capital Cost for 3-axis Large Satellite ...... VII- 16 VII-17 ttectosat Economic Performance ...... VII- 18 VII-18 Summary of APO Economics ...... VII - 19

• • • • • • • • • • • • • • , • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • ° • • • VIII-1 Spinner Scenario Economics ...... VIII- 2 VIII-2 3-Axis Scenario Economics ...... VIII- 3 VIII-3 Assembly/Servicing Economics ...... VIII - 4 VIII-4 Influence of Insurance Rate ...... VIII- 5 VIII-5 Influence of Launch Costs ...... VIII- 5

• • • • • • • • • • • • • • • • ...... • • • • • • • • • • • ...... • • ..... • • A-I.1 Financial Analysis (1985 C-band System) ...... A-2 A-1.2 Income Statement (1985 C-band System) ...... A- 3 A-I.3 Balance Sheet (1985 C-band System) ...... A-4 / A-i.4 Sources and Uses of Funds Statement (1985 C-band System) ...... A- 5 A-l.5a Revenue Assumptions- C-band Transponders (1985 C-band System) ...... A- 6 A-1.6 Capital Expenditure._and General Assumptions (1985 C-band System) ...... A- 6 A-I.7 Operating Expenditure Assumptions (1985 C-band System) ...... A-8 A-I.8 Financing Assumptions (1985 C-band System) ...... A-9 A-I.9 Tax Assumptions (1985 C-band System) ...... A- 10 A-2.1 Financial Analysis (1985 Ku-band System) ...... A- 11 A-2.2 Income Statement (1985 Ku-band System) ...... A- 12

_'k--2 .O D _i,l_t,li£eSheet _.... T__| .... [i_oo Ku-band System ) ...... A- 13 A-2.4 Sources and Uses of Funds Statement (1985 Ku-band System) ...... A-14 A-2.5b Revenue Assumptions- Ku-band Transponders (1985 Ku-band) ...... A- 15 A-2.6 CapitalExpenditure and General Assumptions (1985 Ku-band) ...... A- 16 A-2.7 Operating Expenditure Assumptions (1985 Ku-band System) ...... A- 17

••° Xlll A-2.8 Financing Assumptions (1985 Ku-band System) ...... A- 18 A-2.9 Tax Assumptions (1985 Ku-band System) ...... A - 19 A-3.1 Financial Analysis (1985 Hybrid System) ...... A- 20 A-3.2 Income Statement (1985 Hybrid System) ...... A- 21 A-3.3 Balance Sheet (1985 Hybrid System) ...... A- 22 A-3.4 Sources and Uses of Funds Statement (1985 Hybrid System) ...... A-23 A-3.5a Revenue Assumptions- C-band Transponders (1985 Hybrid) ...... A- 24 A-3.5b Revenue Assumptions - Ku-band Transponders (1985 Hybrid) ...... A- 25 A-3.5c Revenue Assumptions - Other Transponders (1985 Hybrid) ...... A- 26 A-3.6 Capital Expenditure and General Assumptions (1985 Hybrid) ...... A- 27 A-3.7 Operating Expenditure Assumptions (1985 Hybrid System) ...... A- 28 A-3.8 Financing Assumptions (1985 Hybrid Satellite System) ...... A- 29 A-3.9 Tax Assumptions (1985 Hybrid System) ...... A-30

• ° • ° • • • ° . ° ° • • • ° . • ° . • ° , ° • • ° ° ° • ° ° • • • ° , • ° , • • . • ° ° * • ° • B-I.1 Financial Analysis (1995 Spinner System) ...... B - 2 B-1.2 Income Statement (1995 Spinner System) ...... B- 3 B-1.3 Balance Sheet (1995 Spinner System) ...... B- 4 B-1.4 Sources and Uses of Funds Statement (1995 Spinner System) ...... B- 5 B-1.5b Revenue Assumptions - Ku-band Transponders (1995 Spinner System) ...... B- 6 B-1.6 Capital Expenditure and General Assumptions (1995 Spinner System) ...... B- 6 B-1.7 Operating Expenditure Assumptions (1995 Spinner System) ...... B- 8 B-1.8 Financing Assumptions (1995 Spinner System) ...... B- 9 B-1.9 Tax Assumptions (1995 Spinner System) ...... B - 10 B-2.1 Financial Analysis (1995 3-Axis System) ...... B - 11 B-2.2 Income Statement (1995 3-Axis System) ...... B - 12 B-2.3 Balance Sheet (1995 3-Axis System) ...... B - 13 B-2.4 Sources and Uses of Funds Statement (1995 3-Axis System) ...... B - 14 B-2.5b Revenue Assumptions - Ku-band Transponders (1995 3-Axis) ...... B - 15 B-2.6 Capital Expenditure and General Assumptions (1995 3-Axis) ...... B - 16 B-2.7 Operating Expenditure Assumptions (1995 3-Axis System) ...... B- 17 B-2.8 Financing Assumptions (1995 3-Axis System) ...... B - 18 B-2.9 Tax Assumptions (1995 3-Axis System) ...... B - 19 B-3.1 Financial Analysis (1995 Hybrid System) ...... B- 20 B-3.2 Income Statement (1995 Hybrid System) ...... B- 21 B-3.3 Balance Sheet (1995 Hybrid System) ...... B- 22 B-3.4 Sources and Uses of Funds Statement (1995 Hybrid System) ...... B- 23 B-3.5a Revenue Assumptions- C-band Transponders (1995 Hybrid) ...... B- 24 B-3.5b Revenue Assumptions- Ku-band Transponders (1995 Hybrid) ...... B- 25 B-3.5c Revenue Assumptions- Other Transponders (1995 Hybrid) ...... B- 26 B-3.6 Capital Expenditure and General Assumptions (1995 Hybrid) ...... B- 27 B-3.7 Operating Expenditure Assumptions (1995 Hybrid System) ...... B- 28 B-3.8 Financing Assumptions (1995 Hybrid Satellite System) ...... B- 29 B-3.9 Tax Assumptions (1995 Hybrid System) ...... B- 30

xiv Section I

INTRODUCTION

1 Background 3 Approach

Commercial communications satellites form a The technical work is divided into the four tasks high visibility, high benefit use of space and re- described below. Parts of Tasks 3A and 4 were quire a large capital investment. The NASA done first in order to satisfy a November 1985 Space Station may have the potential to provide deadline for inputs to a Space Station require- significant economic benefits to the commer- ments review. cial communications satellite operators, probably with considerable change in satellite design and operation. 3.1 Task 1. Develop Communica- The diverse objectives and lack of standardiza- tions Satellite Financial Model tion in the commercial sector will require NASA coordination and direction to maximize Space 3.1.1 Task 1A: Develop Basic Financial Station benefits. NASA has taken the lead with Model this study which seeks to quantify the benefits of new space-based activities and assess the im- The objective of Task 1A is to develop a finan- pacts on the satellite design and the Space Sta- cial model that describes quantitatively the eco- tion. nomics of the space segment of U. S. domestic Fixed Satellite Service (FSS) communication satellite systems. The model describes the eco- 2 Objectives of Study nomic status of the system throughout the lifetime of the satellite. The model is applicable There are three objectives of this study: over the range of satellites expected to be imple- 1. Develop a quantitative methodology to as- mented over the next 10 years. Output of this task is as follows: sess the viability of a broad range of new space-based activities, procedures, and operations (APOs) when utilized in commer- A financial model which describes the eco- cial communications satellite system opera- nomics of the space segment of FSS sys- tions; tems and a definition of the economic factors comprising the model. 2. Apply the developed methodology to select which of these APOs can be competitively The values of the economic factors of three provided by the Space Station and its asso- representative operating systems calculated ciated operating systems; and on a year-by-year basis over the systems' lifetimes. 3. Determine the economic and functional requirements imposed on the Space Station through the provision of these selected • The model output values of the economic APOs. performance of the three systems.

I-1 3.1.2 Task 1B: Assess Impact of System • The model output values of the economic Characteristics on Financial Model performance for each of the postulated sys- Output tems.

The objective of Task 1B is to identify those 3.3 Task 3. Assess Economics of New communication satellite system technical and Space-Based Activities for 1995 functional characteristics that significantly affect the economic factors (used in the model devel- 3.3.1 Task 3A: Postulate New Space- oped in Task 1A) and the model output values, Based APOs and to perform a sensitivity analysis. The objective of Task 3A is to postulate and The output of this task is as follows: define new or enhanced space-based APOs and An identification of the system character- associated satellite system designs that have istics affecting the economic factors in the the potential to achieve future communications financial model. satellite operations in geostationary orbit with improved economic performance. The availabil- The parametric relationships between the ity of the Space Station and its associated sys- system characteristics and the economic fac- tems as projected by NASA shall be assumed. tors. Output of this task is as follows:

The resultsof a sensitivityanalysisof the • Postulations and definitions of new space- impact of system characteristicson the sys- based APOs to be utilized by future com- tem economic performance. munications satellite systems. • Definitions of changes in satellite system de- 3.2 Task 2. Determine Eco- signs to accommodate the new APOs. nomic Performance of Business- • By 22 November 1985 a preliminary assess- as-Usual Scenario for 1995 ment of the most significant APOs. The objectives of Task 2 are to assess and describe the expected state-of-the-art status of 3.3.2 Task 3B: Evaluate Economics of communications satellite systems and operations APOs for U. S. domestic FSS systems beginning service The objective of Task 3B is to establish economic in 1995. The results of the assessment are uti- target values or target costs for the APOs postu- lized to postulate and describe three representa- lated in Task 3A to provide an incentive for their tive satellite systems. The output of Task 2 is as implementation by the industry. The output of follows: / this task is as follows: An assessment and forecast of the evolution- • Estimates of any increases in system eco- ary improvements in spacecraft and commu- nomic performance requl_red when the nications technologies by 1995. APOs defined in Task 3A are utilized during the system's life cycle. A postulation and description of three systems incorporating the expected improve- • Economic target values for each APO de- ments. fined in Task 3A.

The modified values of the economic fac- 3.4 Task 4. Develop Space Station tors identified in Task 1 resulting from im- Scenarios and Requirements plementing the postulated systems. These modified values shall be on a year-by-year The objective of Task 4 is to describe at least basis throughout the systems' lifetimes. two communications satellite system operating

I-2 scenarios implementing different combinations of the APOs defined in Task 3 through utilization of a low earth orbit (LEO) Space Station and its supporting equipment/systems. The output of this task is as follows:

Descriptions of two communication satellite system operating scenarios utilizing the Space Station.

Descriptionsof the functionaland technical requirementsimposed on the Space Station through implementation of each scenario. Section Task Content of Section By 22 November 1985, descriptionsof two II 1A FinancialModel Description communication satellitesystem operating and Validation scenariosutilizingthe Space Stationand in- III 1B SensitivityAnalysisof Model corporating the preliminaryset of AP0s, and the resultingfunctionaland technical IV 2 Technology Assessment and requirementsimposed on the Space Station. Definition of 1995 systems V 2 Economic Performance of 4 Organization of Report 1995 systems VI 3A Postulation of new The correspondencebetween the Sectionsof this space-based APOs report and the program Tasks is shown in Ta- VII 3B Economics of APOs ble I-1.Appendices A and B containthe Finan- cialModel program output sheetsforthe differ- VIII 4 Space Station Scenarios ent satellitesystems. IX 4" Space StationRequirements

X Recommendations

A 1A Model Outputs for 1985 Satellite systems B 2 Model Outputs for 1995 Satellite systems

Table I-1: Organization of Report

I-3 Section II

FINANCIAL MODEL

1 Introduction 1. System characteristics

This section describes the Task 1A development 2. Capital expenditures and validation of the communications satellite fi- 3. Revenues nancial model (the Model) that describes quantitatively the economics of the space segment 4. Operating expenditures of U. S. domestic Fixed Satellite Service (FSS) communication satellite systems. (Ground ter- 5. Financing activities minals and terrestrialsystem costsaxe excluded from considerationexcept forsatellitetelemetry, 6. Taxes tracking,and controlsystems.) The Model describesthe economic statusof 2.1 System Characteristics the system throughout the lifetimeof the satel- litebeginning with its design and continuing The first step in determining system economic through itsconstruction,launch, and commer- performance is to describe the satellite system cialoperations. The Model can be applied to characteristics which drive both revenues and a range of satellitesizes,communications pay- costs. To validate the Model, three 1985 sys- loads,and lifetimesexpected to be implemented tems were developed. Their characteristics are in the 1985 to 1995 time frame. summarized in Table II-1. Subsection II-2 givesthe model assumptions These three systems are chosen because their which include both satellitetechnicalperfor- characteristics are most representative of the mance and financialfactors. current FSS environment. Although specific The three methods of financialperformance satellite characteristics are used, the economic measurement provided by the Model are dis- factors driven by these characteristics are ad- cussed in SubsectionII-3. justed to reflect industry norms for the partic- SubsectionII-4givesthe Model User Manual ular satellite system. and program description. The three systems are described in the follow- The Model is validatedin Subsection 5 by ing subsections. an analysis of three representativeoperating systems for 1985 initialoperationalcapability 2.1.1 C-Band Satellite (ioc). Table II-2 lists the characteristics of the C-band satellite, which is a small spin-stabilized satel- 2 Model Assumptions lite with a payload of 24 transponders of 5.5 W power and 36 MHz bandwidth. The HS-376 de- Model assumptions consist of the various input sign is widely used for small C-band satellites. data necessary to operate the Model. They can Its relatively modest primary power requirement be grouped in six categories which are discussed of 800 W makes the payload well suited to the in turn:

II - 1 C-Band Ku-Band Hybrid Satellite: Telstar 303 Satcom K2 Spacenet 1 Operatnr: AT_-T RCA Americem Prime contractor: Hughes RCA Astro-Elec. RCA Astro-Elec. EIRP (dBW): 31-34 45-48 33-38 C, 40-44 Ku Design life (yr): 10 10 10 BOL mass (kg): 659 1015 710 Payload mass (kg): 144 175 156 - Antenna (kg): 45 24 46 - Transponder (kg): 99 151 110 EOL power (W): 8OO 2440 1150 Stabilization: Spin Three-axis Three-axis Frequency: C-band Ku-band C and Ku-band Frequency re-uses - C-band: 2 - Ku-band: 2 Number of transponders - C-band: 24 18 - Ku-band: 16 6 Transponder bandwidth - C-band (MHz): 36 36 (12 each) 72 (6 each) - Ku-band (MHz): 54 72 (6 each) Transponder power - C-band (W): 5.5 8.5 (12 each) 16 (6 each) - Ku-band (W): 5O 16 Antenna coverages - C-band: - Ku-band: 2 Satellite EIRP - C-band (dBW): 31-34 33-38 - Ku-band (dBW): 45-48 40-44 Launch vehicle(s): Delta 3920 Ariane Ariane, Delta STS/PAM D STS/PAM D II STS/PAM D

Table II-l: Co_m__p_.risonof 1985 SatelAite Characteristics

II - 2 Na_ne: Telstar 303 Manufacturer & model: Hughes HS-376 EIRP: 31-34 dBW Lifetime: 10 yr (1985 launch) On-board switching: Transponders switchable from TWTA to SSPA Launch vehicle: Delta 3920 or STS/PAM D Ground terminals: 13 m, uncooled paramp LNA, 3 kW HPA 30 m, cooled paramp LNA, 3 kW HPA Frequency band and bandwidth: C-band, 500 MHz - receive: 5.925-6.425 GHz - transmit: 3.7-4.2 GHz - number of times reuse: 2 Antenna - type: Offset paraboloid, dual gridded - number: 1 - size: 1.85 m -- mass: 45 kg - coverage (2 beams): Conus or Conus and/or one spot - polarization: H and V, linear Transponders SSPA's and TWTA's - number: 24 - power: 5.5 W - bandwidth: 36 MHz - redundancy: 5 for 4, (18 SSPA's and 12 TWTA's, switchable) - receiver redundancy: 4for2 -- mass: 99 kg - dc power: 500 W Spacecraft - type: Spin stabilized - size(stowed): dia = 2.17 m, length = 2.85 m - mass, BOL: 660 kg - power, EOL at summer solstice: 800 W - primary power: Solarcells - batteries: 2 x 32 cellNiCad, 20 Ah each - attitude and station keeping: IIydrazinethrusters - attitude pointing accuracy: +0.05 ° - apogee motor: Solid propellant

Table II-2: C-band Spinner Satellite Characteristics (1985)

II - 3 The antenna can provide CONUS (continen- 2.1.3 Hybrid Satellite talUnited States)coveragewith all24 channels, The hybrid satellite supplies transponders at two or alternately,6 channels can be switched to frequency bands and makes better use of an or- Alaska, 6 to Puerto Rico,and 6 to Hawaii. bital position. Although there are few exist- The Telstarseriesisused by AT&T for video ing hybrid satellites, this concept will become and voice.Itcarriesboth 5.5W solidstatepower widespread in the future. amplifiers(SSPA) and 5.5W travelingwave tube Table II-4 lists the characteristics of the Hy- amplifiers(TWTA) which can be switched be- brid satellite. This is a three axis satellite with tween transpondersby ground command in or- 18 C-band and 6 Ku-band transponders. Twelve der to best suitthecommunications trafficbeing of the C-band transponders are 8.5 W SSPAs relayedby the particulartransponder. with 36 MHz bandwidth, while the remaining 6 The more linearSSPAs significantlyimprove transponders are 16 W TWTAs with 72 MHz channel capacity.For instance,the conventional bandwidth. The 6 Ku-band transponders have FDM/FM modulation technique allows 1,800 16 W TWTAs with 72 MHz bandwidth. full-voicecircuitsto be carriedby one transpon- The coverage at C-band is for CONUS, the der usinga TWTA. Using companded SSB mod- Caribbean, Alaska, and Hawaii. Only CONUS ulation,3,900full-voicecircuitscan be carriedby coverage is supplied at Ku-band. one transponderusinga SSPA of similarpower. The spinningcylinderdesign minimizes sta- 2.2 Capital Expenditures tionkeeping fuelrequirements and allowsspace A capital expenditure is normally an outflow of for an integralsolidpropellant apogee stage. cash which is generated from debt or equity and However, space forsolarcellsis limited to the is used to obtain an asset with a useful life usu- surface area of the cylinder,only one side of ally exceeding one year. For a satellite operator, which faces the sun at one time. Small spin- capital expenditures are comprised of the follow- ners axe lessexpensivethan three-a_xissatellite ing costs: designs,but above a certainsizethe three-axis designischeaper. • Satellite

• Perigeestage(ifneeded)

2.1.2 Ku-Band Satellite • STS launch

Table II-3liststhecharacteristicsof the Ku-band • Launch operations satellite,which isthree-a_s stabilizedwith 16 • Mission operations transponders of 50 W power and 54 MHz bandwidth. The 2440 W power requirement due to • Launch insurance the high power transpondersfavorsthethreeaxis The cash outflowsforcapitalexpendituresare designof the Satcom K2 satellite. predeterminedby milestonebillingschedulesne- The coverage of the 16 channels can be gotiatedbetween the operatorand the manufac- switchedon-orbitby ground command between turerforthe satelliteand relatedservices.NASA CONUS coverage and half CONUS coverage. has establishedpayment schedulesfor the STS In addition,the 8 horizontalpolarizationchan- launch. nels can be reconfiguredas a group to include Capital _t_ _. ,_preciated over a useful Caribbean coverage. life consistent with IRS regulations for finan- The use of high power and east and west re- cial accounting purposes. Depreciation is impor- gionalbeams allowstransponders to be used to tant because it determines the amount of taxes distributevideo servicesdirectlyto 1 m receive- paid. It is assumed that five year ACRS (accel- only terminals. erated cost recovery system) depreciation is used

II - 4 Name: Satcom K2 Manufacturer & model: RCA Americom, K2 EIRP: Conus=45, half Conus=48 dBW Lifetime: 10 yr (1985 launch) On-board switching& processing: Switchable among coverage regions Launch vehicle: Ariane IV, STS/PAM D II Ground terminals: As small as 1 m for receive-only Frequency band and bandwidth: Ku-band, 500 MHz - receive: 14.0-14.5 GHz - transmit: 11.7-12.2 GHz - number of times reuse: 2 Antenna - type: Offset paraboloid, dual gridded - number: 1 - size: 1.52 m -- mass: 24 kg - coverage (2 beams): CONUS or E and W CONUS - polarization: HandV Transponders TWTA's - number: 16 - power: 50 W - bandwidth: 54 MHz - TWTA redundency: 11 for 8 - receiverredundency: 4for2 -- mass: 151 kg -dc power: 2070 W Spacecraft - type: 3-axis stabilized - size(bus): 1.57 x 2.18 x 1.77 m - mass, BOL: 1018 kg - power, EOL at summer solstice:2440 W - primary power: Solarcells - batteries: 3 x 22 cellNiH, 150 Ah - thermal control: Heat pipes - attitudeand stationkeeping: Hydrazine thrusters(electrothermalN/S) - attitudepointingaccuracy: 4"0.07 ° - apogee motor: Solid propellant

Table II-3: Ku-Band Satellite Characteristics, (1985)

II - 5 NaJne: Spacenet 1 Manufacturer & model: RCA EIRP: 33-35 dBW (narrow) & 33-38 dBW (wide) C-band 40 - 44 dBW, Ku-band Lifetime: 10 yr (1985 launch) On-board switching: C- and Ku-bands interconnected Launch vehicle: Thor delta3920 or Ariane Ground terminal - C-band: 13 m, 8U K F_,'I-' LINA; 18 m, 33 K uncooied paramp LNA Ground terminal - Ku-band: 9.2 m, 11.3 m, and 13 m; 190 K FET LNA Frequency band and bandwidth: C-band 500 MHz, Ku-band 500 MHz - receive: 5.925-6.425 GHz and 14.0-14.5 GHz - transmit: 3.700-4.200 GHz and 11.7-12.2 GHz - number of times reuse: 2 (C-band) plus 1 (Ku-band) Antenna - type: Offset paraboloid, dual-gridded - number: 2 - size: 1.52 m -- mass: 46 kg - coverage, 2 beams C-band: CONUS, Caribbean, AK, and HI - coverage, 1 beam Ku-band: CONUS - polarization: H (narrow) and V (wideband) C-band, H for Ku-band Transponders SSPA's 8.5 W C-band; TWTA's 16 W C _: Ku-bands - number at C-band: 12 each SSPA (36 MHz), 6 each TWTA (72 MHz) - power at C-band: 8.5 W SSPA (36 MHz), 16 W TWTA (72 MHz) - bandwidth at C-band: 36 MHz (12 channels), 72 MHz (6 channels) - trans, redundancy at C-band: 7 for 6, SSPA's; 8 for 6 TWTA's - receiver redundancy, C-band: 4 for2 - number at Ku-band: 6 - power at Ku-band: 16 W TWTA - bandwidth at Ku-band: 72 MHz - TWTA redundancy, Ku-band: 8for6 - receiver redundancy, Ku-band: 2for 1 -- mass: 110 kg - dc power: 1,000 W Spacecraft - type: 3-axis stabilized - size(stowed): 1.32 m x 1.63 m x 1.5 m - mass, BOL: 670 kg - power, EOL at summer solstice: 1,150 W - primary power: Solarcells - batteries: 2 x 22 cellNiH, 40 Ah each - attitude and station keeping: Hydrazine thrusters(electrothermalN/S) - attitude pointing accuracy: ±0.05 ° - apogee motor: Solid propellant

Table II-4: Hybrid Satellite Characteristics (1985)

II- 6 to achieve the maximum tax shield in the early years of the project (Subsection 2.6.2). Satellite Cost For purposes of the Model, all tax benefits flow Type $ M through to the parent company to be used when C-band 38.2 generated. It is assumed that the owner has Ku-band 48.2 other profitable operations which can use these Hybrid 43.9 tax benefits. The major categories of capital expenditures are discussed in the following subsections. Table II-5:1985 Satellite Costs

• Thermal 2.2.1 Satellite Costs • TT&C Satellite costs are developed from the Ford Aerospace database using the Price H cost • Payload - Antenna model. The PRICE (Parametric Review of Informa- • Payload - Transponder tion for Costing and Evaluation) H (Hardware) Price H outputs another category, integration Model is a computerized method for deriving and test, which is based on the input data. cost estimates of electronic and mechanical hard- The second step involves an interaction be- ware assemblies and systems. Price H contains tween the user and the Price H model to cali- several thousand parametric equations that are brate the Price H output. This process typically not accessible to the user. takes one month for each new satellite database. The fundamental inputs for Price H include The Ford Aerospace satellite cost database is the following: used to validate the satellite costs. Adding a G&A expense of 12% and manufac- • Quantities of equipment turer's fee of 12% to the Price H output results • Schedules in the satellite costs shown in Table II-5. The satellite cost is in 1985 dollars and is an input to • Hardware size, weight, density the Model. The cost excludes STS launch costs, perigee stage costs, launch operations, mission • Amount of new design operations, and launch insurance. • Complexity 2.2.2 Perigee Stage Costs • Maturity of technology For 1985 systems, separate perigee stages are In order to use Price H, the first step is to necessary and are assumed to be provided by create and store the hardware parametric data. the PAM D for the C-band and hybrid satellites Separate data files are created for the seven satel- and by the PAM-D2 for the Ku-band satellites. lite subsystems. (Payload is divided into an- Prices and payment schedules for these stages tenna and transponders to facilitate APO evalu- are determined from McDonnell Douglas pub- ation.) The database for each satellite has eight lished prices. files as follows: For 1995 business-as-usual systems, some of the satellites incorporate hypothesized in- • Attitude control tegrated upper stages which combine perigee, • Power apogee, and station keeping propulsion systems. For display purposes and consistency, the cost of • Propulsion the perigee portion of the stage is separately displayed on the Capital Expenditure Assumption • Structure

II - 7 2.2.3 STS Launch Costs when testing is completed and control is fully assumed by the operator. The STS launch cost is calculated based on the Every insurance policy is unique. Each un- load factor and anticipated launch date. The derwriter has his own methodology for comput- launch costs in the Model are priced using current NASA rates and escalation factors. The ing the formula to determine the value of the sum insured. There are four basic components launch cost is assumed to be contracted directly to consider in determining this value: by the operator with NASA, and thereby avoids an allocation of the manufacturer's G_A expense and fee.

• Cost of a replacement launch 2.2.4 Launch and Mission Operations • Loss of revenues Launch support and mission operations costs are derived from Ford Aerospace experience. These costs include installation of the satellite in the • Possible cost of increase in debt financing launch vehicle and telemetry, tracking, and con- The satellite replacement cost is the manufac- trol during launch and checkout and during station keeping operations. turing cost plus satellite asset value (including perigee stage). The replacement launch cost consists of the charges for launch and mission opera- 2.2.5 Launch Insurance tions, the actual STS launch cost, and launch in- Launch insurance costs are based on rates ap- surance. Some operators also insure against the propriate to the satellite launch date. For 1985 loss of contracted revenues for the period com- satellites, insurance was contracted for two or mencing at the loss or malfunction and ending three years prior to the anticipated launch and when the replacement satellite has been successfully placed in orbit. However, such insurance is is assumed at 14% of the insured value (1982 quoted rates for 1985 launches ranged from 12% becoming increasingly costly. to15%). During this time period, it is also necessary to Current ratesrange from 26% to 31% with hedge against future interest rate increases. It is minimal advance commitment for coverage by possible to insure against such an occurrence in the insurers.These high ratesreflectrecentun- an inflationary period. derwriterlossesin excess of $600 million. In- surance rateswillnot declinesignificantlyuntil 2.2.6 Payment Schedules there are numerous consecutivesuccessfulmis- sionsand the insuranceindustry recoversthese Payment schedules rA_•_, ,,.__._ c_pl_,u_ :,_, expenul_-"ures losses. are shown in Table II-5. Payments for other than Based on conversationswith major insurance STS launch and launch insurance are largely a underwriters and brokers,it is assumed that matter of negotiation between the operator and in 1995 without the Space Station,insurance the manufacturer and vary widely among satel- is priced at 20% of insured value. This price lite systems. assumes that the industry has had successful This study assumes that 90% of the total con- launches and placements of satellitesin service tract value is paid during satellite construction sufficient to allow such a drop_ Otherwise, indus- and the remaining 10% ispaid at completion of try participants agree that current costs of 30% checkout with a warranty payback. Under war- will prevail in 1995. ranty payback, the manufacturer partiallyguar- The period of coverage for launch insurance anteessatelliteperformance by agreeingto reim- is from "intentional ignition" through checkout. burse a portion of the 10% if thereis a failure The last day of the launch insurance period is priorto the end of satellitedesignlife.

II - 8 to allow the comparison of transponders with dif- Yearly Payment, % ferent characteristics, but it is the market that Capital Expense 1 2 3 4 ultimately determines price. Satellite 29 40 21 10 The relationships derived below are based on STS Launch 18 31 51 - market price data. As such, they may change Perigee Stage 29 40 21 10 with time if there are significant technological or Launch Support 29 40 21 10 regulatory changes. For instance, closer satel- Mission Operations 29 40 21 10 lite spacings may lead to an interference-limited Launch Insurance 17 - 70 13 rather than the current noise-limited communications environment.

Table II-6: Payment Schedule, %/yr Bandwidth Transponder price is directly related to band- 2.3 Revenues width ratio according to the formula: Project revenues are affected by five related fac- Price o¢ k (Bandwidth ratio) tors: where k = 0.90. This is because the commu- 1. Characteristics of transponder nications capacity of a channel is directly pro- 2. Degradation of satellite portional to its bandwidth. The factor k is less than unity to account for the technical difficul- 3. Market factors ties of passing multiple signals through a nonlinear channel. 4. Price of transponder (Note that the proper price comparison is to 5. Utilizationof satellite compare, for example, two 36 MHz transponders of 8 W with one 72 MHz transponder of The approach used forrevenue determinationis 16 W.) The basis for this relationship is dis- to calculatebase model revenues forour repre- cussed in more detail in Subsection III-3.2. sentativeC-band system and to index the price of other systems based on theirdifferencesfrom the base case. Frequency Band The following factors influence the relative value of Ku-band versus C-band transponders: 2.3.1 Characteristics of Transponder • Atmospheric attenuation requires a 2 to The following transponder characteristics im- 8 dB increase in Ku-band link margin rel- pact transponder price: ative tc_ C-band margins, depending on ge- • Bandwidth ographical location and required link availability. This translates to relative factors of • Frequency band of operation from 0.8 to 0.5 for Ku-band. (This factor is partially offset by the higher allowed PFD • Power at Ku-band - see Subsection III-2.4.3.) • Transponder type • Terrestrial microwave interference may pre- These items are discussed briefly here, with more vent use of C-band for ground transmitting detail presented in Section III. in certain locations. This increases the relative value of Ku-band transponders.

The reader is cautioned against concluding • The technology at Ku-band is less mature that transponder prices can be reduced to a for- than C-band. This decreases the relative mula. General relationships are derived in order v_T,n_ _f T_,_,_1._,_,,.1

II - 9 • C-band tendsto use SSPAs versusTWTAs at Ku-band. This increasesthe valueof C- band versusKu-band. (See discussionbelow -_ C-band SatelliteKu-bandSystem Hybrid under "Transponder Type".) 1 1.00 1.00 1.00 2 1.00 1.00 1.00 Itisevident thatthe frequencyband factoris 3 1.00 1.00 1.00 based on more than simple communications ca- 4 1.00 1.00 1.00 pacity,and thus thereisconsiderableuncertainty 1.00 .94 1.00 in itsactualvalue.Itsvaluehas been chosen to t_t_ U 1.00 .ISU .UO be consistentwith the transponder pricingdata 7 .95 .87 .90 availableto us atthistime. Relativetransponder 8 .90 .85 .85 frequency factorsaxe 1.00 for C-band and 0.80 9 .85 .78 .80 for Ku-band. More detailisgiven in Subsection 10 .80 .75 .75 III-3.3. 11 .80 .70 .70 The reader is cautioned against jumping to 12 .00 .00 .00 the conclusionthat C-band transponders are worth more than Ku-band transponders. The greater power allowed at Ku-band (typically Table II-7: Satellite Degradation Curves 50 W versus8.5W at C-band) and largerband- widths used at Ku-band (typically54 MHz ver- the modulation method as discussed in Subsec- sus 36 Mgz at C-band) can make the priceof tion III-4.11, and ranges from 4% for FDM single a Ku-band transpondertwice that of a C-band access to 50% for SCPC modulation. transponder. For 1985 systems the difference is small and this factor is not significant. For 1995 sys- Power tems, the case will probably be that all C-band Transponder price varies with transponder transponders use SSPAs and thus can be com- power accordingto: pared directly. Ku-band transponders will still be TWTAs and the SSPA versus TWTA factor ( Powerratio _x Price oc \ ba_th-_d-tatiO ] can be folded into the frequency factor. There- fore, this factor is not explicitly included in the where x = 0.33 and the power ratiorefersto Model. transponderpower dividedby thebaselinetrans- ponder power (same forbandwidth ratio).For 2.3.2 Degradation of Satellite example, a transponder with 5.5 W power and 36 MHz bandwidth compared to the baseline Degradation curves are deve, 1ope,_,-1 __,,_,,,,.K," _1,. i,,_,u_,e"_ _1 ,-1 transponder with 8.5 W power and 36 MHz factors for reliability, redundancy, lifetime, and bandwidth has a relativepriceof (5.5/8.5)"33= the technical characteristics of the satellite. Ta- 0.87. ble II-7 gives the degradation curves used for the Again, thisfactorhas been chosen to be con- three 1985 satellite systems described in Subsec- sistentwith availabletransponder pricingdata, tion 2.1. These curves relate to the generic satel- and may change as technology and regulations lite types, not necessarily to the specific satellite vary.More detailisgivenin SubsectionIII-3.4. types named in Tables II-1 to II-5. The curves are for sate!litos with 10 year lifetimes. Transponder Type These curves are composites of all factors such The greaterlinearityof the SSPA versus the as infant mortality, random failures, wear-out, TWTA improves the communications capacity and the exhaustion of expendables such as proof a channel forthe same bandwidth and power. pellant. The curves are based on past experience However, the amount of increasedepends on for the satellite type. For an individual satellite,

II- 10 variationsin degradationmay be expected. Class of Price, 1985 2.3.3 Market Factors Transponder ($M/yr) Protected 1.90 Transponder lease price is determined by (1) the Unprotected 1.40 cost of providing the transponder; and (2) mar- Preemptible .90 ket considerations which include:

• Added features Table II-8: C-Band Transponder Prices

• Competition These prices represent the annualized lease • Financial status of company cost for protected, unprotected, and preemptible transponders. The figures have been adjusted to • Method of payment include both inflation and an increase in operat- Q Orbital position of satellite ing costs over the life of the satellite, as well as customer discounts for longer leases and use of • Special package deals multiple transponders. Therefore, for 1985 satellites, it is not necessary to artificially inflate the • Supply and demand for transponders base transponder prices at any point during the • Other new or unknown factors life of the satellite.

The market factors are not amenable to quan- 2.3.5 Utilization of Transponder titative analysis but are typically the single most significant factor in determining price. It is For this model, utilization is defined to include beyond the scope of this program to analyze all transponders that have been sold or leased. these factors but they are included for complete- The actual usage of the transponder does not ness. They determine the market price of a basic matter to the Model since it doesn't influence transponder, which is an input to the Model. revenues. There is also a "market factor" input to the In today's market, owners and operators are model in order to allow for changes in the mar- expecting to realize 90% utilization. There is ket. The market factors are 1.00 for C-band and a relative abundance of C-band transponders, 1.13 for Ku-band, reflecting a preference for Ku- however, and their utilization may be consider- band transponders relative to C-band transpon- ably less. Ku-band transponders are in short ders. supply, and their utilization may be higher.

2.3.4 Price of Transponder 2.4 Operating Expenditures The market price of a basic C-band transpon- An operating expenditure is a cost incurred in der is based on a review of current in-house the normal operations of the firm to sustain and databases and verified through interviews with support day to day activities. There are four cat- manufacturers, operators and owners, and other egories of operating expenditures in this model: experts in the communications satellite field. The three types of transponders and their re- • Li_ insurance spective market prices are listed in Table II-8. These prices are for 1985 C-band transponders • Other expenditures which reflect the total market and do not represent any particular transponder. The prices are • Rate of inflation on a per annum basis for an average three to six year lease. • Telemetry, tracking, and control

II - 11 2.4.1 Life Insurance variable component is subject to inflation (assumed to be 4% per annum). This results in an Life insurance is calculated as a percentage of effective inflation rate of approximately 1% of the net present value (NPV) of the future rev- the gross amount. enue stream. Rates are based on conversations with insurance brokers, dealers and underwrit- 2.5 Financing Assumptions ers. Each policy is unique, and is not necessarily calculated using the same methodology or for- Financing decisions should be independent from mula as the others. investment analysis decisions. For this reason, Policies are based on the "sum insured". This a set of representative financing assumptions is value is calculated on the satellite's value less made and held constant throughout the study. depreciation (asset value, not economic value) Factors that influence financing decisions for and lost revenue (business interruption). a given project include debt ratios (debt-to-total The insurance premium currently ranges from assets or debt-to-equity), alternative methods of 3% to 6% of the amount insured. For purposes financing, and interest rates. of this model, the annual insurance premium is assumed to be 4% of the present value of future 2.5.1 Balance Sheet Financing gross revenues. General Balance Sheet financing is assumed. This means that the sources of funds come from 2.4.2 Other Expenditures internally generated equity and general company The other operating expenses are calculated us- debt. A debt-to-total asset ratio of 45% is rep- ing a base amount in the launch year and an resentative for a firm in the aerospace industry inflation premium multiplier for future years. and is incorporated in the Model (i.e. 45% of the Sales and marketing and G_A expenses are capital expenditures is funded with debt and the further divided into pre-operational and post- remaining 55% with equity). operational periods. This split permits us to em- ploy different bases for calculations. The sales 2.5.2 Period of Financing and marketing charge is calculated on a per transponder basis, but is reported on the Income Money is borrowed as needed during the con- Statement for the entire satellite. struction period. During this time, the principal accrues and only interest is paid. 2.4.3 tLate of Inflation Once the satellite becomes operational, level monthly payments of principal and interest are A 4% annual rate of inflation is forecast for the made for five years. Investment bankers agree life of the project. This is consistent with the that a five year loan is acceptable and desirable, current DRI (Data Resources, Inc.) and Whar- relative to the lifetime of most satellites. If the ton Econometrics inflation forecasts. loan period is longer, the def£ult risk increases.

2.4.4 TTg_C 2.5.3 Interest tLates

The telemetry, tracking, and control (TT&C) The interest rate charged during the construc- system is capable of monitoring two or three tion period and during operation is assumed to satellites. It requires a $20 million capital in- be variable and equivalent to the prime rate plus

...... , ,_,,,,.v_,.,.., _, so etc., Vl.S.LLL_O* .L XL_ _I_L_LZ _Z_.JLZCA_L v *vv _.SVZXL_O ,L _X_,_ a and an annual charge of $0.5 million for oper- risk factor based on corporate creditworthiness ations (salaries, maintenance, software replace- and a charge to guarantee a line of credit. Ta- ment and improvement, etc.) ble III-9 gives the prime rates during construc- The TT&C cost includes both fixed (capital) tion of 1985 satellites. These rates are used as and variable (operations) components; only the a basis for calculating interest expenses in the

II - 12 2.6.2 Depreciation

Year J AverageR_te, Prime% All capital expenditures also qualify for depredation. There are three depreciation methods 1983 I 14.91 available (but only ACRS is implemented in the 1984 I 10.83 Model): 1985 I 11.91 1986[ 9.00 • Straight line

* Accelerated method Table II-9: Prime Interest Rates • Straight line/accelerated combination

Model. (The tabulated 1986 rate is an average In 1981, as part of the Economic Recovery Tax for the first quarter.) The prime rate is assumed Act, Congress passed the Accelerated Cost Re- to be 9% for 1995 satellites. covery System (ACRS) method for depreciation. The Act permitted accelerated recovery for most 2.6 Tax Assumptions tangible property (capital purchases) placed in service after 31 December 1980. For tax pur- All tax computations are based on the 1985 In- poses under ACRS, communication satellites are ternal Revenue Code (the Code), including all considered to be five year property. amendments and deletions. The Code and tax The ACRS method is used because it permits laws are subject to change annually as a result the shortest depreciation period and thereby of Congressional legislation. Further, it is as- generates the greatest benefits on a present value sumed that the owner/operator is a U.S. corpo- basis. ration or a subsidiary of a U.S. corporation and is subject to the regulations and procedures of 2.6.3 Marginal Tax Rate the Code. We are evaluating this project inde- The project is taxed at a marginal tax rate of pendently and not in conjunction with other new 49.78%. This is based on a state tax rate of 7% or ongoing projects. All tax benefits are utilized and a federal tax rate of 46% (any corporation at the marginal tax rates when generated. These with earnings greater than $100,000 must be in benefits consist of investment tax credits and depreciation. the 46% federal tax bracket). The aggregate rate of 49.78% reflects the deduction of state tax payments from federal taxable income. 2.6.1 Investment Tax Credit

All capital expenditures qualify for an invest- 3 Financial Analysis ment tax credit (ITC). The ITC is an after- The Model provides the user with three measure- tax credit which lowers the tax liability by the ments of economic performance: amount of the credit and is taken in the tax period when the satellite is placed in service. • Net present value of project The corporation may elect one of two ITC choices; 8% or 10%. The 10% ITC election re- • Internal rate of return on equity duces the basis for depreciation by one half of • Dual terminal rate of return on equity the ITC taken. The 8% ITC election gives an 8% credit after tax without reducing the depre- ciable base. 3.1 Net Present Value

Based on the discount rates used in this model, Net present value (NPV) is determined by using an 8% ITC is chosen since the corporation real- the sum of all current and future cash flows of the izes more benefits. JLJ.LV_ObJ.J.L_ILIJ U_UU_IL_U. I.t<_L,J_ bU b1111_ /,dC_IU. ilL_

II - 13 key to this method of analysis is the selection of - The IRa formula discounts both positive the appropriate discount rate. In this process, and negative cash flows using the computed careful consideration should be given to the rel- IRR. Again, this may be unrealistic. ative risk associated with the investment. Com- monly used discount rates include the investor's - If negative cash flows occur at various times marginal cost of capital, alternative investment over the life of investment, the IRt{ formula rate, and investment hurdle rate. generates multiple solutions. Advantages and disadvantages of using NPV

v_ zAa v ,,,,_u,,uJ.f,,,,,JUL_oQ, AI_,,,_ A'-J,AJLVVV0 q=e• u J.=,e L,aGLA .L_JL JLJLAA,ILAIOI.JLJLIL,(3i.IjI_ UJL ,LI_ _ U.JL JLJL (as indicated by the plus and minus sign respec- The dual terminal rate of return (DTRtL) is sim- tively): ilar in concept to the ItLtL computation. The + It is relatively easy to compute; DTtLtL calculation first discounts all negative cash flows back to time zero using a rate which + It effectively considers the time value of approximates the investor's marginal cost of cap- money; ital. + It inherently utilizes a good reinvestment as- Next, the future value of the positive cash sumption; flows is computed assuming a reinvestment rate which approximates the one an investor could - It gives only the current dollar magnitude earn on other alternatives. This future value is of the transaction without considering the at the end of the sateUite's life. dollar value of alternative investments. Finally, the future value of the positive cash flows is discounted back to time zero using a rate 3.2 Internal Rate of Return that results in a value equal to the discounted value of the negative cash flows (i.e. NPV = 0). The internal rate of return (IRK) for an invest- The discount rate used in this final step is the ment is that rate which, when used to discount investment's DTRR. both positive and negative cash flows back to Advantages and disadvantages of using DTRR time zero, results in a present value equal to zero. to evaluate alternative investments are as fol- Once generated, an investment's IRR is an effec- lows: tive means of comparing alternative investments (regardless of the magnitude) and can also be + It compensates for cash investments that are used to evaluate individual transactions by com- spread over several periods; paring the II{R to the investor's marginal cost of capital or investment rate. + It allows considerable flexibility in making Advantages and disadvantages of using II_R to reinvestment assumptions; evaluate alternative investments are as follows: + It eliminates the ambiguity of multiple so- + It is widely used and understood. lutions; + It effectively considers the dollar magni- + It effectively compensates for variations in tude of alternative investments and the time the relative size of alternative investments; value of money. + It accurately considers the time value of - The IRK formula assumes all positive cash money; flows are reinvested at ,ho TPP rate, -.*_" _ may not be realistic. For instance, if a par- - The method has increased complexity and a ticular project has a very high IRK, it may lack of widespread use and understanding. be unrealistic to assume that cash flow from the project can be reinvested and achieve as It is believed that the DTRR is a superior tool high a return. for project financial analysis. While all of the

II- 14 financialanalysis methods discussed in this section are incorporated in the Model, the DTRR is relied upon throughout this study for sensitivity 1. Input New Data analyses and choice among alternatives. 2. Recalculate Spreadsheet 3. Print Output 3.4 Other Financial Evaluations 4. Save and Quit

The above discussion does not include a descrip- Enter one of the above three choices tion of the Payback or Average Rate of Return techniques for investment analysis. These techniques are relatively simple to compute. How- ever, they do not take into consideration the time value of money or the relative risk associ- Figure II-l: Main Menu Of Financial Model ated with alternative investments. In addition, these methods are no longer believed by most that request information about the system be- investors to provide an accurate means of evalu- ing modeled. It is required that the user have a ating alternative investments. rudimentary understanding of the IBM PC and Lotus 1-2-3 operations. 4 Model User Manual The first step in accessing the Model is to load the Lotus software into the IBM PC. Once this is 4.1 Introduction done, the Model is retrieved from the floppy disk delivered as part of this study. The file name of The FinancialModel (the Model) is comprised the Model is CSSO. The command to retrieve the of a set of pro forma financialstatements and Model is given by typing "/Fit CSSO" followed investment analysestailoredto the satellitein- by a carriage return. dustry.Itis designedforuse on IBM PCs, XTs, and ATs using Lotus 1-2-3Release 2.0software. 4.2.1 Main Menu The user isrequiredto input certainfinancial and technicaldata describingthe satellitebeing Once the Model isretrievedfrom _sk, the main modeled. These inputsare made by responding menu isdisplayedon the screenas shown in Fig- to computer generated prompts forinformation ure If-1.The usertypes thenumber correspond- via displayedmenus. ing to the operationdesiredand pressesthe car- The output of the model consistsof an invest- riagereturnkey. ment analysis,financialstatements,and revenue, cost,financing,and tax assumptions covering 4.3 Model Inputs" the lifeof the project.The purpose of the output isto displaythe economic performance of a The Model is comprised of several layers of satelliteprogram from the perspectiveof a satel- assumptions incorporated into a spreadsheet liteoperator.Economic performance issumma- model to show economic performance of satellite rized on page 1 of the output by the various operations. Certain assumptions change more measures of return on investment. The follow- frequently than others and are incorporated into ing subsectionsdescribein detailthe operation the menus for update by the user. of the Model. The assumptions that are not incorporated into the menus are available for update by leav- 4.2 Accessing the Model ing the menu and updating the spreadsheet directly. For example, satellite cost changes fre- The Model requires an IBM PC, XT, or AT quently and is a selection possible in the menu. computer and Lotus 1-2-3 Release 2.0 software. However, the debt ratio of the operator doesn't The Model is "user friendly" and displays menus necessarily vary with type of system and is not

II - 15 GENERALASSUMPTIONS GENERAL ASSUMPTIONS

C K__u Other C Ku Other Are thereany transponders 1 0 0 Bandwidth (MHz) 36 54 72 of the following types? Power (W) 5.5 50 16 ("1" = Yes, "0" = No) Degradation curve 1 2 3 (!= C, 2=Ku, Note: If there are transponders of different 3 = hybrid, bandwidths, also use the Other column. 4 = user supplied) Utilization (%) 90 90 90 Number of Transponders C Ku Other Annual inflation (%) 4 Protected 18 0 0 Unprotected 4 0 0 Figure II-3: General Assumptions Menu 2 Pre-emptable 2 0 0

and Other. The Other frequency assumes the Figure II-2: General Assumptions Menu 1 revenue parameters for the C-band system. In- putting a "1" indicates that a particular fre- part of the menu. The insurance rate is another quency is present on the satellite and a "0" in- parameter that is not part of the menu and must dicates that the frequency is not present. Trans- be changed directly in the spreadsheet. ponders are input by typing the number of trans- The model is menu driven and prompts the ponders corresponding to each frequency. user for inputs to the most frequently varied as- Once a particular selection is made, the next sumptions as well as for file maintenance func- selection on the menu is made by moving the tions such as saving and printing. cursor using the arrow keys on the keyboard to User inputs are effected by" typing in num- the information requiring update and typing in the new information - either a "1" or "0" for bers corresponding to menu options or system frequency or number of transponders. characteristics (e.g., bandwidth, power, satellite cost). The numbers are input by placing the cur- Once all the desired selections are made, the information is incorporated into the Model and sor (a high-lighted cell on the PC screen) where the input is to be made, typing in the appropri- the menu moves to the next set of assumptions when the return key is pressed. This method of ate number, and pressing the "enter" key. This causes the information to be entered into the update of satellite information is the same for all menus. Model and also causes the next menu to be dis- The second General Assumptions menu, played. shown in Figure II-3,inputs bandwidth, transponder power, degradation curve, utilization, 4.3.1 General Assumptions Menu and annual inflationrateforoperatingexpense. Except forinflationrate,alldata correspond to Ifselection1 ismade from the main menu (Fig- the frequency bands previouslyselected. ure If-l),the firstof four input menus appears on the screen. Figure II-2 is the firstof two "General Assumptions" menus and requiresse- 4.3.2 Financing and Tax Menu lectionof frequencyband and number of operat- The next menu to appear is the Financing and ing transponders. Tax Assumptions menu shown in Figure II-4. It Up to threefrequencybands may be specified allows the user to determine the required return for an individualsatellite;C-band, Ku-band , on equity, capital structure, loan term, prime

II - 16 Financing and Tax Assumptions Capital Expenditure Assumptions

P_equired return on equity (%) 18.00 Satellite cost ($M) 38.23 % Debt in capital structure 45.00 STS launch cost ($M) 17.19 Term of loan (months) 60 Perigee stage cost ($M) 6.21 Annual average prime rate (%) 9.00 Launch support cost ($M) 1.64 Depreciation method ACRS Mission operations cost ($M) 2.55 Investment tax credit (%) 8.00 OMV/OTV cost ($M) 0.00 (0% or 8%) Space Station support ($M) 0.00

Operating Assumptions Figure II-4: Financing and Tax Menu

Month placed in service 36 rate, and investment tax credit election. The Useful life (months) 120 prime rate assumption is the basis for project interest rates, calculated at prime plus 2 points. Figure II-5: Capital & Operating Menu The depreciation method is fixed at ACRS (see Subsection II-2.6.2). Financing and tax assumptions do not nec- 4.4 Recalculate Spreadsheet essarily change with different satellite systems. Once the user has updated the Model for a par- The assumptions provided in the delivered model ticular satellite system, the spreadsheet is usu- and shown is Figure IL4 are believed to be repre- ally recalculated and then printed. sentative of the industry. Therefore it is recom- To incorporate updates of Model assumptions mended that these assumptions remain as they and to calculate revised economic performance, are unless a change is specifically desired. the user must recalculate the spreadsheet. This procedure is invoked by typing "2" on the main menu (Figure ILl) followed by "return". 4.3.3 Capital Expenditures and Operat- The recalculation procedure requires several ing Assumptions Menu minutes of computer processing time, depending upon the speed of the computer. While the The next menu to appear is the Capital Expen- computer is working, the message "wait" flashes diture menu shown in Figure II-5. It includes the in the upper right corner of the CRT screen. cost of the satellite, STS launch, perigee stage, When complete, the message "ready" is shown launch support, and mission operations. Space is and the user may select another item from the also provided for input of costs associated with main menu. At this point, the user would nor- OMV/OTV and Space Station support activi- marly print out the results. ties. Operating assumptions are displayed on the 4.5 Print the Spreadsheet same menu as the capital assumptions (Fig- ure II-5). These assumptions are comprised of 4.5.1 Model Output the number of months from the beginning of The output of the Model consists of nine spread- satellite construction to the beginning of satellite operation and the useful life of the satellite. sheet pages:

Once all of the assumptions have been up- 1. Financial Analysis dated, the user presses enter and the program returns to the .m___Anmen,_,.

II - 17 3. Balance Sheet is time consuming and may take several hours with a dot matrix printer or ten minutes with a 4. Sources and Uses of Funds laser printer. It is therefore recommended that 5. Revenue and General Assumptions only the spreadsheets of interest be printed. If "print support schedules" is selected, an- 6. Capital Expenditure Assumptions other set of printing options appear at the top of the screen: 7. Operating Expenditure Assumptions • Revenue 8. rT_. lnancing• &ssumpuonsa . • • Capital 9. Tax Assumptions • Operations The Financial Analysis page presents economic performance measures of dual terminal • Finance rate-of-return (DTRR), internal rate-of-return •Tax (IRR), and net present value (NPV) as well as the projects equity cash flows, required return • Return on equity, and reinvestment rate. The Income Statement, Balance Sheet, and This set of print options allows the user to se- Sources and Uses of Funds pages are typical fi- lect any of the Model assumption spreadsheets nancial statements. for printing. Once a selection is made and the The Capital Expenditures, Operating Expen- output printed, the user is returned to the main ditures, Financing, and Tax Assumptions pages menu. display the assumptions for the particular sys- If "print statements" is selected, another set tem being modeled. of print options appears at the top of the screen:

• All statements 4.5.2 Print Options * Analysis Item "3" is selected from the main menu (Fig- ure ILl) in order to print the spreadsheet. A set • Income statement of printing options is displayed at the top of the main menu when option 3 is selected: • Balance sheet

• Print all schedules • Sources and uses

• Return • Print support schedules

• Print statements These options allow the user to print all of the financial analysis and financial statements, work- • Return sheets or individual worksheets. Again, once a selection is made and printing is completed, the The user selects a particular option by placing user is returned to the main menu. the cursor by use of the arrow keys over the op- If the user selects the "return" option included tion desired and pressing the return key. The in all the print menus, the Model returns to the desired output then begins printing. menu previously displayed. While the output is printing, the user cannot access the Model and the word "wait" flashes 4.6 Save and Exit Spreadsheet on the screen. When printing is complete, the program returns to the main menu. The "save and quit" option on the main menu If "print all schedules" is selected, all nine (Figure ILl) saves the updated version of the spreadsheets are printed in sequence. Printing Model and exits the Model software.

II - 18 4.7 Use of Model The satellite bus costs are derived from costs associated with the CS-2 contract performed at The base model is used in Section III for sen- Ford Aerospace. (The C-band payload is de- sitivity analysis. For instance, it is possible rived from another satellite program.) Once the to change revenue assumptions by altering the parameters and the cost input converge on the satellite characteristics such as power or lifetime. cost of a CS-2 satellite system, the Price H out- Also, net cash outflows are affected by changes put is considered to be calibrated. By varying in individual capital expenditures and payment the parameters to those of the ttS-376, the satel- schedules. Changes in capital expenditures are lite costs for the Model are derived. The other also linked to financing considerations and asset capital costs are developed from current Ford size on the Balance Sheet. Aerospace experience or published prices. Sensitivity analysis can also be performed in As shown in Table A-1 of Appendix A, the the operating costs category. It is possible, for total capital expenditures of $76.5 M for the C- example, to change the annual rate of inflation. band satellite yields a DTRR return of 18.1%. If economic conditions either improve or worsen, Since the component requirements are validated new inflation factors can be incorporated into and the return is within the expected range, the the Model. Model is considered validated. Examples of the Model output are given next in Subsection II-5, with the actual output being given in Appendix A. 5.2 Ku-band Satellite

The nine pages of Table A-2 in Appendix A give 5 Model Outputs for 1985 the Model results for the Ku-band satellite de- Satellites scribed in subsection 2.1.2 and Table II-3. The DTtttt return is 19.8% with a total capital ex- Appendix A contains the Model outputs for the penditure of $104.3 M. The better return than three 1985 satellite systems. The Model outputs the C-band satellite is due to the higher revenues are validated by: from the high power Ku-band transponders. The Ku-band satellite costs are based on costs 1. Calibrating each capital expenditure ele- derived from two Ford Aerospace programs; In- ment; and telsat 5 and Ford Satellite. The costs are validated in a manner similar to the C-band satellite 2. Examining current experience for expected described above to yield an expected cost for a returns for each satellite system. RCA K2-type satellite system.

By far the single greatest capital cost is satellite construction. The costs are computed 5.3 Hybrid Satellite through the Price H cost model and are derived from the Ford Aerospace database and validated The nine pages of Table A-3 in Appendix A against existing satellite systems. The basis for give the Model results for the hybrid satellite de- the satellite cost validation and the Model results scribed in subsection 2.1.3 and Table II-4. The are described below. DTRR return is 21.9% with a total capital expenditure of $83.1 M. This satellite achieves the best return due to the combination of high rev- 5.1 C-band Satellite enues from a mix of C and Ku-band transponders The nine pages of Table A-1 in Appendix A give at a relatively modest increase in costs. the Model results for the C-band satellite de- The hybrid satellite costs, like those of the Ku- scribed in subsection 2.1.1 and Table II-2. The band satellite, are based on costs derived from DTRtt return is 18.1% with a total capital ex- the Intelsat 5 and Ford Satellite programs. The penditure of $76.5 M. costs are validated as described above to yield an

II - 19 expectedcost associated with a GTE Spacenet type satellite.

5.4 Discussion

These results achieved are in agreement with the current experience of satellite system owners and operators. Because of this and the validation process involved in calibrating each capital expenditure element, the Model is considered to be validated. Section III will further validate the Model by investigating its sensitivity to variation of system parameters, and verifying that the results correspond to the current experience of the industry.

II - 20 Section III IMPACT OF SYSTEM CHARACTERISTICS ON ECONOMICS

1 Introduction channel capacity:

This section describesthe selectionof system H- B log_ (1 + _) (1) characteristicsand the Task IB sensitivityanalysisof the impact of the system characteristics where on system economic performance. Subsection 2 gives the basis for the choice H is the link capacity of communication satellitesystem technicaland B is the link bandwidth functionalcharacteristicsthat significantlyaf- fectthe economic factorsand output valuesof C is the carrier signal power the financialmodel (the Model). These system characteristicsaredividedintothreegroups;pri- N isthe noisepower mary, secondary, and financial,which are discussed in Subsections3, 4, and 5. Bandwidth is chosen as a primary characteristic since communications capacity varies directly with it. 2 Choice of Characteristics C/N has a logarithmic influence on capacity; changes cause greater effects for small values of C/N than for large values. In practice, The fundamental purpose of a communications the multiple access technique and modulation satellite is to provide communications capacity. scheme determines the required minimum value Shannon's equation (Equation 1) gives the max- for C/N, and the system is engineered to exceed imum communications capacity of a channel. thisvalue. The satellite link equation, solved for the uplink and downlink cases, gives the technical factors 2.2 Satellite Link Equation that influence the carrier-to-noise ratio (C/No). Regulatory constraints in order to limit inter- The factors that influence C/N can be seen by ference are a third factor that impacts system consideration of the satellite link equation. The performance. fundamental equation for RF link performance is as follows:

2.1 Channel Capacity - (2) The value of a communications link depends on where its capacity, which is defined as the maximum rate at which information can be transmitted C is carrier signal power without error. Without telling how to achieve it_Shannon's equation givesan upper bound on No is the noise power per ,_,nit bandwidth

III- 1 EIRP (effective isotropic radiated power) Link performance depends on the receive an- is the transmit gain times power tenna efficiency and area, the transmit power, and the receiver system noise temperature, but k is Boltzman's constant not on operating frequency. (However, frequency is important for the other reasons of device per- Ls is the free space path loss formance, atmospheric losses, and interference Gr is the receive antenna gain with other systems.) It is the physical size of the ground receive antenna that is important, Ts is receiver system noise temperature nnt it_ _n ...... _.

The relationship for path loss is as follows: 2.2.2 Uplink Communications

/;s - R) 2 For the up[ink segment, if the satellite receive A2 (3) antenna coverage on the earth is specified, Equa- where R is distance and A is wavelength. tion (2) becomes: The relationship between the receive antenna C K1 (rlA,'_ gain, Gr, and effective area, At, is as follows: No - 4_r k S \ Ts ] P' (6) where Gr - 4_r _/Ar ,_2 (4) At is transmit (ground) antenna area where 77is antenna efficiency. Ts is receive system temperature of sat. The satellite communications link is made up of both down]ink and up[ink segments which are Pt is transmit power of the earth station analyzed in turn. Thus, for both up[ink and down[ink, performance depends on the ground antenna physical 2.2.1 Downlink Communications size and not on the transmit and receive antenna gains. An important constraint for FSS satellite com- The satellite transmitter power, rather than munications is that the satellite transmit an- EIRP, is the important factor since satellite antenna gain is fixed by the required earth coverage tenna gain is a constant for systems with fixed area. Using Equations (3) and (4), Equation (2) coverage areas. Power can be compared across can be rewritten as follows: frequency bands, provided atmospheric attenuation is taken into consideration. N0- 4rkS \ 7'8 ]Pt (5) 2.2.3 Technical Characteristics -@-here The factors influencing C/No can be seen from K1 is a constant equal to the transmit an- Equations (5) and (6). Satellite transmit power tenna gain times its solid angle is a primary characteristic. The ground terminal requirements including k is Boltzman's constant antenna size, antenna efficiency, system noise S is the coverage area temperature, and ground transmit power are secondary characteristics due to the limited relative is the receive antenna efficiency vab,_...... of the ground ,_ogrnont far 1QR.__ _j_xr_torn_...... The satellite receiver system noise temperature Ar is the receive (ground) antenna area is a secondary characteristic due to its limited Ts is receiver system noise temperature range of variation. The coverage area is a secondary characteristic due to constraints on its Pt is the transmit power of the satellite variation.

III- 2 2.4.1 Interference Rain Margin CITY Zone (dB) Interferencefrom unwanted signalsactslikesys- Eastport,Maine B 3.7 tem noiseto degrade performance. The maxi- Miami, FL E 10.0 mum allowablecarrier-to-interferenceratioC/I New York City D 6.3 depends on the toleranceof the communica- San Diego, CA F 2.0 tionssystem to interference.For instance,FM Seattle,WA C 3.3 modulation,used forTV transmissions,exhibits a strong "capture" effectwhereby the receiver locksonto the strongestavailablesignaland is Table HI-l: Ku-Band Weather Margins relativelytolerantto unwanted signals. As long as C/I is greater than C/N, the system is noise limited. Such a design is desirable in 2.3 Atmospheric Attenuation that it permits individual users to improve their quality of service by using lower noise figure re- The atmosphere attenuates the transmitted sig- ceivers. (Users could improve their performance nai and decreases C/N. The amount of attenua- even more by purchasing larger receive anten- tion depends on transmission frequency and path nas.) length through the atmosphere (i.e. elevation Interference can arise from either misdirected angle of satellite as viewed from the ground and energy (in antenna sidelobes) or from the out-of- altitude above sea level). There is both a fixed band energy of one channel falling in the band component of the attenuation due to molecular of another channel. An extensive body of regu- oxygen and a variable component due to water lations exist in order to keep interference at tol- in its various forms. erable levels. Rain, snow, and ice in the atmosphere also depolarize the signal, resulting in interference 2.4.2 Ground Antenna Sidelobes between the orthogonally polarized signals com- moniy used for frequency reuse. In order to allow 2° satellite spacing, the FCC At C-band, the normal atmosphere has negli- specified in the 1983 Orbital Assignment Order gible attenuation and heavy rain can cause up to the maximum sidelobe levels for ground transmit 2 dB loss. At Ku-band, the normal atmosphere antennas. Table III-2 gives the maximum allow- has 0.5 dB loss and heavy rain can cause up to able antenna gain in dBi as a function of off-axis 10 dB loss. angle 0 in degrees in the plane of the geosta- Table III-1 gives the link margins recom- tionary orbit and elsewhere (orthogonal to the mended for 99.9% Ku-band signal availability by plane). (Limits are also placed on cross polar- RCA Americom for commercial satellite opera- ization isolation by the FCC.) tors in different cities. Minimum transmit antenna sizes of around 80 wavelengths (4 m at C and 2 m at Ku-band) For 100% reliability, operations in areas of fre- are required to meet these FCC specifications. quent heavy rainfall, or operations at elevation For ground receive-only applications, it is the re- angles below 10 °, C-band has a considerable ad- sponsibility of the user to provide sufficient pro- vantage over Ku-band. tection from interfering signals.

2.4.3 Power Flux Density Limits 2.4 Regulatory Considerations WARC regulations place limits on the maximum The purpose of international and national regu- ground power flux densities (PFD) in any 4 kHz lations on telecommunications is to limit inter- bandwidth to -142 dBW/m 2 at C-band and - ference among different communications links. "1._ (]RW/m 2 _.t TCu-h:_n_ f_r _no'l_ nf _.v-;,_l

III- 3 There are no motivating factors to go to high GEO ORBIT PLANE power as the present day 34 dBW EIRP gives Maximum Off-Axis Angle 99.98% availability and allows 3 m TVRO re- Allowable (degrees) ception. Furthermore, the 2 ° satellite spacing Gain (dBi) From [ To places lower limits on ground antenna sizes. Ku-Band +8 7.0 9.2 In order to achieve the same performance as 3229-251og01- 25 log 0 9._1"021 48.07.0 In C-band with the same ground antenna size, more --u I 'i°'Ul lou.u power is required for the Ku-band signal to compensate for atmospheric attenuation (Sub- ELSEWHERE section 2.3). Transponder power is limited to 50 W for 32- 251og-10 # I 48.0180.01.01 48.0 1985 and an estimated 100 W for 1995 satellites, for. CONUS coverage and 54 MHz bandwidth. Table III-2: Ground Antenna Standards These correspond to 45 dBW and 48 dBW EIRPs. Interference considerations drive these limits. A satellite power of 50 W allows use of greater than 25o above the horizon. Equivalent EIRPs over a 36 MHz bandwidth are 60 and 2 m earth terminals for two-way, low data rate customer premise services. There are diminish- 64 dBW respectively. ing returns with higher satellite power in that However, satellites providing CONUS (conti- more ground transmit power is required to offset nental United States) coverage from an extreme the smaller ground antenna. orbital position such as 65 ° W or 130 ° W are 10° or less above the horizon at the opposite north- ern CONUS corner. The PFD limit drops up to 3 Primary Characteristics 10 dB for low elevation angles. (The additional path loss and atmospheric attenuation loss at 10 ° The following primary factors that most directly are 2 dB and the EIRP limits only drop 8 dB.) affect the satellite economics have been identi- The effective operational limits are even lower. fied: The 1979 ITU Radio Regulations for Region 2 1. Communications capacity of satellite limit Ku-band EIRPs to approximately 50 dBW maximum for FSS. (The DBS band is set aside 2. Bandwidth of channel for higher power, direct transmissions of TV to customers.) 3. Frequency band of operation The important consideration is to limit inter- 4. Power of channel ference among satellites. Higher power in one satellite translates to higher interference in its 5. Utilization of satellite neighbors. The regulatory trend is to keep all Satellite costs are typically around one third of satellites to the same radiated power per unit bandwidth. the revenues. Thus, those characteristics that influence revenues are more significant than those C-Band that influence cost and have greater effects on The emphasis is on low cost and low risk for the rate-of-return to the operator. medium, to high ao,_t,,...... j ,._, . Op erational 1;,_.... _,_,_,_._ ,a_o,s, discussed ;-'lxt c.._.o^_.:^.tJ tt uo,ot._ tl.t ult ,TTJk_JL-- its on transponder power for 1985 satellites, are 2.3.3, have a great impact on transponder price 8.5 W for CONUS coverage and 36 MHz band- and hence economic performance. It is beyond width, and may increase slightly to 10 W for the scope of this study to discuss and analyze 1995 satellites. This corresponds to 36 dBW their impact. However, the results for variation EIRP. in utilization as given in Table III-6 are the same

III- 4 as for market factor variation. (For instance, a 25% drop in transponder price is equivalent to a Revenue Return 25% drop in utilization.) Satellite Capacity $M per (DTRI_) The market factor may change with time. For Type MHz transponder % instance, the present conditions of a plentiful Telstar 864 1.6 18.1 supply of C-band transponders and a shortage K2 864 2.0 19.9 of high power Ku-band transponders is not likely Spacenet 1,296 1.2 21.9 to last through 1995.

Table III-3: Capacity Versus Cost

3.1 Communications Capacity 3.2 Bandwidth of Channel

Communications capacity is determined by the The bandwidth of the communications channel total bandwidth used by the satellite, including is directly related to the communications capac- possible frequency reuse via two polarizations ity of the channel as per Equation (1). Thus, and multiple spacially separate beams. it is expected that transponder price is directly The World Administrative Radio Conference proportional to bandwidth: (WARC) 1985 Allotment Plan would expand the Price c< k (Bandwidth) presently available 500 MHz of bandwidth by 300 MHz to 800 MHz at C-band and by 500 MHz where k = 1. However pricing data suggests a to 1 GHz at Ku-band. Since this is still in the lower value, k = 0.9, which we use in the Model. planning stage, for purposes of this program the One reason for this is the technical difficulty of present 500 MHz allotments are used. passing a greater number of signals through a A satellite design is optimized in terms of size single transponder that is nonlinear. and power for its communications capacity. To Bandwidths smaller than 36 MHz are unlikely reduce capacity for a particular design does not to be available for FSS service by 1995 due to make economic sense. To increase capacity re- the increased manufacturing costs and limited quires increased mass, power, and launch capa- demand. bility, typically meaning a new satellite design. The total available communications band- Thus, analysis of communications capacity width is the limited resource. The choice is one must compare different satellite designs. Ta- of number of channels, i.e. 16 @ 54 MHz or 24 ble III-3 compares the economic performance of 36 MHz, and the total satellite revenue tends to the three 1985 satellites. The revenues are for remain the same. Note that the total power of all equivalent 36 MHz transponders. The different transponders also remains the same in order to revenue figures also reflect other features of the keep the power radiated per unit bandwidth the satellites such as different transmit powers and same. However, satellite cost and mass increases operating frequencies. The dual terminal rate of as channel bandwidth is reduced on account of return (DTRR) is the bottom line financial re- the increased number of components. sult, reflecting satellite manufacturing, launch, and operating costs. 3.2.1 C-Band Channel Bandwidth

The conclusion is that due to economies of C-band FSS supplies low cost, low risk commu- scale, satellites with greater communications ca- uications for medium to high density traffic. The pacity yield better rates of return on the capital 500 MHz available band is packaged in 36 MHz invested. However, there are limitations on the bandwidth units for the historical reason that maximum capacity of one satellite due to launch it allows FM transmission of a single TV chan- vehicle constraints, insurance limitations per sin- nel. The 36 MHz bandwidth can also transmit gle launch, and market for transponders.

III - 5 Use of this standard transponder bandwidth will - Ku-band requires from 2 dB to 8 dB more continue, with some larger channels available for weather margin than C-band links (Subsec- wider band transmissions. tion 2.3).

3.2.2 Ku-Band Channel Bandwidth Note that the the link equations given by Equations (5) and (6) contain ground antenna Both 54 MHz and 72 MHz bandwidths are size and not gain. Thus, for a given desired in use at Ku-band. Bandwidths wider than system performance, ground antennas at Ku- 36 MHz were used because of satellite mass and band must be larger than those at C-band on ac- power limitations. Some satellite designs could count of rain attenuation. The reason Ku-band not carry enough payload for 24 transponders ground antennas are generally smaller is that (36 MHz each), but 16 transponders (54 MHz more power may be transmitted by the satellite. each) were possible. The relative value of Ku-band and C-band With 1995 technology, this limitation is re- transponders of similar power and bandwidth moved and smaller bandwidths such as 36 MHz can be estimated by taking 2 dB from the Ku- will be used for FSS service. This facilitates in- band transponder power to account for added terconnection between C and Ku-bands. Band- weather effects. The 2 dB power assessment widths smaller than 36 MHz are unlikely to be against Ku-band is equivalent to a 0.9 relative offered because of the inefficient use of band- price for Ku-band versus 1.00 for C-band (see width due to increased guard bands. The non- Subsection 3.4). (If 100% availability is required, linear characteristics of the TWTAs required for there is an 8 dB difference, and a 0.65 cost fac- high power operation makes the use of very wide tor.) bandwidths inefficient. The conclusion is that these points counterbal- ance each other and the frequency band factor is 3.3 Frequency Band of Operation 0.80 for Ku-band versus the baseline 1.00 for C- band for 1985 and 1995 satellites. The frequency The need for greater communications capacity factor may change with time, particularly if FCC than that offered by the 500 MHz at C-band regulations change. led to the utilization of the 500 MHz bandwidth at Ku-band and leads to use of the 2,500 MHz 3.4 Power Transmitted in Channel bandwidth at Ka-band. The relative advantages and disadvantages of In the past, satellite mass and primary power Ku-band compared to C-band are as follows: constraints have limited available transmit power. New, larger satellites overcome these + Ku-band has higher allowable transmit constraints and now the scarce resource is avail- power. able communications bandwidth. Power is primarily limited by considerations of interference + C-band transmissions are regulated to avoid among satellites and efficient use of bandwidth. interference with existing ground microwave Another important factor is to make efficient facilities. Ku-band does not have this prob- use of bandwidth. Higher power transponders, lem. depending on modulation method, may "waste" + C-band typically uses SSPAs which are bandwidth by requiring use of more dispersal or more linear than the TWTAs used at Ku- spreading waveforms for the carrier. (It has been •o_na,"1 and can resm_1, in greater channel capacity (see Subsection 4.11). use of a 25 W C-band transponder for CONUS coverage.) Ku-band has less mature technology than Table III-4 gives typical values for "protected" C-band with the consequent greater costs transponder price as a function of power. Data and risks. is presented separately for the two bands, refer-

III- 6 Relative Transponder Price K2Mass(kg)-] Power (S/month) Satellite Subsystem 0Wl 25W I C-band Ku-band (dB) Power 233 166 0 150,000 300,000 ADCS 39 39 -3 122,000 244,000 TT&C 27 27 -6 198,000 Propulsion 100 96 Structure/thermal 159 145 Comm. payload 175 175 Table III-4: Transponder Price versus Power Harness 40 40 RCS (10 yr) 242 164 enced to different baselines. The 0 dB or baseline TOTAL (BOL mass) 1,015 852 power are 8.5 W for a 36 MHz C-band SSPA Deployed Mass (STS) 5,538 4,775 transponder and 50 W for a 54 MHz Ku-band TWTA transponder. The data was obtained by comparing a number of satellites. Table IIt-5: Effect of Power Reduction on Mass Transponder price varies according to the transmit power per unit bandwidth: transponder price. This can be done by adjust- f Power _x ing the transponder power input to the Model Price 0¢ \ B_tl_ J such that it is in the same ratio to the base- where x = 0.33. Future Ku-band services are line satellite power as is the EIRP. The baseline planned using high power satellites and small 8.5 W C-band satellite with CONUS coverage has 36 dBW EIRP. The baseline 50 W Ku-band ground terminals. For these systems, the ground segment may consist of 100,000 terminals and satellite with CONUS coverage has 45 dBW EIRP. represent a larger investment than the satellites. In this case, transponder price could depend For satellites with regional beams, the actual more strongly on power, perhaps x = 1. power must be adjusted according to the coverage area. For example, 50 W into a half CONUS 3.4.1 Impact on Satellite beam would be input into the Model as 100 W.

Table III-5 summarizes the impact on satellite mass of a 3 dB reduction in transponder power for the Satcom K2 satellite. The 16% reduction in BOL mass was reflected by an equal percent- 3.5 Utilization age reduction in satellite cost as determined by the Price H cost model. This is consistent with Utilization is the percentage of the satellite's the transponder price data. communications capacity that is sold or leased, and thus has a direct effect on economic perfor- 3.4.2 EIR.P mance. Table III-6 gives the rate of return for As pointed out in Paragraph III-2.2.2, satellite a K2 type satellite for different utilization rates. power rather than EIRP is the important factor In order to achieve the targeted 18% rate, 85% for links with fixed coverage area since antenna utilization is required. gain is a constant. This is the case as utilized by Both dual terminal (DTRR) and internal rate the Model; fixed CONUS coverage is considered. of return (IRR) are plotted. The DTRR invests If the satellite system utilizes frequency reuse profits at the target 18% rate, while the IRR via spatially separate beams, the beam EIRP assumes investment of profits at the same rate rather than power must be used to determina being re_ized from the sate!_te.

III- 7 4.1 Attitude Control and Pointing Satellite Total Rate of Return Utilization Revenue The accuracy of satellite attitude control and an- (%) ($M/yr) IRR(%) I DTRR(%) tenna pointing depends on the size of the individual antenna beams or service coverage areas. 40 21.3 0.5 9.7 Edge-of-coverage gain and isolation between dif- 50 26.6 5.4 12.2 ferent beams are affected by pointing inaccuracy. 60 31.9 9.5 14.2 In general, pointing accuracy requirements in- 70 37.2 13.2 15.9 crease with frequency of operation due to the 75 4, ,, ,,IL"t.,_ JLU. feasibility of obtaining smaller beam sizes from 80 42.5 16.5 17.4 the same size antenna. More spatial frequency 85 45.2 18.0 18.0 90 47.8 19.7 18.6 reuse also increases pointing accuracy requirements due to the smaller coverage areas of the 95 50.5 21.0 19.2 individual beams. Accuracies of 4-0.1 ° are ade- 100 53.2 22.4 19.8 quate for C-band and 4-0.05 ° for Ku-band. Higher pointing accuracy requires more fre- Table III-6: Utilization Versus Rate-of-Return quent satellite station-keeping adjustments. Al- ternately, active steering of the antenna could be 4 Secondary Characteristics used. In either method, greater satellite costs are incurred. Since the attitude control and point- The following set of secondary factors are dis- ing requirements follow directly from the choice cussed below: of payload, this item is not of primary impor- tance. 1. Attitude control and pointing accuracy 4.2 Coverage Area 2. Coverage area Service coverage areas are determined by the lo- 3. Ground terminal requirements cation of traffic and the number of frequency reuses. A single reuse covers CONUS with one 4. Lifetime of satellite horizontal and one vertical polarization beam. 5. Mass and volume of satellite Further reuse divides CONUS into a number of alternating H and V beams. The coverage area 6. Maturity of technology shapes must be matched to the communications traffic. 7. Noise temperature of satellite receiver The key tradeoff is the desired communications capacity. Greater capacity requires more 8. On-board switching coverage areas which generates more revenue but cost more to provide. 9. Primary power of satellite

10. Reliability of satellite 4.3 Ground Terminal Requirements

11. Time delay in launch The scope of the present work excludes ground terminals and terrestrial systems, except for 12. Tr_u_i,...... uuu_,.1__ r_-ype.... TT£rC _'_,-111¢_,_=from consideration in the financial model. Thus, economic analysis is not done The majority of these secondary characteristics for the ground terminal, but a brief discussion is are interrelated with each other and the primary included for completeness. characteristics, and thus are not amenable to As seen from the link equations, satellite quantitative analysis. transmit power is inversely related to ground

III- 8 antenna area and performance. The impact of the ground terminal on economic performance Satellite Satellite Mass depends on the relative costs of the space and Lifetime (kg) ground segments of the system. (yr) Spacenet Satcom K2 The following characteristics of the ground 7 670 974 terminal are important: i0 710 1,018 12 750 1,083 • Antenna efficiency

• Antenna effective area Table III-7: Satellite Lifetime Versus Mass • System noise temperature The analog modulation schemes in use at C- • Transmit power band do not allow a user to take advantage of a Improvement in C/N can be obtained by im- larger or smaller ground antenna. A certain size proving satelliteperformance or ground termi- antenna is required to achieve the required C/N, nal performance. Less expensive (i.e.smaller) and increased size achieves only a small improve- ground terminalsare indicatedforsystems with ment in quality. Digital modulation schemes, largeground segments. likely to be employed at Ku-band, will allow variation in communications capacity with C/N. A 4.3.1 C-Band Ground Terminals mix of different size ground terminals could use the same satellite, each using the appropriate ca- The C-band ground terminal is typically a pacity. Class B 10 m antenna. The present large number of C-band antennas for TVRO (TV receive only) 4.4 Lifetime will move to Ku-band in the future. The use of C-band will be for high priority voice communi- Satellite Lifetime is determined by the lifetimes cations where weather outages are not tolerable. of its critical components and the amount of Therefore, C-band will continue to use relatively station-keeping fuel that is carried. Components large and expensive ground stations and will not with short lifetimes can be utilized if there is have low cost ground terminals as a cost driver. enough on-board redundancy to replace failures occuring during the design life of the satellite. 4.3.2 Ku-Band Ground Terminals Current satellite designs are optimized regard- Ku-band typically uses 6 m antennas for busi- ing payload, mass, and lifetime. Thus, the inclu- ness services - two-way video conferencing or sion of more mass for redundant components in multiple voice circuits. The benefit of a higher order to extend life requires a mass reduction elsewhere. power satellite transponder in reducing ground antenna size is offset by the uplink requirement The use of a larger upper stage to launch for increased ground transmit power. A further the heavier satellite is not economically feasi- minimum on antenna size is determined by the ble. Typically the satellite is designed at the 2° satellite spacing. mass limit for the upper stage, and a larger up- TVRO business is coming to this band, and per stage does not exist or is grossly oversized. the issue of small low-cost ground terminals is Table III-7 shows the variation in satellite very important in terms of being a cost driver. BOL (beginning of life) mass with lifetime for the (TVRO is not as sensitive to outages as voice, 1985 satellites Spacenet and K2. Mass changes and thus Ku-band with its weather outages is by +5% as lifetime is varied from 10 years to more appropriate for TVRO.) There will con- 7 or 12 years. The impact on satellite cost is tinue to be large 10 m ground terminals used for approximately the same. high density trunking,

III - 9 4.5 Mass and Volume Lifetime Rate-of-Return, % The satellite communications capacity deter- (yr) DTRR IRR mines the satellite mass and volume. Thus, in- 9 22.0 27.8 10 21.9 28.1 creased capacity gives increased mass which in turn leads to increased revenues and increased 11 21.8 28.2 costs. Larger satellites tend to have lower costs per unit communications capacity. Table III-8:Rate-of-Return V_.rsuaI,ifptime Limitations on mass and size, and thus pav- load, are placed by the launch vehicle mass capacity and size envelope. Limitations may also from an overall design tradeoff of component life- be placed by insurance capacity for a single pay- times, satellite mass, and launch costs. For 1985 load and transponder market capacity. satellites this optimum is 10 years. Longer lifetime satellites are not feasible due to launch ve- 4.6 Maturity of Technology hicle constraints an the unfavorable economics of including and paying for redundant equip- New technology initially has lower reliability, ment that does not generate revenue for 10 years. and is incorporated into the satellite design only after careful testing. This process results Short lifetime satellites are not economically fea- from the inherent conservatism of the commer- sible since there is little reduction in cost but a cial satellite users and their bankers and insur- large reduction in revenue. A further problem with longer lifetimes is technical obsolescence. ance brokers. Higher frequency band technology is typically less mature than that at lower fre- A quantitative analysis of a 12 year lifetime quency bands, and costs more to use. 1995 hybrid satellite was done and compared with the baseline 10 year satellite. The 1995 4.7 Noise Temperature of Receiver satellite design was used on account of the requirement to use the OTV for upper stage launch The total system noise temperature is the sum of the heavier (+65 kg) 12 year design. of many components:

The satellite cost increased $4.3 M, chang- • Receiver noise figure ing capital cost from $138.8 M to $144.3 M. The DTRR return decreased .5% from 21.9% to • Feed and beamforming network losses 21.4% for the longer life satellite. The reason • Radiation from ground or spacecraft for the poorer performance, in spite of two additional years of revenue, lies in the requirement • Radiation from atmosphere for redundant components to be purchased and As seen from Equation (5), a decrease in incorporated into the satellite, but not used for receiver system noise temperature directly im- 10 years. proves C/No. Consistent with reliability, the A further sensitivity study was done by keep- satellite receiver utilizes the best available LNA ing capital costs the same but changing lifetime (low noise amplifier) in the receiver. Cooled re- from 9 to 10 and to 11 years. For the 1995 hybrid ceivers are as yet too heavy and unreliable for satellite, Table III-8 shows how the DTRR and space use. Other contributions to system noise IRR rates-of-return change. The IRR improves temperature such as antenna pattern and feed

rJy ou.k,,.....:*"'..... ':_-*" _ae m_- noiseare a_u IUlIIAILli_2U..L11.¢¢conclusionisl;ll_•' [, counted value of revenues far in the future. The althoughimportant,the satellitereceiversystem DTRR actually decreases slightly. As shown in noisetemperature is optimized and willnot be Subsection III-4.11, a time delay at launch has varied. much more effect than a similar time loss at the For the ground station on the downiink perfor- end of life. mance, there will be variation depending on its

III - 10 Relative Price I Frequency I Transponder Power (W) Type of Transponder Band Type 1985 I 1995 Interconnection (%) Interconnected pair 101 C TWTASSPA [] 1,200500 I 410[660[ Hybrid pairs gu TWTA 13,60012,500 I C receive, Ku transmit 140 Ku receive, C transmit 120 Table III-10: Payload Power Requirements

Table III-9: Price of Interconnectivity divides CONUS into smaller regions, each requiring less power. size and cost. Smaller, low cost ground stations Assuming a baseline of full use of the band have less expensive receivers. Large ground sta- (i.e. 24 transponders of 36 MHz bandwidth) and tions may have cooled receivers. However, the use of the maximum allowed transmit power, the contribution from the warm ground and the at- required satellite primary power depends on the mosphere keeps a lower limit on practical system following factors only: noise temperatures. • Transponder efficiency 4.8 On-board Switching • Antenna efficiency Technology is allowing the feasibility of other • Coverage area than satellites with "bent pipe" transponders, but marketing is needed to establish demand. • Frequency band(s) covered On-board interconnectivity between different beams at the same frequency or different chan- Table III-10 gives the total payload power re- nels at different frequency bands appears to be quirement for CONUS coverage, assuming am- attractive to users. Note that bandwidth must plifier efficiencies appropriate for 1985 and 1995 be the same at C and Ku-bands in order to allow satellites. There are 24 C-band transponders of interconnection. 36 MHz bandwidth and 8.5 W, and 16 Ku-band Table III-9 gives an estimate of the relative transponders of 54 MHz bandwidth and 50 W transponder price for the different types for in- power. Power is included for receivers, switches, terconnectivity. The price refers to the percent- and payload control. age of the standard C and Ku-band transponder The cost of providing power is in the solar price. cells, the batteries required for eclipse opera- On-board processing such as planned with tion, and the thermal control subsystem. An in- NASA's ACTS satellite is judged to be too ex- crease in power directly increases satellite mass pensive for commercial application in the 1985 and cost. Since only one third of its solar cells to 1995 time-frame. face the sun at one time, the spinner satellite is less efficient than the three axis satellite de- 4.9 Primary Power sign in terms of mass required to supply large amounts of power. Satellite primary power requirements are mainly determined by the payload, in particular the 4.10 Reliability number and power of the transponders. How- ever, for a given coverage area and frequency Reliability has the same consideration as life- band, there is a regulatory limitation on total time: redundant components increase reliability transmit power regardless of the number of spa- at the expense of increased satellite mass. Cur-

l_lLb _d,b_Clllbe¢ _IU_I_IL_ are U_I, IIILIZ_U Wll, lL l'_i,l'U

III - 11 Class of Price($M/year) Time Delay Rate-of-Return,% Transponder C-Band Ku-Band (months) DTRR IRR Protected 1.90 3.6 0 21.90 28.06 Unprotected 1.40 2.6 1 21.50 27.70 Preemptible .90 1.7 3 21.31 26.96 6 21.00 25.84 9 20.65 24.71 Table Ill-11:Transponder Price J._ to probability of failure and redundant compo- Table III-12:Effectof Launch Delay nents. The degradation curves for the different satellite types have been given in Table IL7. Transponders are priced in three categories - protected, unprotected, and preemptible - ac- % Increasein cording to degree of protection from failures. Channel Capacity Spares are used while available to replace failed Modulation Method SSPA vs. TWTA components as required for all customers. As FDM (single access) 4 further failures occur, first the preemptible and FDM (multiple access) 10 then the unprotected customers are removed re- (2 carriers) gardless of which particular transponder fails. FDM (multiple access) 25 Table III-11 gives current prices of the differ- (6 carriers) ent classes of transponders for C-band and Ku- SCPC 50 band satellites. A five year lease with monthly payments is assumed. Baselines are a 8.5 W C- Table III-13: SSPA Transponder Capacity band SSPA transponder with 36 MHz bandwidth and a 50 W Ku-band TWTA transponder with 54 MHz bandwidth. 4.12 Transponder Type

4.11 Time Delay at Launch The greater linearity of the SSPA compared to the TWTA results in increased channel capac- A time delay at launch has a relatively large ef- ity. Table III-13 shows how channel capacity de- fect on rate-of-return since the satellite capital pends on modulation method and access meth- expenditures have been made but no revenue is ods. The nonlinear nature of the TWTA requires being generated. Table III-12 shows the effect it to be operated at reduced power when multi- on the rate-of-return (DTRR and IRR) of time ple signals are present. The increased linearity delays at launch from one month to one year. of the SSPA results in less intermodulation dis- The 1995 hybrid satellite design was used. A tortion of other signals. 1% decrease in DTRR is equivalent to an initial capital expenditure of $12 M. The improved linearity of the SSPA results in +1. _k;l;+., ,_ +. - _:+ + .... +_1^..'_: _ - - 1-- Th_ ray assumption is that the ,_.....;_*;--'_ _lLe C,I,UJ..L,ILIsJ _',J blalLSlttJb bWU b_l_VJkoJ.Oll _h_nne,_ (ACRS) is written off the first year regardless of through one 36 MHz transponder, and the use whether the satellite is launched that year. The of single sideband amplitude modulation (SSB- result would be even worse if the time delay were AM) techniques to transmit as many as 6000 such that the depreciation had to slip to the next one-way voice channels through one transpon- fiscal year before being utilized. der.

III - 12 cash outlays, total project costs are lowered on Debt Interest Rate of Return a present value basis. This benefit of deferring Ratio Expense IRR DTRR costs is more than enough to offset the added % $M % % interest expense, less taxes, arising from the use 0 - 17.1 17.7 of debt. 25 12.6 19.3 18.5 Satellite project risks also increase with debt 45 22.6 21.9 19.6 because the cost of debt is lower than the cost 50 25.2 22.8 19.9 of equity. It is better for shareholders to bor- 75 37.7 29.5 22.7 row money from the bank at 11% (5.6% after 100 50.3 54.9 31.0 tax) than to use their own funds which carry an implicit cost of 18% and are not tax deductible. This sensitivity analysis varies the amount of Table III-14: Return Versus Debt Ratio debt, holding all other economic factors constant, and leads to the conclusion that satellite 5 Financial Characteristics operators achieve the best return by financing 100% of the project capital cost. This conclusion The influence of the following financial charac- is not supported by reality, however, since it is teristics on system economic performance is in- rare to find projects 100% leveraged. This is due vestigated: to the bank's unwillingness to finance most of a 1. Debt ratio project, the higher borrowing rates accompany- ing higher debt levels, and the increased cost of 2. Insurance equity due to increased shareholder risk. 3. Interest rate The existence of an optimal capital structure (debt versus equity) for businesses is debatable. 4. Period of financing Some empirical studies show that returns do increase with leverage. This increase is limited, 5. Taxes however, and the assumption is usually made that the optimal capital structure for a particu- 5.1 Debt Ratio lar business is that which is used by its particular industry. The debt ratio of 45% used in this Table III-14 shows the effects of varying the per- study is representative of the satellite industry, centage of debt (debt ratio) used to finance the and is assumed to be at an optimal level. capital expenditures for the 1985 Ku-band satellite. The interest expense is stated in incurred 5.2 Insurance economics and the percentage of debt assumed in the Model is 45%. Table III-15 shows the variation of rate-of-return The result is that a 25 point increase in debt with launch insurance cost for the 1985 Ku-band ratio causes a 0.8 point increase in the DTRR satellite which has a total cost of $108 M. Insur- and a $12.5 M increase in interest expense. ance cost is calculated using incurred economics. Higher rates of return accompany increased debt For 1985 satellites, the average rate is 14%, levels primarily because equity cash outflows are c.urrent quotes for launch insurance are around deferred and the cost of debt is lower than the 30%, and the projection is for rates to return to cost of equity. 20% for 1995 satellites. Rates above 30% lead to By financing, the equity shareholders are able self insurance, and rates below 10% are unlikely to defer cash outlays for the capital cost of the due to the inherent risk of relatively new space project. Because all of the cash flows are ad- operations. justed for the time value of money, expenditures The potential variation in insurance rates is in beginning periods are worth more than those likely to be 10 points: for example, a reduction _rr_rn 9N@/_ *_ 1 flO/'_ f .... _ _" _¢^, 1 ..... 1_ ,,_1,:,1^ in later neriofls. Ther_fnrp hv rh_f,_rr;,_,T.... ;t_, _VALA _v/v v_ _.#/v avL _o_._ v_ a O_bA_A A_btA_L%AL V_;ALI_A_

III - 13 Insurance Insurance Rate of Return Period of Interest Rate of Return Rate Financing Expense IRR DTRR % Cost$M IRR]% DTRR% yrs SM % % 0 - 25.0 20.8 1 2.9 19.2 18.5 5 4.9 23.9 20.4 2 5.8 19.9 18.9 10 10.2 22.8 20.0 3 8.7 20.6 19.2 15 16.0 21.7 19.5 4 11.7 21.3 19.4

ZU 22.3 20.6 19.1 5 14.8 ZI._ l_.O 25 29.1 19.4 18.6 6 18.0 22.6 19.8 3O 36.3 18.2 18.1 7 21.3 23.2 20.0 35 43.9 17.1 17.6 8 24.6 23.8 20.1 9 28.1 24.3 20.3 10 31.7 24.8 20.4 Table III-15: Return Versus Insurance Cost

Table III-17: Return Versus Financing Period Interest Interest Rate of Return Rate Expense % SM IRR]% DTRR% 9 11.9 22.2 19.7 Interest rates vary with economic conditions and are difficult to forecast with confidence. The 11 14.8 21.9 19.6 13 17.7 21.7 19.5 assumptions for interest rates used in the Model 15 20.7 21.4 19.4 is the prime rate at the end of the first quarter 17 23.8 21.2 19.3 of 1986 plus 2 points (11%).

Table III-16: Return Versus Interest Rate like the OTV. This would translate to a $12 M 5.4 Period of Financing savings in insurance cost and a 1 point increase in the DTRR return. Table III-17 shows the results of varying the period of financing for the 1985 Ku-band satellite. 5.3 Interest Rate The interest expense is stated in incurred economics and the base assumption for period of Table III-16 shows the results of varying the in- financing used in the Model is five years. terest rate for the 1985 Ku-band satellite. The interest expense is stated in incurred economics On the average, a one year increase in pe- and relates to the permanent financing. The riod of financing results in a 0.2 point increase base assumption in the Model is an interest rate in the DTRR and a $3 M increase in interest of 11%. expense. The DTRR increases despite increased The result is that a two point increase in in- interest expense because a longer period of fi- terest rate causes a 0.1 decrease in the DTRR nancing allows the equity shareholder to defer •_.,,,, ,_,,,, a $2.9 M increase in :...... 1 ...... 1-_ ^...4.1 .... C ..... :4-_1 :_ 1 ...... pense. The small decrease in DTRR is due to overall costs on a present value basis (see III- the dollar magnitude of interest expense relative 5.1). The potential range of variation for period to total project cash flows, and the deductibility of financing is thought to be two years, with the of interest which makes a 1 point rate increase Model assumption of five years being the maxi- equivalent to 2 points after taxes. mum period.

III - 14 5.5 Taxes 6 Discussion

5.5.1 Depreciation The conclusion is that current satellite designs are optimized for their technical characteristics The baseline Model uses accelerated cost recov- in order to maximize revenues. The trend is ery (ACRS) depreciation. For purposes of com- to build larger satellites with increased capacity parison, a straight line depreciation over a pe- and economies of scale. riod of ten years is analyzed. The result, for the 1985 Ku-band satellite, is a drop in DTRR of 1.4 points. Accelerated depreciation permits the operator to realize the greatest immediate benefit for tax purposes. By spreading the benefits of depreciation over the life of the satellite, the net income cash flows axe considerably smaller than if an accelerated method is elected.

5.5.2 Investment Tax Credit

The investment tax credit (ITC) is a direct reduction in tax liability calculated as a percentage of capital expenditures. It is designed to provide an incentive for capital investment. The Model assumes an 8% ITC taken when the satellite is placed in service. There is currently proposed legislation to dis- continue the ITC. For the Ku-band satellite, removal of ITC results in a tax expense increase of $8.6 M in incurred economics and the DTttR declines 0.9 points.

5.5.3 Loss Carry Forward

Depending on a corporation's capital and operating structures, it may carry forward its losses for tax purposes or pass them through for corporate use. To achieve the maximum return on investment, these losses should be passed to the corporation assuming the corporation has enough profits to offset the losses. The method used by the Model is to assume that the losses are used when incurred. For purposes of comparison, use of the loss carry forward method was analyzed for the Ku- band satellite and resulted in a 2.9 point decrease in DTttR. Immediate tax loss write-off is much

III- 15 Section IV

TECHNOLOGY FORECAST FOR 1995 & DEFINITION OF FSS SYSTEMS

1 Introduction 1. Attitude Control

This section presents the following Task 2 results 2. Communications Payload for 1995 satellites: 3. Primary Power • Technology assessment and forecast;

• Description of three FSS systems. 4. Propulsion

In addition, a large (2,200 kg) high-capacity 5. Structure and Mechanisms satellite design is postulated which takes advantage of the launch capacity of the space-based 6. Telemetry, Tracking, and Command OTV. The next section (Section V) will use the Fi- 7. Thermal Control nancial Model to analyze the performance of the Also assessed will be the following: three postulated FSS (fixed satellite service) systems. This will form the baseline for comparison 8. Space transportation with the Space Station APO Scenarios of Tasks 3 and 4. 2.1 Attitude Control Subsystem 2 Technology Assessment The business-as-usual attitude and orbit control subsystems contain a variety of autonomous and An assessment is made of the expected state-of- manual control modes. During basic on-orbit op- the-art status of communications satellite sys- eration, as many as 20 people (over three shifts) tems and operations for U. S. domestic FSS are required for manual control of attitude and systems initially entering service in 1995 - ini- orbit parameters. The number of people could tial operational capability (IOC) in 1995. The be reduced to 6 if some basic jobs could be assessment is constrained to a consideration of made autonomous. The station keeping or or- business-as-usual satellite system operations em- bit control, which has the highest manual con- ploying evolutionary improvements in satellite trol requirements, could be made autonomous system technology and the allied field of space by advances being made in navigation and com- transportation. The business-as-usual constraint puter software and hardware. The reliability of means that Space Station or space-based opera- this type of autonomy is currently being studied tions will be excluded from the assessment. This through various programs. assessment will be used to develop, in the next Use of the TDRSS relay satellites would al- subsection, a description of three satellite sys- low contact to be kept with the satellite even tems representative of the 1995 state-of-the-art. when on the other side of the earth from the The assessment will consider each of the seven control station. Use of the planned global posi- rnmmllnle:_.tlnn_ _t_ll;to R11hQyf_mR.

IV - 1 allow constant and more accurate position determination. Together, GPS and TDRSS will sim- Frequency ]Efficiency,% [ plify the support required for initial positioning Band in orbit and for attitude determination during the lifetime of the satellite. Ku The recently-developed ring laser gyroscope Ka has several advantages over conventional gyros:

• Higher accuracy; Table IV-l: TWTA Efficiency

• Reduced calibration time; and reuse willrequiresmallerbeam sizeswhich re- • Quicker start-up. quirelargerdiameter antennas. Use of the Shut- tlelimitsthe diameter of a solidreflectorto 15 ft At present the ring laser gyro is heavier, uses (4.5m), which correspondsto 1.25° C-band and more power, and is more expensive than a con- 0.30° Ku-band beam sizes.The Ka-band reflec- ventional gyro. However, developments of this tor size is limited to 4 m, corresponding to a technology are expected to overcome these prob- 0.3° beam size,by the overallsatellitepointing lems by 1995. It is forecast that the ring laser accuracy. gyro will be used rather than the digital inte- grated rate assembly (DIRA) for sensing of satellite attitude. 2.2.2 Transponders and Receivers

Most significant advances in the transponder 2.2 Communications Payload subsystem will be in the area of better device performance - lighter and more efficient trav- Due to development of higher strength materials eling wave tube amplifiers (TWTA's) and solid and increased compactness of electronic compo- state power amplifiers (SSPA's) and better noise nents, a 15% mass reduction for the payload is figure low noise amplifiers (LNA's). projected for 1995. Table IV-1 gives TWTA efficiency (dc to rf power) predictions for 1985 and 1995 satellites. 2.2.1 Antennas The 50 W per transponder required at Ku-band Modest technical advances are projected over the will be supplied by a TWTA with an expected next decade for the antenna subsystem, but they 10 year lifetime for 1995. will be offset by the increased performance re- Solid state power amplifiers (SSPA) will be quirements imposed by closer orbital spacings. available with 10 W per device at. C-band and Increased component efficiency will be offset by 35% efficiency for 1995. SSPA have advantages the reduced antenna efficiency of tapered illumi- of increased reliability and lifetime, and much nation functions required to control sidelobes. less mass than the equivalent TWTA. A satellite Antenna subsystem development costs will be with 36 10 W C-band channels would require reduced by improved analysis programs that al- 1030 W of dc power for SSPA's versus 570 W for low skipping of the breadboard antenna design TWTA's. For 1995, SSPA's will be preferred for step. C-band use. Antenna manufacturing costs will be reduced LNA's using GaAS Fet's will be available for by near field range fa_ci!ities which allow faster 1995 with 2.5 dB noise figures, a 1 dB improve- and more accurate adjustment of the antenna ment from 1985. The most significant change for subsystem. receivers will be in a 50% mass reduction due to The development of higher strength materials large scale integration techniques. will allow a 10% reduction in antenna subsystem The development of dielectric filters and oscil- mass for 1995. However, increased frequency lators will allow a great reduction in transponder

IV-2 subsystem mass. Improvements in modulation techniques, particularly digital coding schemes, will allow more Type IYear I°°°l(%) P°wer'weig(W hr kg -1) t efficient use of the available bandwidth. NiCad 1985 55 21.3 MMIC technology has the potential to greatly NiH 1985 70 24.5 reduce payload mass and add capability, but is NiH 1989 70 31.0 judged to be immature for commercial satellite NaS 1992 70 58.0 applications in 1995. NaS 1995 70 70.0

2.2.3 On-Board Processing Table IV-2: Battery Comparison The most dramatic change in technology could be in the area of on-board signal processing due 2.3.2 Solar Cells to advances in VHSIC technology and high speed Current practiceuses silicon(Si)solarcellswith digital control systems. Benefits can be achieved 13.5% BOL efficiency.Developments are under- in the areas of: way to reduce cellthicknessand thus mass. Al- • Increased connectivity; though thincellsaremore expensiveto manufac- ture,the reduced mass willgive a lower overall • Increasedcapacity; costin geosynchronous orbit. The gallium arsenide(GaAs) cell,currently • Increased communications link efficiency; under development, has a 21% BOL efficiency and is relativelyimpervious to radiation,but is • Increasedflexibility. 2.5 times heavier than siliconand much more expensive. By 1995 GaAs may be equal to Si Increasedconnectivityviaswitchmatricesispos- cellsfor space applications(same on-orbitcost tulatedfor 1995 satellites,but on-board signal forequal capacity),but willonly be used when demodulation isjudged to be immature technol- area available for solar cells is limited (as for a ogy. high power spinner satellite). The 1995 solar collector technology for com- 2.3 Primary Power mercial satellites will remain silicon, but collec- 2.3.1 Batteries tor mass should be reduced by 25% for similar capacity systems. The standard method of power storage in commercial satellites has been the nickel cadmium 2.4 Propulsion Subsystem (NiCad) battery. In 1984, the Intelsat V and G- Star satellites became the first commercial satel- Two types of propulsionsystems are being used lites to use nickel hydrogen (NiH) batteries. The today: main advantage of the NiH over the NiCad bat- • Hydrazine stationkeepingsystem plussolid- tery is its higher depth of discharge (DOD) which propellantapogee motor; effectively increases its power to weight ratio. The sodium sulfur (NaS) battery is presently • A bipropellant system [nitrogen tetraoxide under development and promises to have a power (N204) and monomethylhydrazine (MMH)] to weight ratio three times the NiCad battery. used for both station keeping and apogee Although the NaS battery operates at a temper- motor firing. ature of 350 ° C, the required technology exists. Table IV-2 compares battery performance. The The hydrazine thruster has the advantage of NaS battery is the preferred technology for 1995 being able to supply smaller force-time incre- satellites. ments of thr,,_st_However, the biprop system

IV-3 results in mass savings due to the higher spe- takes place at C-band. However, a number of cific impulse of the fuel and is the preferred tech- new satellites are planning to use Ku-band since nology, allowing less fuel mass or longer station C-Band is becoming saturated with users. Like- keeping time with a given fuel mass. The biprop wise, as more satellites shift to Ku-band, there technology will be assumed for 1995 satellites. may need to be a shift to Ka-band with conse- Hydrazine thruster performance can be im- quent rain attenuation problems. For 1995, C proved by heating the fuel at the thruster. De- and Ku bands will be adequate. vices known as augmented catalytic thrusters impulse than today. However, significant elec- 2.7 Thermal Control tric power is required to operate these thrusters Present satellites use passive thermal control (approximately 10 kW for a 1 N thruster). Sys- plus heater augmentation. Passive radiators are tems requiring solid apogee motors win use this mounted on the north or south-facing panels of technology for station keeping, but bipropellant a 3-axis satellite or on the despun portion of a systems remain the preferred technology. spinner satellite. These systems are relatively in- expensive, but are heavy and limited in capacity 2.5 Structure and Mechanisms per radiator area. Business-as-usual satellite structures are primar- New generation satellites may incorporate ily constructed of aluminum or aluminum hon- heat pipes with passive or active pumping to re- eycomb materials with two main exceptions: duce mass and improve thermal dissipation ca- • If the satellite has a mass problem due to pacity. A single phase pump system using freon launch vehicle constraints, some structure fluid has been demonstrated which is efficient up may be manufactured from either graphite to 4 kW dissipation. A pumped heat pipe sys- fiber reinforced plastic (GFRP) or beryl- tem has more accurate temperature control than lium. a passive system, but this is not a critical factor for communications systems which can typically • Parts of the satellite critical to thermal dis- tolerate ±50 ° C. tortions, such as antenna related structures, The passive heat pipe rather than the pumped are usually constructed from GFRP due to system is the preferred technology for 1995 FSS its extremely low coefficient of thermal ex- systems. It has: pansion.

However, the additional expense of these exotic • Greater design maturity; materials will continue to keep their use to a minimum for commercial communications satellites. • Higher reliability and life; Higher strength graphite materials will be available in 1995, which should reduce mass by • Less complex integration. 10%. However, for the major part of the satellite, there will be no use of new structural technology. Heat pipes are imbedded within the honey- Satellite appendages are typically deployed comb structure of the equipment panel for best with one-shot spring motor devices or electrome- efficiency, minimum system weight, and less chanical actuators. No significant changes are complex spacecraft integration. expected for 1995o Radiators fixed to the body of the spacecraft rather than deployable radiators are preferred 2.6 TT_C Subsystem due to their lower weight and cost, and higher Little technology change is expected in the reliability. Deployable radiators would be used telemetry, tracking, and command (TT&C) sub- only if adequate area does not exist for fixed ra- system for 1995 satellites. At present TT&C diators.

IV - 4 2.8 Space Transportation Satellite The Ariane launch vehicle places the satellite Type Payload I 11 Task 31 4 in a highly elliptical orbit known as a geosyn- HS376 C-band chronous transfer orbit (GTO) with a perigee HS393 Ku-band altitude of 200 km and an apogee altitude of K2 Ku-band 36,000 kin. A high thrust apogee kick motor C&Ku (AKM) is then used to circularize the GTO or- Spacenet Fordsat C&Ku bit. Finally, fine orbit adjustments are made. The Space Transportation System (STS) or Shuttle places a payload in low earth orbit Table IV-3: FSS Systems (LEO) at an altitude of 260 kin. Further means are then required to transport the payload to geosynchronous earth orbit (GEO), which is a 5 along with the anticipated technical benefits. circular orbit of 36,000 km altitude. Table IV- Cost savings due to the improved technologies 4 gives transportation alternatives for transport will be incorporated into the Financial Model from LEO to GTO and from LEO to GE0. All of results of Section V. these alternatives are relatively costly (approxi- mate costs given in 1986 dollars). Another alternative method of transportation 3 Description of Three FSS from LEO (or GTO) is the integral perigee stage, Systems for 1995 which is controlled directly from the satellite. This results in lower mass and significantly re- This subsection presents descriptions based on duced launch costs. the preceding technology assessment of three It is anticipated that these means of space representative systems incorporating the fore- transportation will not change for 1995. casted improvements. Section V will use the A new development will be the ground based Financial Model to determine economic perfor- orbital transfer vehicle (OTV) which is sched- mance. uled to become operational in 1995. If the cur- Table IV-3 summarizes the satellite types used rent prices projected for its use are correct and for the different phases of the study. The Task 1 fuel costs do not become excessive, the STS Model Validation relates to 1985 satellites, while OTV combination will become the least expen- Tasks 2, 3 and 4 relate to 1995 satellites. The se- sive alternative. lected satellites types are named according to the The space-based (SB) OTV is scheduled to be- present type they most nearly resemble. How- come operational in 1998, and should have a dra- ever, they are generic types and payload and bus matic effect on upper stage launch costs. The do not necessarily match the existing or planned SB-OTV will be based at the Space Station and satellite of that name. will not have to be carried up from earth for each Table IV-6 compares the characteristics of the use. It will be reusable and return from GEO or- three 1995 satellites. Cost includes 12% G&A bit via aerobraking to conserve fuel. The capac- and 12% fee. ity of the SB-OTV is planned to be 12,000 kg, which will allow much larger satellites to be put into orbit. 3.1 Spinner Satellite System

2.9 Summary of Technology Devel- The spin stabilized satellite is similar to the opments Hughes HS-393 design. Table IV-7 presents a summary of its characteristics. Most prominent The anticipated technology developments for is the enhanced payload possible from the use of

IV-5 Capacity (kg) Cost Type Manufacturer Propulsion Leo-Gto Leo-Geo Status ($M) Pare D McDonnell Douglas Solid 1,250 0 Operational 8 Pain D II McDonnell Douglas Solid 1,842 Operational 9.5 TOS OSC Solid 5,942 In qualification 20 TOS/AMS OSC Solid/liquid 8,437 2,950 In development > 30 Delta Astrotech Liquid In development >_ 22 IUS Boeing Solid 2,268 Operational 30 Centaur G General Dynamics Cryogenic 11,022 4,536 Progam stopped 50 Centaur G I General Dynamics Cryogenic 11,160 5,988 Program stopped 50

Table IV-4: Available Orbital Transfer Rockets

Category Change Benefit Structure None Thermal Passive heat pipes Reduced mass of thermal subsystem. Higher thermal dissipation. Propulsion Bipropellant system Reduced fuel mass. Attitude Control Use of GPS & TDRSS More accurate and faster position determination. Ring laser gyro Increased reliability, less calibration time. Power Na£ batteries Improved power/weight ratio. Thinner Si solar cells Reduction in mass. GaAs solar cells Greater efficiency (21% vs 13%) TT&C None Comm. Payload Better design tools Reduced development time 8z cost.

Near field testing Reduced testing*:-^blAIl_ ° More efficient TWTA's Less power required. SSPA's at C-band Greater reliabilityand lifetime,lessmass Improved modulation More efficientuse of given bandwidth. VHSIC & microprocessors Bettercapacityforprocessingand switching. High strength materials 15% mass reductionfor antenna subsytem Large scale integration 15% mass reductionforelectroniccomponents Space Transport Ground-based OTV Reduced launch costs. Space-based OTV Greater launch capacity.

Table IV-5: Satellite Technology Developments (1995 Launch)

IV - 6 Spinner Ku-Band Hybrid Baseline satellite Hughes HS-393 RCA K2 Ford FS-1300 Design life (yr) 10 10 10 BOL mass (kg) 1377 1044 1540 Payload mass (kg) 261 261 342 - Antenna (kg) 29 29 52 - Transponder (kg) 232 232 290 EOL power (W) 2900 3000 4200 Stabilization Spin Three-axis Three-axis Frequencies Ku-band Ku-band C and Ku-bands Number of transponders: - C-band 24 - Ku-band 24 24 24& 6 Transponder bandwidth: - C-band (MHz) 36 - Ku-band (MHz) 54 54 36 & 72 Transponder power: - C-band (W) 10 - Ku-band (W) 50 5O 35 Antenna coverages: - C-band - Ku-band 3 3 Satellite EIRP (Conus): - C-band (dBW) 36 - Ku-band (dBW) 46 46 46 Launch vehicle(s): Ariane IV Ariane IV Ariane IV STS/PAM D II STS/PAM D II STS/Ford perigee Satellite Cost ($M, 1985) 54.2 50.9 64.6

Table IV-6: Summary of 1995 Satellite Characteristics

IV-7 Manufacturer & model: Hughes HS-393 Baselinesatellitename: Galaxy K EIRP (Conus): 46 dBW Lifetime: 10 yr On-board switching: Among coverage regions Launch vehicle: Ariane 4 or STS Frequency band and bandwidth: Ku-band, 500 MHz - receive: 11.7- 12.2 GHz - transmit: 14.0 - 14.5 GHz Antenna - type: Offset parabolic, dual gridded - number: 1 - size: 2.44 m - mass: 29 kg - feed array: 2 each 80 elements - coverage (3 beams): CONUS and E & W CONUS - polarization: Dual linear Transponders - number: 24 - power: 50 W - bandwidth: 54 MHz - TWTA redundancy: 5 for4 - receiver redundancy: 6for3 -- mass: 232 kg - dc power: 2,522 W Spacecraft - type: spin stabilized - size(stowed): dia = 3.64 m, length = 3.35 m - mass, BOL: 1377 kg - power, EOL at summer solstice:2900 W - primary power: Solar cells (Si) - batteries: NaS - attitudeand stationkeeping: BipropeUant thrusters - attitudepointingaccuracy: +0.05° - apogee motor: Liquid propellant

Table IV-7: Spin-Stabilized SateUite Characteristics (1995)

IV - 8 Gain Coverage (dBi) 29.0 East Half CONUS Full CONUS 35.2 West Half CONUS 29.7

Table IV-8: Ku-Band Predicted Antenna Gain

to any of the three coverage regions. The receiver preamplifiers are low noise GaAs Figure IV-l: Ku-Band FSS Coverage Regions FET's with 2.5 dB noise figure. Monolithic gain block amplifiers and dielectric oscillators will be used to reduce size and power consumption of 3.1.1 Ku-Band FSS Payload the receiver. The 11 GHz input multiplexers will This payload is similar to the Scenario If, Ku- use dielectric resonators to reduce size and mass band FSS concept developed under NASA Con- by a factor of three. TWTA efficiency of 53% tract No. NAS3-24235, Communication Plat- overall (TWT plus power supply) and 10 year form Payload Definition Study. It has three lifetime is projected. times frequency reuse with one CONUS beam, one eastern beam, and one western beam, all interconnected via electronic switches. 3.1.2 Other Design Features

The payload features a dual gridded reflec- The large payload power requirement of 2,500 W tor to obtain the polarization purity required for poses a problem for this spinner satellite design, reuse and two 80 element feed arrays. The re- in terms of area for solar cells and waste heat flector assembly is stacked with a slight offset in radiators. However, high efficiency GaAs solar focal point to give a physical separation of the cells could be used to allow 50% more power to feed assemblies. One feed assembly is for the be obtained from the same solar cell area, but at horizontal and the other for the vertical polar- the expense of added mass. ization. The feed array is connected to a diplexer which separates the receive and transmit channels on the basisof theirdifferentfrequencies. 3.2 Ku-Band Satellite System Figure IV-1 shows the antenna coverage regions;fullCONUS ishorizontalpolarizationand The 1995 Ku-band satellite is a 3-axis design similar to the RCA K-2. Its characterisitics are the east and west beams are verticalpolariza- shown in Table IV-9. tion. The spacialseparationbetween east and west beams is necessary to achieve 25 dB co- polarizationisolation. Table IV-8 gives pre- 3.2.1 Ku-Band FSS Payload dictededge-of-coverage(EOC) antenna gains. Each coverageregionhas eight54 MHz band- This isthe same payload as describedin subsec- width transponders,giving a totalof 24 chan- tion3.1.1forthe Spinner satellite. nels.The receiveroutput foreach coverage region is fed to an input multiplexerwhich sub- 3.2.2 Other Design Features dividesthe 500 MHz IF band into eightchan- nels.The threeinput multiplexersaxe followed The 3-axis satellite design allows for additional by eight3 x 3 switch matricesin orderto permit solar cell area. Use of thin Si solar cells will lead signalsin any uplinkchannel to be retransmitted to overall mass reduction.

IV-9 Manufacturer & model: RCA Americom, K2 Baselinesatellitename: Satcom K2 EIRP (Conus): 46 dBW Lifetime: i0 yr On-board switching: Among coverage regions Launch vehicle: Ariane 4 or STS/PAM D2 Frequency band and bandwidth: Ku-band, 500 MHz - receive: 14.0-14.5 GHz - transmit: 11.7-12.2 GHz Antenna - type: Offset parabolic, dual gridded - number: 1 - size: 2.44 m -- mass: 29 kg - feed array: 2 each 80 elements - coverage (3 beams): CONUS and E & W CONUS - polarization: H and V, linear Transponders - number: 24 - power: 50 W - bandwidth: 54 MHz - TWTA redundancy: 5 for4 - receiver redundancy: 6for3 -- mass: 232 kg - dc power: 2,522 W Spacecraft - type: 3-axis stabilized - size(bus): 1.57 x 2.18 x 1.77 m - mass, BOL: 1200 kg - power (EOL) at summer solstice: 3000 W - primary power: Solar cells (thin Si) - batteries: 4 NaS, 150 Ah - thermal control: Heat pipes - attitude and station keeping: Hydrazine thrusters (ACTS) - attitude pointing accuracy: 4-0.07° - apogee motor: Solidpropellant

Table IV-9: Ku-Band Satellite Characteristics (1995)

IV - 10 Manufacturer _ model: Ford Aerospace FS-1300 Baseline satellite name: Ford Hybrid Satellite EIRP (Conus): 36 dBW C-band, 44 dBW Ku-band Lifetime: 10 yr On-board switching: Among coverage regions, also C- and Ku-bands interconnected Launch vehicle: Ariane 4 or STS/Ford perigee Frequency band and bandwidth: C-band 500 MHz, Ku-band 500 MHz - receive: 5.925-6.425 and 14.0-14.5 GHz - transmit: 3.700-4.200 and 11.7-12.2 GHz Antenna - type: Offset parabolic, dual-gridded - number: 2 - size: 1.4 m x 1.8 m C-band, 2.44 m Ku-band -- mass: 17 kg C-band, 35 kg Ku-band - coverage (2 C and 3 Ku beams): CONUS and E & W CONUS - polarization: H and V linear for both bands Transponders - number of C-band: 24 - power at C-band: 10 W - bandwidth at C-band: 36 MHz - SSPA redundency (C-band): 5for 4 - receiver redundency (C-band): 4for 2 - number of Ku-band: 30 - power at Ku-band: 33 W - bandwidth at Ku-band: 36 MHz (24), 72 MHz (6) - TWTA redundency (Ku-band): 5for4 - receiver redundency (Ku-band): 4for 2 -- mass: 72 kg C-band, 290 kg Ku-band - dc power: 660 W C-band, 2,780 W Ku-band Spacecraft - size(stowed): 2.5 m x 1.88 m x 2.64 m - mass, BOL: 1540 kg - power (EOL) at summer solstice:4000 W - primary power: Solar cells (thin Si) - batteries: 4 NaS, 232 Ah (total) - attitudeand stationkeeping: 3-axis stab, biprop thrusters - attitudepointingaccuracy: ±0.I° - apogee motor Liquidpropulsion

Table IV-10: Hybrid Satellite Characteristics (1995)

IV - ii 3.3.2 Ku-Band FSS Payload

Function Coverage Region This is the same payload as described in subsection 3.1.1 for the Spinner satellite, except Transmit Conus 27.0 V that transponder bandwidth is now 36 MHz and Transmit Alaska 27.0 V thus there are a total of 36 transponders (12 per Transmit Conus + Alaska 25.0 V 500 MHz band coverage region). Receive Conus + Alaska 25.0 V Transmit Conus 27.2 H

elbl 0 TT 3.3.3 Other Design Features .L _r _IAIt$1LU b n_w_IAi O.L.O ££ Transmit Puerto Rico 29.7 H The use of 36 MHz bandwidth at C and Ku- Transmit Conus + Hawaii 25.1 H bands isto allowinterconnectionbetween bands, Transmit Conus + PR 25.1 H i.e.uplinkat Ku-band and downlink at C-band. Receive Conus+HI+PR 25.1 H There isalsoprovisionforswitchingof transponders among the differentcoverageregions.

Table IV-11: C-Band Antenna Predicted Gain 3.4 Comparison of Systems

3.3 Hybrid Satellite System The spinner and Ku-band systems carry the same payload, but the spinner design requires The 1995 hybrid satellite design is a 3-axis satel- 15% greater mass. The hybrid system is larger lite similar to the Ford Satellite design. Its char- yet with 54 transponders versus the 24 transpon- acteristics axe summaxized in Table IV-10. The ders of the other two systems. payload includes both C and Ku-band transponders, which are interconnectable and therefore of 4 Large Satellite System the same bandwidth (36 MHz).

A 2,200 kg (5000 lb) satellite design is postu- 3.3.1 C-Band FSS Payload lated which takes advantage of the space-based orbital transfer vehicle (SB-OTV). This satellite This payload is similar to the Scenario II C- is named "Hectosat" after its 100 transponder band FSS concept developed under NASA Con- payload. tract No. NAS3-24235, Communication Plat- The satellite is at the upper mass limit of the form Payload Definition Study. It provides two study. Satellites of larger than 5,000 lb are con- times frequency reuse via horizontal and vertical sidered platforms and are outside the scope of polarized beams. The coverage is full CONUS this study. with selected coverage for Alaska, Hawaii, and Puerto PJco. The motivation for this design is the SB-OTV which can more efficiently transfer mass from The C-band antenna consistsofa dual-gridded LEO to GEO. The SB-OTV will be based in offset-fedreflectorand two 7-elementfeedarrays LEO at the Space Station and is reusable. Its foreach polarization.Both transmitand receive mass doesn't need to be lifted from earth to LEO bands use the same feed array.Antenna sizeis for each launch. It uses atmospheric aerobraking 1.4m x 1.8 m. Table IV-11 givesthe calculated to reduce fuel costs for return from GEO to the edge-of-coveragegainsforthe verticaland hori- Space Station. zontallypolarizedantennas. The C-band transponder uses SSPA's with 4.1 Satellite Design 35% dc to RF power efficiency.The receiverhas low noiseGaAs FET's with 1.5dB noisefigure, The satellite is a three axis design of 2,144 kg monolithicgainblocks,and dielectricoscillators. mass and 3,100 W primary power. It has a pay- The multiplexerswilluse dielectricresonatorfil- load of 108 Ku-band transponders. Table IV-13 ters. gives the characteristics of the design.

IV- 12 -- f ll+ /zv - I Gain I Power I (dBW)_(W) _

.. 6.6 %.>, \,-+" : ; __"T,._=_-_'C_ 6.4 5.6 -0.9 6.8 13.1 10.7 10.0 7.2 Figure IV-2: Ku-Band FSS Coverage Regions

Table IV-12: Power Requirements for Beams 4.2 Payload

Hectosat uses the Scenario V Ku-band FSS tor input to the Model is 77%. (Note that concept developed under NASA Contract No. transponders are not sold as entities in this sys- NAS3-24235, Communication Platform Payload tem. Rather parts of nine transponders are used Definition Study. It provides nine times fre- by each trunking location, depending on volume quency reuse via horizontal and vertical polar- of traffic and destination.) ized regional beams over CONUS (see Figure IV- 2). Its capacity is three times the Ku-band pay- 4.3 Launch Vehicle load on the spinner and Ku-band satellite designs. Transponder bandwidth is 36 MHz versus The 2,144 kg BOL mass would require the IUS or Centaur upper stage (see Table IV-4) to be used the 54 MHz of the previously used Scenario II with a Shuttle launch for a business-as-usual sce- Ku-band concept. nario (no Space Station). The economics of a The payload features a 4.6 m (15 ft) dual- space-based OTV should be much better, and gridded reflector assembly for the spacial reso- will be analyzed in Section VII. lution and polarization purity required for frequency reuse. Table IV-12 shows the projected antenna gain and amplifier power levels necessary to maintain a 46 dBW EIRP with single carrier operation. As described in the Communication Platform Payload Definition Study, this payload addresses voice trunking traffic and data trunking traffic. Single side band (SSB) modulation techniques allow a total voice traffic capacity of 519,000 half voice circuits. Satellite switched time division multiple access (SS-TDMA) is used on 12 transponders to give a data trunking traffic capacity of 519 Mb/s. The coverage and interconnectivity of this system has been designed to match the traffic mix forecasted for the future by region. Traffic bal- ancing is such that the fill factor for the 108 transponders is 0.77. Thus the ntiliza.tion_ fac-

IV - 13 Baseline satellite name: Hectosat r.,, r_r _<,ouus): 4D (ID VY Lifetime: 10 yr On-board swit ching/interconnectivity: Among coverage regions Launch vehicle: STS and Centaur or OTV upper stage Frequency band and bandwidth: Ku-band; 500 MHz - receive: 14.0-14.5 GHz - transmit: 11.7-12.2 GHz Antenna - type: Offset parabolic, dual-gridded - number: 1 - size: 4.6 m -- mass: 161 kg - coverage (9 beams): 9 regional beams - polarization: H and V linear for both bands Transponders - number: 108 - power: 20 W maximum; 8 W average - bandwidth: 36 MHz - SSPA redundency: 5 for4 - receiver redundency: 4 for 2 -- mass: 586 kg -dc power: 2,700 W Spacecraft - size (stowed): 2.6 m x 1.9 m x 2.8 m - mass, BOL: 2,144 kg - power (EOL) at summer solstice: 3,100 W - primary power: Solar cells (thin Si) - batteries: 4 NaS, 45 Ah - attitude and station keeping: 3-axis stab, hydrazine thrusters - attitude pointing accuracy: +0.05 ° - apogee motor None

Table IV-13: Large Satellite Characteristics (1995)

IV - 14 Section V

BASELINE ECONOMIC PERFORMANCE

1 Introduction A 50% increase in number of transponders on the same size satellite is pos- " This section presents the Task 2 results on the sible. This allows a 50% increase in determination of the economic performance of communications capacity and revenues the business-as-usual scenario for IOC 1995, as (1985 Ku-band to 1995 Ku-band 3-axis defined in Section IV. The Financial Model redesign). sults for the three FSS systems forms the baseline for comparison with the Space Station APO Use the Financial Model to calculate the Scenarios of Tasks 3 and 4. 1995 satellite system initial rate-of-return. In addition, a large (2,200 kg) high-capacity The DTRR (dual terminal rate-of-return) satellite design which takes advantage of the averaged 4.4% higher for the 1995 than the launch capacity of the space-based OTV is an- 1985 returns as shown in Table V-1. alyzed. The large satellite is called "Hectosat" Adjust the 1995 transponder price until the after its 100 transponder payload. average 1995 return equals the average 1985 The economic performance for the following return. This requires iteration of the Finan- four 1995 satellite types is presented: cial Model.

• Ku-band spin-stabilized satellite; The logic for this step is that market forces will eventually push the return • Ku-band 3-axis satellite; down to the original level.

• Hybrid (C and Ku-bands) 3-axis satellite; This results in a 33% decrease in the basic transponder price from $1.9 M to • Large Ku-band 3-axis satellite. $1.27 M. This is equivalent to a 4.1% per year transponder price reduction The Financial Model results for the 1985 satel- for 10 years. lite designs are given in Subsection II-5 and Ap- • The result is the baseline 1995 satellite eco- pendix A. nomic performance as shown in the "Final" column of Table V-1 and described in the 2 Methodology following subsections.

The following methodology is used to obtain the Table V-1 gives the dual terminal rate-of- 1995 baseline satellite economic performance: return (DTRR) for the four satellite types that are analyzed. The 1985 column gives the Fi- • Start with 1985 satellite designs; nancial Model results for the 1985 IOC satellites with basic transponder price (C-band, 36 MHz) • Predict 1995 technology; of $1.9 M per year. The "initial" 1995 returns are for the 1995 • Evolve 1985 satellites to 1995; satellite designs (50% more capacity) and the

V-1 output is calibrated. The satellite costs for the DTRR Return, % Model are derived by altering the parameters Satellite 1985 L______5 to reflect a HS-393 design utilizing 1995 tech- Design [ Initial [ Final nology. The other capital costs are developed C/Ku Spinner 18.1 23.4 18.9 from current Ford Aerospace experience or pub- Ku 3-axis 19.8 23.3 18.8 lished prices. Since the component capital re- Hybrid 3-axis 21.9 26.5 21.9 qnirements axe validated, the Model itself is also Large 3-axis - 29.6 25.1 considered to be validated.

Table V-l: DTRR forSatelliteSystems 4 Ku-Band 3-Axis System

Satellite Cost ($M 1985) The nine pages of Table B-2 in Appendix B Design 1985 I 1995 give the Model results for the Ku-band 3-axis satellite. The DTRR is 18.8% (after the 33% C/Ku Spinner 76.5 115.1 Ku 3-axis 104.3 116.8 transponder price reduction) with a total capital expenditure of $116.8 M. Hybrid 3-axis 83.1 138.8 The costs are based on the Price H output as Large 3-axis - 215.4 a result of altering the 1985 K2 parameters to reflect the predicted 1995 technology. Table V-2: Capital Costs for Satellites

5 Hybrid 3-Axis System same basic transponder price. The "final" 1995 returns were adjusted 4.4 points lower so that the The nine pages of Table B-3 in Appendix B give average return equals the average 1985 return. the Model results for the hybrid 3-axis satellite. This required a 33% decrease in basic transpon- The DTRR is 21.9% (after the 33% transponder der price. price reduction) with a total capital expenditure The Large satellite is a 1995 design. Its "ini- of $138.8 M. tial" and "final" returns axe 29.6% and 25.1% The costs axe derived from the Ford Satellite respectively. The higher return implies that program, but are somewhat generalized. By first transponder prices will be further reduced. calibrating against the currently known satellite Table V-2 gives the capital costs of the base- and then altering the parameters to reflect the line satellites. The greater costs of the 1995 1995 technology, the Price H satellite costs are satellites are due to the increased number and considered validated. power of the transponders.

6 Large Satellite System 3 Ku-band Spinner System The nine pages of Table B-4 in Appendix B The nine pages of Table B-1 in Appendix B give give the Model results for the large satellite with the Model results for the Ku-band spin-stabilized Shuttle launch and Centaur upper stage. The satellite. The DTRR is 18.9% (after the 33% DTRR is 29.6% after the 33% transponder price transponder price reduction) with a total capital reduction and application of the 0.77 utilization _-_._t,,_o of $115 I M. factor as discussed 1£1:_ _.m)c.._._^_:^_U.U_I._blU.L/. _nr¥ -'-_./._._ ,_ __J. JLJLt_ The satellite bus costs are derived from the total capital expenditure is $215.4 M. CS-2 program at Ford Aerospace. The Ku-band A target DTRR return of 23% is used for the payload costs are derived from another program. large satellite. The one point premium over the By varying the parameters against these costs hybrid system is judged to be a necessary incen- and converging on the known costs, the Price H tive for such a large system with its concentra-

V-2 tion of risk. To reach this target return, a further 18% reduction in transponder price (from $1.27 M to $1.04 M) would be required. The implication is that the large satellite system can either be more profitable than the three business-as-usual cases or that transponder prices can be reduced. Market forces will cause more large satellites to be built with their better economic performance, and transponder prices will eventually be reduced.

7 Discussion

There islittleto choosebetween the capitalcosts and ratesof return for the spinner and 3-axis Ku-band systems. However, due to itsgreater number of transponders,the hybrid system has a 3% greater rate of return. This is achieved without selling any cross- connected transponders; i.e. transponder prices are based on all C and all Ku-band sales. As discussed in Subsection IIL4.8, sales of hybrid pairs of transponders bring a 30% premium and would further increase the return. For the purposes of this analysis, we take the conservative assumption that revenues from sales of hybrid pairswillbe offsetby a decreasein utilizationof the remaining "wrong way" pairs. The impressiveresultsfor the large satellite are due to economies of scale.The implication is clearlythat thisis the satellitedesign of the future.A 18% transponderpricereductionfrom the best performing 1995 satelliteis achieved, and a 45% pricereductionfrom the 1985 satellite systems.

V-3 Section VI

NEW SPACE-BASED ACTIVITIES

1 Introduction equipment, and internal and external vehicular activity (IVA/EVA) skill types. This section presents the Task 3A results in pos- These APOs can be combined for advanced tulating and defining new or enhanced space- scenarios, and in some cases may require another based Activities, Procedures, and Operations APO to become possible. (APOs) and associated satellite system designs Subsection 2 defines eleven APOs, and Sub- that have the potential to achieve future com- section 3 discusses the impact on the design of munications satellite operations in geostationary the satellite. orbit with improved economic performance. Selection has been made of the most promis- ing space-based APOs that have the potential 2 Postulation of APOs to achieve future communications satellite operations in geosynchronous orbit with improved The criteria for selection of the APOs are in- economic performance compared to the business- creased communications satellite technical and as-usual scenario for 1995. economic performance. The selection of APOs For each APO, the functional description is is made based on predicted available technology followed by three paragraphs: and judgment of economic value. The eleven APOs described in subsections 2.1 1. Scenario describing sequence of operations to 2.11 are as follows: with the APO; 1. Emergency retrieval from LEO 2. Requirements placed on the communica- 2. Ground-based Orbital Transfer Vehicle tions satellite, including changes in satellite system design to accommodate the new (OTV) launch to geostationary transfer or- APOs; bit (GTO)

3. Requirements placed on the Space Station 3. Ground-based OTV launch to geosynchro- and its supporting equipment: nous earth orbit (GEO)

- Functional and technical requirementS _' 4. Deployment of appendages at shuttle

- Support equipment includes the Or- 5. Space-based OTV launch to GTO bital Maneuvering Vehicle (OMV), Or- bital Transfer Vehicle (OTV), Remote 6. Space-based OTV launch to GEO Manipulator System (RMS) on Shut- 7. Deployment of appendages at Station tle, and the Mobile RMS (MRMS) on the Space Station. 8. Checkout at Space Station The types of requirements include characteris- 9. Fueling at Space Station tics and capabilities, facilities, interfaces, operational timelines and schedule constraints, special 10. Assembly at Space _qtation

VI- 1 11. Servicing/replacement for GEO satellites satellite could not be designed for every con- tingency, certain safety devices could be im- - Transport to low earth orbit (LEO) for plemented to facilitate repair at the Shuttle or servicing Space Station and relaunch without return to - Servicing in GEO Earth.

The APOs are listed in order from simplest to 2.1.1 Scenarios most complex, which is approximately the same as chronological for availability. There are three scenarios for emergency retrieval APOs 2 to 4 use the Ground-Based (GB) OTV corresponding to repair at the shuttle, at the and thus would be performed directly from the Space Station, and on an OTV launch. Shuttle without need for the Space Station. The Retrieval by Shuttle only difference between APOs 2 and 3 is the orbit achieved by the OTV: APO 3 reaches full GEO 1. Satellite is deployed from Shuttle bay and orbit, while APO 2 requires the spacecraft have fails to activate. an apogee stage. APOs 5 to 7 are the space-based OTV ver- 2. Shuttle approaches within safe distance of sions of APOs 2 to 4 and would involve use of satellite. the Space Station. APO 8 is enhanced checkout 3. MMU (manned maneuvering unit) or of the satellite, such as antenna pattern testing, RMS/EVA activity to dock satellite. and is to be differentiated from health checks which would be performed at some level during - Attach docking device. each APO. - Despin satellite if required.

2.1 Emergency Retrieval 4. Satellite is brought back to Shuttle and grappled with RMS. Emergency retrieval is the most realizable AP0 and is being used today in limited form from the 5. Satellite may be repaired via extra-vehicular Shuttle. Insurance underwriters can be expected activity (EVA) and ground-originated in- to require emergency retrieval provisions a stan- structions. dard feature on communications satellites. At present, the faulty satellite usually must wait un- . If repair is satisfactory, relaunch satellite. til a future Shuttle flight for repair or retrieval, 7. If repair is not satisfactory, use safety fea- and often suffers thermal injury. Short-notice tures designed on satellite to make it comply availability of a space-based Orbital Maneuver- with Shuttle safety regulations. vemem tviv,¥ ) that can dock with a generic satellite would be required. 8. Return satellite to Earth via Shuttle. If a satellite is launched with the OTV, the OTV must wait until the satellite is deployed 9. Repair satellite and relaunch. and operational. If a major failure occurs, the satellite could be retrieved and returned for re- Retrieval at Space Station pair. However, this option requires that additional mass be carried in the form of extra OTV 1. Satellite is deployed from Shuttle bay and fails to activate. grappling fixtures on the multiple payload car- 2. OMV is released from Space Station (SS) rier. and rendezvous with failed satellite. The non-functional satellite being returned from GEO must be able to withstand the OTV 3. OMV grapples satellite with remote arm or atmospheric aerobraking forces. Although the RMS.

VI - 2 4. Satellite is brought back to SS. Slow OMV Safety features to remotely inhibit all satel- spin for thermal control. lite propulsion and pyrotechnic devices.

5. Rendezvous with SS, use MRMS to take Tank purging capability through controlled OMV and satelliteto storage. firing of thrusters or advanced methods.

6. Undock OMV and satellite,storesatellite Accurate, detailed documentation of satel- untilservicing. lite available for emergency use.

7. Place OMV back in itsstoragehanger. Grappling fixture in accessible location if OMV retrieval is required. 8. Repair satellitevia EVA with ground link to satelliteexperts. 2.1.3 Requirements on Shuttle, Space Station and OTV 9. If repair not possible, make safety precautions and return to ground on next available Requirements on Shuttle Shuttle. • Retrieval equipment flown on launch. I0. If repairis successful,use OMV to place satellitein new orbit away from SS, re- • Crew prepared for emergency retrieval. launch. Requirements on Space Station Retrieval on OTV Launch • OMV available at Space Station for emergency use. 1. Satelliteis deployed from OTV per OTV APO (2,3,5or 6). • OMV grappling capability.

2. OTV deploys other satellitesbut does not • Dedicated servicing equipment at Space de-orbit. Station (servicing bay).

3. Satellitebegins deployments while being • Storage facilities at SS to provide thermal monitored by ground facilities. control.

4. Satellitecontinuesuntilon-orbitcapability • Standard power supply for battery charging. isachieved. Requirements on OTV 5. If satellitefailsduring deployment, contin- genciesare attempted or option to retrieve • Orbit holding capability. with OTV is chosen. • Remote docking capability. 6. OTV closesin on satellitein low thrust mode. • Additional fuel allocated for return of a satellite. 7. Redocking of satellite by OTV. 2.2 Ground-Based OTV Launch to 8. OTV returns to LEO and servicing occurs Geostationary Transfer Orbit with Space Station or Shuttle. The Ground-Based Orbital Transfer Vehicle (GB-OTV) can substitute for the business-as- 2.1.2 Requirements on Satellite usual perigee or integral stages used to place • Standard configuration for attachment of most communications satellites into an elliptical RMS grappling fixture.

VI - 3 little change in satellite design and minimize im- 2.2.2 Requirements on Satellite plementation costs. Commercial perigee stages or alternate launch vehicles such as the Ariane • OTV and back-up launch system attach- would he major cost competitors, but could also ment compatibility. serve as reliable back-up systems. • Standard hardpoints for satellite handling The OTV should simultaneously launch sev- during OTV connection. eral satellites in order to be cost effective. This requires a reusable Multiple Payload Carrier • Thermal requirements/shroud design for (_4"D/'_ ...T._k .... 1.1 k^ _ _I^ + ..... I_I.^ _+ .... _AVAA _), vv _.J._..AL _.,vu, Lv,J. u,_ ck OIAAA_A_ _A ULOO-XAZ'a_ O_A rAp- t)'l'v launch. ture with variable lengths to connect the pay- loads. • Automatic activation of satellite systems Note that thisAPO isa launch to geostation- upon detachment from OTV. ary transferorbitonly.This would be more ben- eficialfora satellitewhose launch cost isdeter- • Thermal and power data telemetered to mined by length,and thereforeenablingthe fuel ground. mass requiredfor the apogee maneuver to be transported without cost. (An example would 2.2.3 Requirements on Shuttle and OTV be a spinner satellite design with surface area and hence volume determined by power require- Multiple payload carrier, capable of sup- ments.) porting several satellites of different masses for the same launch without loss of effi- 2.2.1 Scenario ciency.

1. Satellites and GB-OTV in Shuttle bay, Sufficient space on Shuttle with GB-OTV launch to LEO orbit. for efficiently packaged satellites (mass vs. length) that optimize OTV use. 2. Deploy GB-OTV plusMPC, (assembleifre- quired). • Slow spin capability for OTV to allow satellite thermal control. 3. Affixpayloadsto MPC in predefinedconfig- uration. Data communications contact through MPC to satellites if telemetered data can- 4. Attach deployable thermal shrouds if re- not be sent by several satellites on OTV. qnired.

2.3 Ground-Based OTV Launch to 5. Release OTV, launch to geosynchronous transferorbit. Geosynchronous Orbit

The ground-based orbital transfer vehicle can 6. Ifspinningof OTV isrequired,must despin in GEO transferorbit. be used to place satellites into circular geosynchronous orbits. For many types of satellite de- 7. Deploy satellitefrom MPC. signs, the elimination of the large amount of fuel and propulsion system required to perform the

. Turn on S/C systems,begin pre-orbittrans- apogee maneuver can lead to a simpler and more 4"...... _:^--. compact design. This, however, _._._...... ;..... jo.;_ redesign of the satellite bus and additional imple- 9. Confirm 0TV - S/C distance, begin con- mentation cost. Satellite designs that are driven trolled apogee burn to GEO. by other factors, such as the surface area available for solar arrays on a spinner, may not ben- 10. Begin satellite on-orbit operation. efit from this APO.

VI - 4 No commercial system yet exists that can satellitedesign and launching. New designsfor launch an unintegrated satellite into geosynchro- appendages could be implemented that are not nous orbit in a low thrust mode. Several types of dependent upon automatic deployment mecha- systems have been studied but none have been nisms. However, the deployed satellitewould implemented to date. It is imperative to have have to be able to withstand the 0.1 G forceof such systems developed and tested before a satel- the OTV or otherlow thrustlaunch vehicle.Re- lite design would incorporate the use of the OTV liabilitywillbe increasedthrough simplification for this APO. of automatic procedures and testingof the on- orbitconfiguration.Some cost savings can be 2.3.1 . Scenario gained through the elimination of zero-gravity simulations of complicated deployment schemes. Same as scenario in paragraph 2.2.1 for GB-OTV launch to GEO transfer orbit, except that OTV 2.4.1 Scenario releases and activates satellite in circular GEO orbit. I. Deploy GB-OTV from Shuttle.

2.3.2 Requirements on Satellite 2. Remove satellitefrom cradlewith RMS.

• Same as OTV launch to GEO transfer orbit 3. Affixsatelliteto raisedwork stationor lock (paragraph 2.2.2), plus arm to hold satellitein usable orientation within EVA reach. • No apogee motor needed: large decrease in fuel results in smaller fuel tanks and possi- 4. EVA assisteddeployment of antennas,solar ble change to less efficient fuel requiring less arrays,other appendages. hardware 5. Test on-orbitconfigurationas per checkout Possibleredesignof satellitebus to be more APO (subsection2.8). space efficient 6. Launch satellitewith low thrustlaunch ve- Fewer controland deployment modes: one hiclesuch as OTV (2.3). setof on-orbitdeployment modes, no trans- ferorbitmodes needed. 2.4.2 Requirements on Satellite • Operates with OTV or commercial external transfersystem. • Standard hardpoint for RMS handling.

• Separate connection interface for worksta- 2.3.3 Requirements on Shuttle and OTV tion. • Same requirements as OTV launch to GEO transfer orbit (paragraph 2.2.3). • Devices for deployment of appendages cr methods of safe deployment via EVA. • The Shuttle must maintain its orbit for a longer time until the GB-OTV can return • Deployed appendages must withstand low from the full GEO launch. thrust (0.1 G) of OTV.

• For operation before 1997, a back-up mode 2.4 Deployment of Appendages at required for deployments. Shuttle

The externally-assisted deployment of ap- • Satellite appendages are limited in size by pendages in low earth orbit would create a ma- the MPC and other satellites being simulta- •"^. _1-._ " 41.^ k.. ,I"...... 1 .... 1. ju_ _L_LE, e in _L_ uuS_u,_o-o_-u_uo_ apt_v_Cn to neously launched.

VI - 5 2.4.3 Requirements on Shuttle and OTV • Automatic activation of satellite systems upon detachment from OTV. • Trained satellite handlers on Shuttle flight.

• Low-thrust mode (0.1 G) for OTV. 2.5.3 Requirements on Space Station • Raised workstation accessible for EVA. • Efficient scheduling of OTV with Shuttle to avoid satellite storage. 2.5 Space-Based OTV Launch to 4- 4...... • -- CI^C .... .'-I.A .... _-- _t__z_11!L__ t .... fit .-zzl - ,.)O.A_, _.ltt.lLlk blO, IL_l,C.t UI _¢i,b_JA1bt_ IIUILL ,DIi.Ut, 51_ bay to OTV. The use of the Space-Based (SB) OTV which is based at the Space Station has the advantage • Quick connect docking procedure. of not having to carry the GB-OTV up in the Shuttle for each launch. (The GB-OTV will not Multiple payload carrier, capable of sup- be left in LEO between missions.) porting several satellites of different masses for the same launch without loss of effi- 2.5.1 Scenario ciency.

o Satellites are launched via Shuttle direct to • Slow spin capability for OTV to allow satel- Space Station. (Optional use of OMV to lite thermal control. move satellites from nominal Shuttle orbit to Space Station, or unmanned cargo transport to Space Station via expendable launch 2.6 Space-Based OTV Launch to Ge- vehicle (ELV). osynchronous Orbit

2. RMS and MRMS used to transfersatellites The Space Station and SB-OTV can be used for to SB-OTV. launch to full geosynchronous orbit. The scenario would be the same as paragraph 2.5.1, ex- 3. Attach satellite to MPC on SB-OTV, ready cept that the launch is into full geosynchronous to launch. orbit. The requirements on the satellite are the same as paragraph 2.5.2, and the requirements 4. Dock OMV to 0TV. on the Space Station are the same as paragraph 2.5.3. 5. Use OMV to move 0TV away from Space Station. More details of OTV performance is given in Subsection VII-2.3. 6. Launch OTV to GEO transfer orbit.

7. Release each satelliteand launch to full 2.7 Deployment of Appendages at GEO orbit. Space Station

8. Return OTV to Space Station. The deployment of appendages at the Space Sta- tion and the subsequent testing before OTV 2.5.2 Requirements on Satellite launch has advantages over a Shuttle based operation (Subsection 2.4) where operation time is • OTV and back-up launch system attach- limited. More elaborate deployments can be per- ment compatibility. formed at the Space Station, and concentrated • Hardpoints for satellite handling during testing of the on-orbit configuration could add OTV connection. significant value to the operational life of the satellite. This APO requires that additional • Thermal requirements/shroud design for equipment and a dedicated servicing and testing OTV launch. bay for satellites be placed on the Space Station.

VI - 6 2.7.1 Scenario 2.8 Checkout at Space Station

1. Satellites in Shuttle bay transported di- Health checks via a standard data port are made rectly to Space Station. on the satellite before launch and deployment. It may be desirable to have a more thorough check- 2. RMS and MRMS used to transfersatellites out of satellite systems, especially if the satellite to Space Station storage bay and servic- can be deployed into its on-orbit configuration ing/deployment area. at the Space Station (or Shuttle). New tests could include power levels, accurate 3. Positionsatellitein servingarea. center of mass properties, and detailed communication systems tests. Basic tests such as power 4. EVA assisteddeployment of appendages. levels could be conducted from the Shuttle, but 5. Test on-orbitconfiguration(Checkout APO, more complicated tests such as antenna patterns subsection2.8). that could enhance the reliability of full on-orbit operation would require extra equipment more 6. SB-OTV launch scenario(paragraph 2.5.1) appropriately placed on the Space Station. with low (0.1G) thrust. Requirements for this AP0 may be driven by the insurance underwriters and trends in satellite failures. The checkout AP0 will become more 2.7.2 Requirements on Satellite cost effective when combined with other AP0s. • Deployment requirements same as at Shut- tle (paragraph 2.4.2). 2.8.1 Scenario

• More advanced testing capability required. 1. Satellite is brought to Station and stored.

. Routine health checks are carried out while 2.7.3 Requirements on Space Station at Space Station. (Other APOs would also • Dedicated servicing/deployment area with incorporate routine health checks.) adequate area to deploy booms, arrays, etc. without obstructing other operations• . After launch configuration is completed (possible exception of fueling), satellite is • This may require external facilities, electri- placed in servicing bay or dedicated testing cal power, and communication systems if facility. there is not enough space in the service bay. 4. Tests are performed. If anomalies are found, • Storage facilitiesprovidingthermal control attempt repair and/or transport to ground. forundeployed satellitesor satellitesrequir-

ing active thermal control prior to OTV , Far field antenna testing may require use of launch. an OMV to transport satellite to adequate test range and control pointing during tests. • Battery charging facilities for satellites. 6. Antenna testequipment at SS isused to test • Efficient scheduling of OTV-Shuttle to patterns. avoid excessive satellite storage. 7. Satellite on OMV is returned to Space Sta- • Trained EVA satellite handlers on Space tion. Station crew. 8. Launch satellite on OTV or alternate launch • Low thrust (0.1 G), efficient OTV. system.

VI - 7 2.8.2 Requirements on Satellite when used to launch satellites into geosynchronous orbit, will decrease the advantage of fuel- • Standard hardpoints to allow satellite han- ing. Systems designed to use liquid fuels for inte- dung. grated launch stages wiU see increased advantage of low cost Space Station fueling. • Equipment to allow communications with systems being tested. 2.9.1 Scenario

• Added abilities to facilitate testing such 1. Satellites in Shuttle, launch to Space Sta- on/off control of single transponder. tion.

2. RMS and MRMS used to move satellitesto • Standard data port. storagefacilityor fuelingdepot.

2.8.3 Requirements on Space Station 3. Affixsatelliteto fuelingports.

Servicing/testing bay facility with power 4. Fuel satellitewith propellantand pressur- supply, data communications port, thermal ant. control. 5. Disconnect satellite.

• Storage facility with thermal control. 6. Transfer to OTV if self-propeUed or to OMV for maneuvering away from SS. • Standard testing equipment and small mobile equipment for servicing. Requirements on Satellite

• Antenna testing system with power level Standard hardpoints to allow satellite han- and pointing capability. dling.

Standard quick disconnect fueling ports. 2.9 Fueling at Space Station Fueling meters for internal pressure, tem- The fueling of mono- or bi-propeUants at the perature with communications port to al- Space Station will allow the optimization of low constant interface during fueling and for satellite tank designs in order to achieve min- health checks. imum mass and volume. Additional hardware such as fuel pressure and temperature meters • Dust covers and shielding to prevent dam- and shut-off valves will be required on the satel- age or contamination of ports. lite for safety. The fuel could be carried in bulk by a transport vehicle to the vicinity of the Space 2.9.3 Requirements on Space Station Station or else scavenged from the Shuttle fuel • Propellant/pressurant storage and fueling tanks. facility. Safety concerns may lead to fueling becoming a requirement for Space Station operations. Pos- • Availability of low cost fuel; i.e. scavenged sible concerns are (1) to provide adequate safety fuel with no charge for mass transport to during EVA handling, assembly or testing, and LEO (2) to _,,,,;,t...... th ...... ;t,,s f_,.v...... t-_-g---_,;,,_, the ta.,ko" ° if • Standard fueling ports that provide safe, faults are found in a satellite during testing and quick connect/disconnect. return to a ground-based facility is required. The advantage that fueling may offer is di- • Satellite storage facilities to provide thermal rectly related to the amount of fuel that a satel- and contamination protection prior to OTV lite requires. The use of the OTV, especially connection.

VI - 8 • Standard communications port to allowfu- Shielding,dust coversfor ports and fragile elinginterfaceand healthchecks. equipment to prevent damage during docking or assembly,and to prevent contamina- • Battery charging facilitiesto allow maxi- tion. mum satellitecharge on launch. All loose and protruding equipment must be • OTV requirements per paragraph 2.5.3. non-sharp to avoid damage of EVA suits or other Space Station equipment. 2.10 Assembly of Satellite at Space Station All tools required for assembly should be sti_ndardized. Assembly of satellites at the Space Station may offer advantages, but will require satellite re- Simple assembly/construction features, yet accurate alignment is required. design. Fewer structural restraints such as size, shape, and position of appendages will allow Sizeand mass limitedby Space Station. mission-specialized designs with better performance. The satellite design could make more Efficientpackaging of satelliteequipment. efficient use of Shuttle space, with appendages Full assembly sequence of events priorto able to be stored separately from the satellite. launch,subjectto NASA approval.

2.10.1 Scenario Fully assembled and deployed satellite must withstand standard low thrust. 1. Satellite and appendage modules in Shuttle, transport to Space Station. 2.10.3 Requirements on Space Station 2. RMS and MRMS transferpartsto SS stor- • Communication/data, and battery charg- age area with thermal control. ing/power ports at assembly site.

. Transfer bus to servicing/assembly bay, af- • Dedicated area for construction/assembly fix communications and electrical ports, that will definemaximum size and mass perform bus health check. as well as other constraints(forces,safety, etc.). . Use MR.MS to bring modules and appendages to assembly site,affixvia EVA or • Storage area forunassembled parts(passive teleoperator,perform healthcheck test. thermal and physicalprotection).

5. Continue assembly untilcomplete. • Manipulator system (MRMS) to aid in assembly. 6. Perform checkout afterassembly. • Cherry picker or mobile foot restraint 7. Launch with low-thrust OTV as per sce- (MFR) arms to accesssatellitepartsnot ac- narioin paragraph 2.5.1. cessiblefrom standard work area.

2.10.2 Requirements on Satellite • Ability to pre-flighttest communications systems (coulduse OMV to separatesatel- • Communications/data, and battery charg- literequireddistancefrom Space Station). ing/power ports. • Efficient method of remember- • Standard hard points or hand holds for bus ing/communicating (heads-up display). and all parts to avoid damage during un- packing and assembly. For use in handling • Low thrust(0.05to 0.1G) OTV forefficient by B_MS or EVA. single satellite (1500 kg) transfer t,_ _1_.¢_

VI - 9 2.11 Servicing for GEO Satellites 4. OTV rendezvous with satellite.

Before satellites can be serviced in GEO, years 5. Docking with OTV completed. of success in LEO and testing in GEO will be required. This APO defines GEO servicing as 6. Return to Space Station or Shuttle. retrieval and repair/refurbishment of a satellite in geostationary orbit. This can be as simple 7. Service satellite as in emergency retrieval APO (subsection 2.1.1). as remote retrieval and return to the Space Sta- tion or Shuttle, or as complicated and refined as rbbotic repair/replacement of modules and on- orbit refueling. For this APO, two specific sce- Servicing in GEO narios will be addressed: (1) retrieval and return to the Space Station will be studied as an early . Launch OTV with robotic servicing unit type of servicing; and (2) replacing modules in which is an OMV with smart servicer (some a satellite designed for servicing with a telepres- autonomous operation capability, not nec- ence system. essarily artifical intelligence) and Orbital The retrieval scenario requires added OTV ca- Spacecraft Consumables Resupply System pabilities as briefly addressed under emergency (OSCRS). retrieval (paragraph 2.1.3). The design changes on the satellite to be retrieved become the de- 2. Shutdown satellite propulsion and control ciding factor. Assuming the aerobraYdng capa- systems. bility of the OTV is to be used, the satellite appendages extending beyond the aerobraking 3. Rendezvous and dock with servicer. envelope must be automatically stowed. In ad- 4. Fuel, exchange modules, attempt servicing dition, all unstowed appendages must withstand repair on non-modularized equipment. the transport forces. The replacement scenario requires several 5. Undock and separate from satellite. technology developments on both the satellite and servicing system. Telepresence is defined as 6. Test new on-orbit operation. an operation that is performed roboticaily under the control of a remotely manned system 7. Re-service if errors still exist, make retrieval decision. whose inputs axe entirely artificial sensors such as video or force sensors. For GEO servicing, this 8. Return servicer to storage bay in GEO or requires accommodating several tenth-of-second return via OTV. time delays, and achievement of modularized sateUite designs and remote docking/servicing 2.11.2 Requirements on Satellite equipment. Retrieval from GEO to LEO 2.11.1 Scenarios Appendages automatically stowable or Retrieval from GEO to LEO within OTV aerobraking envelope.

1. OTV with retrieval capabilities or OMV is All unstowed appendages able to withstand launched into GEO orbit. forces of aerobraking and OTV de-orbiting accelerations. 2. S/C automatically stows appendages and despins. Propulsion and control system shutdown capability. 3. Prepare for docking, propulsion system shutdown. Standard grappling fixture for retrievability.

VI - 10 Servicing in GEO 3 Impact on Satellite Design

• Standard grappling fixture. The APOs discussed in the previous subsection each require that modifications be made to the • Propulsion and control system shutdown ca- baseline satellite design. pability. Tables VI-1 and VI-2 give the impact of the APOs on the spinner and 3-axis satellite designs • Modularized components that may fail or respectively. These tables show how the APOs become obsolete. change the mass of the different satellite subsystems. The communications payload, power, and • Detailed satellite documentation. thermal subsystems are not tal_ulated as they are not affected by the APOs. • Refueling capability if desired(Subsection The integration and test column indicates how 2.1.9). much the cost of integration changes. It is pro- portionai to mass change unless the complexity of the task changes. Satellite integration and 2.11.3 Requirements on OTV and Space Station - Retrieval to LEO test costs are increased 5% for APOs using the Space Station. For example, EVA and IVA (ex- • Docking capability via video/laser system. ternai and internal vehicular activity) may be tested in a zero gravity tank before use in space. • Standard grappling feature with grapple fixture. 3.1 Retrieval APOs

• Envelope characteristics; controllability The primary change to the satellitedesignforre- with excess mass during aerobraking. trievalis the implementation of a grapplingfixture. • Requirements as per emergency retrieval For a Shuttle-basedretrieval,the implementa- APO (paragraph 2.1.3). tionisdesignedto have the minimum impact on the satellite.Fixtures are added to the satelliteto allow manual attachment via EVA of 2.11.4 Requirements on Servicer Used in the standard grapplingfixture(SGF) with grap- GEO plingbars and trunnion pin attachment devices. • Standard grappling and docking features. The standard grapplingfixtureisrequiredto allow handlingby the Shuttleremote manipulator • Telepresence/robotic capability. system (RMS) or mobile RMS (MRMS) at the Space Station. • Modular replacement capability. For a Space Station retrieval,the satellitede- signmust containthe SGF in a locationaccessi- • Mobilized fueling capability (0SCRS). ble to the OMV which isused to transportthe satellitefrom itslaunch orbitto the Space Sta- • Docking and servicing provides safe envi- tion. Additionalstructuralmass is requiredto ronment for satellite (plume impingements, support the SGF. arm movement, etc.). A GEO retrievaluses the SGF for retrieval by the OMV and requiresadditionalstructural • Ground-based control via telepresence sys- mass for an OTV docking interfacethat can tem. withstand the forcesexperienced during aerobraking with the OTV. • Low cost operations (satellite value versus Additional changes to the baselinesatellite cost and value gained)...... I_. be to _ow_11 .....re-sating of tl_e propulslon sys-

VI - 11 Mass of Subsystem (kg) Total Int. & Attitude Mass Test Control Propulsion Structure TT&C (kg) (%) BaselineSatellite 40 113 227 37 I 1,059

Difference from Baseline (kg) Int. & APOs at Shuttle Attitude Propulsion Structure TT&C Total Test LEO Retrieval - - +4.1 - +4.1 +.1 GB-OTV to GTO +5.6 +5.3 +23.4 - +34.3 +2.2

Difference from Baseline (kg) Int. & APOs at Station Attitude Propulsion Structure TT&C Total Test LEO Retrieval - - +28.0 - +28.0 +.9 SB-OTV to GTO +5.6 +5.3 +28.0 - +38.9 +12.0 Checkout +3.0 - +28.0 +3.0 +34.0 +1.0 Fueling - +2.0 +18.0 - +20.0 +1.2 GEO Retrieval - - +33.0 - +33.0 +.2 Combination +8.6 +7.3 +23.0 +3.0 +41.9 +14.0

Table VI-I: Impact of APOs on Spinner Satellite Design

Mass of Subsystem (kg) Total Int. & Attitude Mass Test Control Propulsion Structure TT&C (kg) (%) Baseline Satellite 48 114 176 35 1,176 -

Difference from Baseline (kg) Int.& APOs at Shuttle Attitude Propulsion Structure TT&C Total Test LEO Retrieval - +4.1 +4.1 +.2 GB-OTV to GEO -60.0 +13.4 -46.6 -6.0

Deploy appendages M

Combination w -60.0 +13.4 -46.6 -6.0

Difference from Baseline (kg) Int. & APOs at Station Attitude Propulsion Structure TT&C Total Test LEO Retrieval - - +28.0 - +28.0 +1.0 SB-OTV to GEO - -60.0 +18.0 - -42.0 -1.1 Deploy appendages - - +28.0 - +28.0 +6.1 Checkout +3.0 - +28.0 +3.0 +34.0 +11.6

Fueling - +5.0 +23.0 - tLO.U +12.3 GEO Retrieval - - +33.0 - +33.0 +6.3 Combination +3.0 -55.0 +18.0 +3.0 -31.0 +7.1

Table VI-2: Impact of APOs on 3-Axis Satellite Design

VI- 12 tem. These changes have been studied and have piing fixture to allow RMS handling and an relatively little cost impact on the satellite. OTV/MPC (multiple payload carrier) interface. Handling at the Space Station may require ad- 3.2 OTV APOs ditional hand holds and fittings. Integration and test increases due to the com- 3.2.1 Spinner Design bined effects of the additional mass and the ad-

The design of a spin-stabilized satellite is such ditional ground testing required to simulate op- that the power requirements of the payload de- erations at the Space Station. termine the size of the satellite, or conversely the satellite size limits the power available for the 3.2.2 3-Axis Design payload. The spin-stabilized satellites are lim- A 3-axis design can benefit by using an OTV to ited in power based on the solar array area on the replace the current perigee and apogee stages. surface of the satellite. Additional area is some- This allows the satellite size to shrink slightly times created by deploying an external shroud of by removing the apogee motor and bipropellant solar cells to effectively double the length of the tanks. The bipropellant system would be re- satellite. The area inside the new length is not placed by a less complicated hydrazine system used because it is not needed. for station keeping and altitude control. The If the OTV were used, the propellant tanks central structure mass is reduced since it is no that exist inside the main cylindrical body would longer required to support the large propellant not be needed. However, the cost of a Shuttle mass at launch. launch for the satellite would not decrease be- A standard grappling fixture is added to allow cause launch costs are determined by length and ttMS handling while additional handling devices not mass for this satellite design. Therefore, it allow EVA handling at the Space Station. is more economical to include the propellant and Integration and test decreases because of the apogee motor (no increase in launch costs) and significant decrease in satellite mass. use the OTV in place of the conventional perigee motor for transport to geosynchronous transfer orbit (GTO). It is uneconomical to use the OTV 3.3 Deploy Appendage APOs to transfer a spinner satellite to full geosynchro- Deployment or assembly of appendages at the nous orbit. Shuttle or Space Station requires that the satel- Use of an OTV for transfer to GTO requires lite be launched into geosynchronous orbit by that the satellite be capable of spinning up by a low thrust vehicle. An SGF must be added itself. The satellite must also be capable of ob- to allow handling outside of the Shuttle bay or taining a 3-axis reference after separation from Station storage area. There is no appreciable the OTV. The equipment required for these op- mass change to the design as the handling fix- erations are non-radial spin thrusters or rockets, tures mass offsets the mass lost due to removal an additional axis sensor (rate gyro coupled with of the automatic deployment and partial deploy- earth and sun sensor or other system), and on- ment devices. board software to perform the new maneuvers. Use of this APO with spinning satellites is not Control of the satellite thermal environment is desirable. The spin-up procedure requires simul- achieved by spinning the satellite at a relatively taneous stabilization of the equipment platform high rate. The OTV is not capable of spinning to avoid large angular acceleration forces on the at the required rate and it is necessary to in- appendages. This requires active acceleration clude a thermal shroud for each satellite during sensors that are coupled with the platform sta- OTV launch. Even if the 0TV provides a slow bilization device, creating a complicated control spin, the thermal protection would be required loop which decreases the reliability of the single for transfer operations from Station to OTV. failure point in the despun platform configura- _luiL. l_Le _uwuLage in deployment reliability

VI - 13 is offset by the spin-up problemsof the spin- not fit in the Shuttle or expendable launch ve- stabilizedsatellite. hicle envelopes. If a satellite is fully modular, it For the 3-axissatellitedesign,thereis a deft- may be assembled and easily serviced by replac- nite advantage for larger appendages which can ing failed or old equipment. be packaged separately for Shuttle transport and The impact on the design can range from then attached at the Station. alignment guides and manual locks for appendages to precise electrical connectors and me- 3.4 Checkout APOs chanical housings for each subsystem or part of .... 1- .... J.___ rlr_l. _ • __._ e ...... 1.1..

Checkout capability at the Space Station would are also variable. Analysis of various modular require additional on-board failure detection. designs show that up to a 20% increase can be This feature can be approximated by added ca- required. The integration and testing of a satel- pability in the attitude control and TT&C sub- lite to be assembled would show a significant in- systems. A satellite using antenna pattern test- crease due the greater amount of equipment. ing would require deployment of appendages as Contemporary communication satellite sizes previously discussed. The satellite being tested do not appear to benefit from assembly in the at the Space Station would require the standard- sense that a large platform would. In addi- ized power and communications ports and han- tion, spin-stabilized designs do not readily show dling capability which includes the addition of a assembly capability or the potential for simple SGF. These last changes are required of all satel- equipment replacement. lites utilizing the Space Station.

3.7 Servicing APOs 3.5 Fueling APOs Satellite servicing in geosynchronous orbit re- The capability to fuel a satellite at the Space quires that the satellite design be modular is Station creates two major impacts to the base- discussed in the previous paragraph. In addi- line satellite design. tion, the satellite may need to be designed for autonomous changeout of modules for designs 1. The fuel tanks must be modified in order before the year 2000. This may require addi- to allow on-orbit fueling. Current tanks tional specialized fittings and ports that must be are designed to withstand the on-orbit pres- tested by changeout simulations prior to launch. sures. Standardized fueling ports would be Servicing may also require fueling impacts as dis- required. Bipropellant systems would re- cussed in Subsection VI-3.5. quire either two ports or a common port with line switching capability. 4 Discussion 2. Removal of fuel on launch reduces the strength requirements on the satellite cen- 4.1 APO Timelines tral support system and allows a savings in structural mass when lower thrust upper Table VI-3 gives possible timelines of the stages are used. Use at the Space Station first communication industry implementation for each APO. The timelines are based on the fol- will require handling capability and standardized communication and power inter- lowing three assumptions: faces ° • OTV and Space Station timelines as given in Table VI-3; 3.6 Assembly APOs • Reliable backup systems are available; and Implementation of assembly of the satellite may take many forms. Simple assembly of ap- • Successful APO proof-of-concept tests have pendages may be used if future appendages do been carried out.

VI- 14 based APOs (2,3,4) will combine with each other APO (deployment and launch with GB-OTV) but will 1. Emergency retrieval not interact with the Space Station based APOs - at Shuttle (5-10). - at Space Station The Space Station based APOs allow for a - at OTV wider variation of capabilities that may be used 2. GB-OTV to GT0 together for technological and economic advan- 3. GB-OTV to GEO tage over the business-as-usual operations. Some 4. Deploy appendages APOs will be assumed to occur for most satel- - at Shuttle lites, others may eventually be required for safety 5. SB-OTV to GTO reasons, while still others may see only limited 6. SB-OTV to GEO use in the foreseeable future. Some level of 7. Deploy appendages checkout will be performed on every satellite, if - at Space Station only for the business-as-usual thermal and health 8. Checkout checks. Advanced forms of checkout such as an- - at Space Station tenna pattern measurement will need detailed 9. Fueling analysis to ascertain the reliability gained ver- - at Space Station sus the cost to perform the checkouts. 10. Assembly Fueling will need to be cost effective before it - at Space Station is implemented, although Space Station safety 11. Servicing for GEO requirements may make it necessary for some - Transport to LEO types of operations such as assembly or storage. - Servicing in GEO Finally, more advanced APOs such as assembly and GEO servicing may be included, although these types of operations are presently consid- Table VI-3: Timelines for APOs ered high risk, and therefore will probably not be initially used by the commercial communica- It is expected that the communications indus- tions industry. try will take advantage of economically or tech- nically advantageous APOs within two years after successful test. The more advantageous the APO, the sooner it Will be implemented. With- out backup systems such as alternate launch vehicles, the satellite design would be postponed until proof of reliability was available. Satellite design and fabrication will continue to require 3 to 5 years from start to launch. Thus an APO might not be implemented for 6 years after successful testing.

4.2 Combinations of APOs

Several APOs can be combined to build a more efficient scenario. Some APOs such as checkout do not appear to be efficient as a stand-alone option, but when combined with others such as deployment or assembly become quite attractive as _.dded capability. !t is expected that the Shuttle

VI - 15 Section VII

ECONOMICS OF APOs

1 Introduction • The Financial Model is run with inputs appropriate to the APO: This section presents the Task 3B evaluation of the APO's economics. Results are presented - Satellite cost from Price H; forthe spinner and 3-axissatelliteAPOs as de- - Shuttle launch costs based on mass and scribedin SectionVI. The baselinesatellitede- volume of satellite and upper stage; signsare the spin-stabilizedKu-band and 3-axis - Perigee stage costs; hybrid cases described respectivelyin Subsec- tionsIV-3.1 and IV-3.3. The case of the OTV - Launch support costs; launch of the largesatellitedesign describedin - Mission operations costs; and SubsectionIV-4 isalsopresented. - Launch insurance costs. The methodology used to determine the APO valueis describedfirstin Subsection2. Subsec- - The "fee" for the APO is not initially tions3, 4, and 5 discussrespectivelythe eco- included. nomic performance of the APOs for the spin- stabilized,3-axis,and largesatellitedesigns. The Model output is compared with baseline results to determine the value of the APO. 2 Methodology The Model is iterated with the APO "fee" 2.1 APO Values as input until the baseline rate-of-return (DTRR) is equaled. The result is the AP0 The APO value is defined as the "fee" NASA value. (Note that the APO value is subject could charge for the APO that would result to launch insurance.) in the same economic performance as for the business-as-usual scenario. An additional incen- 2.2 Launch Insurance tive of 1% increase in DTRR return, which is equivalent to a capital expense of approximately A major item for potential cost savings from im- $12 M, is judged to be necessary as an incentive plementation of APOs is launch insurance. for potential users but has not been subtracted from the APO fee. 2.2.1 Launch Insurance Rates in 1995 The following procedure is used to determine the APO values: The baseline 20% rate predicted for launch insurance in 1995 is derived from the 16% average • The baseline satellite design is modified to loss rate over the last decade plus 4% overhead. reflect the APO requirements. The satellite The actual future rate will be based on the in- payload is not altered. surance underwriters past experience. Since the launch insurance cost is itself insured, the base- • The new satellite costs are determined via line 20% rate is effectively 25% of all other costs the Price H cost m,,3_1 except, the insurance.

VII- 1 Failures Rate Decrease FailureCategory Number Percent APOs at Shuttle Points Initial stage 5 16 LEO retrieval 1.0 Perigee stage 11 34 Deploy appendages 1.0 Apogee motor 10 31 GB-OTV to GTO 2.0 Spacecraft 3 9 GB-OTV to GEO 5.0 Satellite 3 9 APOs at Station Points

Table VII-I: Incidents of Satellite Failure Fueling 0.0 LEO retrieval 1.0 Deploy appendages 1.0 A variation of the rate by 5% changes the cost Checkout 1.0 of insurance by $5 M and the DTRR return by GEO retrieval 1.0 0.5% (Subsection III-5.2). SB-OTV to GTO 3.0 SB-OTV to GEO 7.0 2.2.2 Incidents of Satellite Failure Spinner combination 6.0 Table VIL1 categorizes 32 incidents of satel- 3-axis combination 9.0 lite failure from 1963 through 1985:7 in the 1960%, 9 in the 1970's, and 16 in the 1980%. Table VII-2: Influence of APOs on Insurance The primary cause of satellite losses and the claims paid by underwriters has been launch vehicle malfunctions [Space, Vol. II, No. 11]. Of to OMV to Station to OTV), it is judged that the 30 civilian communications satellite malfunc- rates will decrease a somewhat lesser amount as shown in Table VII-2. tions up through 1985, 24 have been related to launch systems. The ground-based OTV will act as a test bed The distinction between spacecraft and satel- for the space-based OTV and will have to un- lite in Table VII-1 is that the spacecraft becomes dergo launch from Earth. Therefore, it is judged a satellite the after initial operational capability to be somewhat less reliable than the space- is achieved; i.e. on orbit acceptance testing has based OTV. The other APOs were judged to been successfully completed. have the potential to affect rates as shown in Table VII-2. The fueling APO has no positive impact on in- 2.2.3 Influence of APOs on Rates surance rates. Fueling at the Space Station does T_ble VII-2 gives the projected insurance rate not directly decrease chance of failure. In fact, changes in points for the individual APOs. A the increased complexity of additional valves in one point decrease in insurance gives a rate of the fueling system slightly increases the risk of 19% versus th_ baseline 20%. failure. It is assumed that NASA safety pre- The APO with the largest potential effect on cautions will offset any potential risks associ- launch insurance is the use of the OTV. As ated with fueling. A related issue is that these shown in Table VII-l, 34% of historical losses safety precautions may impose addition paper- occurred in the perigee stage and 31% in the work and hardware requirements that increase apogee ._uu_ai...._ stage. A penec_ly_ __r ., rename,. •, OTV satellite cost. to GTO could resultin a 5.4% reduction and A one point decrease in launch insurance the OTV to GEO could result in a 10.4% reduc- rate is shown for the retrieval APOs. The ac- tion in launch insurance (out of the total 16% tual value of retrieval depends on the details of loss rate). Since the OTV APOs will introduce the satellite design and risk of failure, and on new risks (i.e. transfer of satellite from Shuttle whether there is a satellite repair facility at the

VII - 2 Space Station. The capability for retrieval may 2.3.2 Integral Upper Stage be required as a condition of launch insurance The integralupper stageor perigeestagealterna- and by international treaty to avoid the prolifer- tivesareshown in Table IV-4.The perigeestage ation of space junk. ispurchased and integratedwith the satelliteon The 3-axis combination APO at the Space the ground. The apogee motor isalsointegrated Station of deploy appendages, checkout, fuel- with the satellite.The combined satelliteplus ing, OTV launch to GEO, and GEO retrieval upper stagesisplaced in a reusablesupport cra- is judged to result in a nine point drop in launch dle and transportedby Shuttleto LEO. insurance. This is less than the simple sum of the individual APO impacts since the act of checkout 2.3.3 Orbital Transfer Vehicle at the Space Station reduces the risks of failure from all other causes. Likewise the spinner com- The orbitaltransfervehicle(OTV) isplanned to bination APO at the Space Station of checkout, be availablefor 1995 launches as an alternative fueling,OTV launch to GTO, and GEO retrieval to the conventionalintegralupper stage. Two has a six pointdrop in launch insurance. types of OTVs are planned: Launch insurance rates for combinations of APOs areunlikelyto drop below 10% due to the • Ground-based OTV. The GB-OTV is unknown risksinherentin new space activities. brought up to LEO by the Shuttleforeach It must be emphasized that insurance rates for use, possiblyin the aftcargo carrier(a re- movable cover on back end of the Shut- APOs will not drop until there has been successful demonstration of the APOs. tle externalfuel tank used to give added launch volume). The GB-OTV is a proof- 2.3 Launch Costs of-conceptvehicleindependent of the Space Stationand is planned to be operationalin Launch costs are a significant element in the cost 1995. of the baseline satellite and APO value. Launch costs include the following items: Space-based OTV. The SB-OTV isbased in LEO at the Space Stationand isplanned • Fees for use of the Shuttle; to be operationalin 1998. • Cost of refurbishment of the cradle on which The advantage of the SB-OTV is that its mass the satellite plus upper stage is carried; is already in LEO and Shuttle transportation • Costs of the perigee and apogee stages (up- charges are not incurred. Only the fuel for the per stages); SB-OTV must be transported from Earth to LEO, and this can occur at low cost as filler ma- • Fee for use of the OTV (if used); terial on Shuttle flights that are not completely • Launch support cost (before launch); full. However, the large capacity of the OTV requires that at least two satellites be simulta- • Mission operations costs (from launch to op- neously launched in order to be economical. erational satellite).

Space transportation alternativesare dis- 2.3.4 Comparison of Launch Costs cussedin SubsectionIV-2.8. Launch cost consistsof Shuttlecharges and upper stage charges. Table VII-3 gives launch 2.3.1 Shuttle to LEO chargesforthe 1995 spinnerand 3-axissatellites The Shuttle launch cost is based on mass and with integraland OTV upper stages,forsingle volume of satellite and upper stage mounted on and dual satellitelaunches at 1 G. a cradle. APOs which reduce satellite mass may The 1985 scheduleof Shuttlecharges isused. not affect Shuttle launch charges if length is the OTV charges(in1985 dollars)are based on cur- determining factor, rent NASA estimates as follows:

VII- 3 Spinner Costper satellite Launch Perigee No. Upper Thrust Time in Hours Stage Sats. Shuttle Stage Total (G) GTO GEO Integral - 29.9 3.1 33.0 .05 17 42 SB-OTV 1 21.0 16.2 37.2 .10 11 36 SB-OTV 2 21.0 10.3 31.3 .15 11 36 2.0 1 12 o-_ Cost per satellite Perigee No. Upper Table VII-4: OTV Launch Time versus Thrust Stage Sats. Shuttle Stage Total Integral 35.4 6.9 42.3 2.3.5 Time for OTV Launch GB-OTV 1 51.5 24.1 75.6 GB-OTV 2 34.3 14.1 48.5 The OTV is planned to have several G thrust SB-OTV 1 16.1 30.6 46.7 and be variable over at least a factor of ten. The SB-OTV 2 16.1 18.4 34.5 deployment of appendages at the Space Station (Subsection VI-2.7) requires use of low thrust (0.05 to 0.1 G). Table VII-3: Launch Cost Comparison Table VI-4 gives OTV launch times to GTO and GEO orbits as a function of OTV thrust. • OTV use fee: $7.8 M total; The launch times are quantized due to the use of an integer number of perigee maneuvers for • OTV launch services:$2.7 M singleand greater efficiency. (Three perigee maneuvers are $3.7 M double satellitelaunch. required for .05 G and two for .10 and .15 G launches to GEO. A half orbit wait is required for • Propellant cost at $1200/kg for GB-OTV the 2 G launch to GEO.) These are the minimum and $1765/kg for SB-OTV. times thought possible, and assume that trim burns are not required and that accurate orbit • GB-OTV transported on Shuttle for each determination can be made in one orbit. launch; full cost of transport charged to The time for return of the OTV from GEO satellite launch mission; $2.4 M fee for use to the Space Station is still uncertain because of aft cargo carrier (behind external tank). of the unknown efficiency of the aerobrake. Sev- • 1 G launch;0.i G isrequiredifsatelliteap- eral passes through the upper atmosphere are required. The estimate for return time is 15 hours pendages are deployed,and propellantcost is i0% higher. for a high G return plus OMV rendezvous time. The conclusion is that a low-thrust OTV mis- Propellant cost is the major cost with exact sion will take a minimum of 2 to 3 days.. value depending on launch orbit and number of satellites launched. There can be considerable 2.3.6 OTV Performance variation in SB-OTV propellant cost depending on whether the full cost of transport to LEO is The major cost for use of the OTV is its fuel. paid or if some scavenging or dunnage scheme is Depending on payload mass, number of perigee used_to lower costs.A figureof _¢I7_.v,,/n',_,....,,,uc,,_'"_" maneuvers, ,_uu^-'_ OTV thrust, the ratio of propel- is in the middle of the estimatesof $440/kg to lant mass to payload mass ranges from two to $2200/kg for scavenging,isour best estimateof four for transfer from LEO to GEO. Estimates OTV fuelcost.Non-scavenged fuel,transported for fuel cost (1995) in LEO are $1800/kg with at fullcost by the Shuttleto a fuelingdepot at extremes of $700/kg and $3500/kg. the Space Stationisestimated to cost$3500/kg. The following points about OTV fuel use can

VII - 4 be made: in insurancerate make thisAPO economically attractive.As pointed out under discussionof For a given payload, use of higher OTV insurance,retrievabilitymay be requiredin or- thrust resultsin lesspropellantconsump- der to obtain insurance. tion. The actualcostfora retrievalmissionisnot in- cludedin thisAPO. In caseof a problem with the The OTV is more efficient when fully satellite,the cost of retrievalwould have to be loaded. This means simultaneous launch of weighed againstthe value of the retrievedsatel- multiple satellites using a multiple payload lite. carrier.

OTV aerobrakingforreturnto LEO is very 3.1.2 GB-OTV Launch to GTO efficientand can save 5,000kg of fuelon a 4,000 kg payload launch to GEO. Use of the ground-based (GB) OTV allows a decrease satellite launch costs, removal of the OTV costs are discussed further in Subsec- perigee stage, and a two point reduction in tionVII-2.3. launch insurance. The satellite design has increased cost to allow for the OTV interface, ther- 3 Spinner APO Values mal shielding during OTV launch, and additional capability for three axis orientation and The valuesof APOs forthe spin-stabilizedsatel- spin-up in GTO. Increased mission operations litearesummarized in Table VII-5and discussed are also required. in turn. The spin-stabilizedsatellitedesign was The $12.5 M APO value needs to be signifi- presented in SubsectionIV-3.1. The discussion cantly more than the fee NASA will charge for of spinner APO valueis divided intotwo parts: use of the GB-OTV in order to create savings for APOs at the Shuttle and APOs at the Space the customer. This is unlikely, even for a shared Station. satellite launch, if the cost of Shuttle transport for the GB-OTV must be paid by each mission. Subsection 4.1.2 estimates fees of $14 M for use 3.1 Spinner APO's at Shuttle of the GB-OTV. There are two APOs that are envisioned for the

spin-stabilized satellite at the Shuttle: (1) re- 3.2 Spinner APOs at Station trieval and (2) use of the GB-OTV for launch to geostationary transfer orbit (GTO). There are five APOs plus one combination that axe envisioned for the spin-stabilized satellite at 3.1.1 LEO Retrieval Capability the Space Station. Their values are shown in Table VII-5 and they are discussed in turn. The capability to support a future retrieval mission impacts the baseline satellite design. This 3.2.1 LEO Retrieval Capability impact is described here and in Subsections VII- 3.2.1 and 3.2.5. Since the actual retrieval mis- The retrieval of satellites that have failed to sion is not a normal part of the mission, the costs achieve GEO orbit (on account of perigee motor of retrieval missions are discussed separately in malfunction or other LEO failure) can be carried Subsection VII-5. out by the OMV and the satellite returned to Only minor changes in satellite design are re- the Space Station. There is a one point decrease quired to enable LEO retrieval capability of the in insurance rates. Although more costly than satellite by the Shuttle. The added cost to the baseline system, this APO may be required the satellite is small and there is no increase in as a condition of insurance. The real value of launch cost as the spinner satellite is charged by this APO depends on the probability of failure,

VII - 5 COST(millionsof 1985 dollars) Launch Mission Launch Satellite Shuttle Perigee Support Ops. Insure Total Baseline Satellite $54.3 $29.9 $3.8 $1.6 $2.6 $23.0 $115.1

Difference from Baseline (better/) APO ] APOs at Shuttle Satellite Shuttle Perigee Support Mission Insure Value I

I GB-OTVLEO Retrievalto GTO I <<.1>1.4 > 8.8.... 3.1 - < .3 > 5.01.4 12.51.11

Difference from Baseline (better/) APO APOs at Station Satellite Shuttle Perigee Support Mission Insure Value LEO Retrieval < .6> .... 1.3 .5 SB-OTV to GTO < 2.1> 8.8 3.1 - < .3 > 6.1 13.0 Checkout < 1.6> .... 1.0 < .5 > Fueling < 1.2> .8 - - - < .1 > < .4 > GEO Retrieval < 1.3> .... 1.1 < .2 > Combination < 2.6> 8.8 3.1 - < .3 > 9.5 15.9

Table VII-5: APO Values for Spinner Satellite there is a satellite repair facility at the Space Station.

3.2.2 SB-OTV Launch to GTO

The same considerations apply to use of the space-based OTV as for the ground-based OTV. Spinner Satellite Launch Cost ($M) However, the costs of providing the space-based Launch to GTO Single Dual OTV are estimated to be substantially less since SB-OTV fee 7.8 7.8 it does not need to he carried up to LEO in the OTV fuel 5.7 9.1 Shuttle for each mission. i Launch Services 2.7 3.7 Table VII-6 shows a OTV cost of $10.3 M per OTV Total 16.2 20.6 satellite for a two satellite launch, which is to be Shuttle ::_ 21.0 42.0 compared with the APO value of $13 M. This is Total 37.2 62.6 not enough by itself to provide an incentive for Cost per satellite 37.2 31.3 this APO.

Table VII-6: Launch Costs with SB-OTV 3.2.3 Checkout at Station

The negative value of this APO is due to the increased satellite cost in spite of a one point benefit in insurance rate. However, combination of checkout with an OTV launch is advantageous.

VII - 6 3.2.4 Fueling at Station 4 3-Axis APO Values

The use of fueling has little benefit with the spin- The values of APOs for the 3-axis satellite are stabilized satellite design since the launch cost summarized in Table VII-7 and discussed in continues to be based on satellite length. The turn. The 3-axis satellite design was presented spinner satellite cannot be made smaller due to in Subsection IV-3.31. The discussion of 3-axis the requirement for solar cell area on its surface. APO value is divided into two parts: APOs at It is unlikely that fuel and fueling costs could be the Shuttle and APOs at the Space Station. less at the Space Station than on Earth.

4.1 3-Axis APOs at Shuttle 3.2.5 GEO Retrieval Capability There are three APOs plus one combination that The reasons for including retrieval capability in are envisioned for the 3-axis satellite at the Shut- GEO satellites, other than for a failure in LEO, tle. are not clear. The cost of a retrieval mission is high and any deployed appendages may be dam- aged by an aerobraking maneuver. The safety 4.1.1 LEO Retrieval Capability status of the satellite is another factor that must The capability to support a future retrieval mis- be considered. Finally, the worth of the retrieved sion impacts the baseline satellite design. This satellite must be taken into account. This APO impact is described here and in Subsections VII- would be implemented only if there is potential 4.2.1 and 4.2.6. Since the actual retrieval mis- positive benefit. sion is not a normal part of the mission, the costs of retrieval missions are discussed separately in 3.2.6 Combination of Spinner APOs Subsection VII-5. Only minor changes in satellite design are re- The following combination of APOs shows the quired to support retrieval by the Shuttle. The best potential savings for the spinner satellite added cost to the satellite is small and its greater design: mass slightly increases the Shuttle launch cost. The projected savings of one point in insurance • Space-based OTV to GTO; rate make this APO economically attractive. As • Checkout at Space Station; pointed out under discussion of insurance, retrievability may be required in order to obtain insurance. • Fueling; The actual cost for a retrieval mission is not in- • GEO Retrieval capability. cluded in this APO. In case of a problem with the satellite, the cost of retrieval would have to be If the $15.9 M APO value is compared with weighed against the value of the retrieved satel- the estimated $10.3 M cost of the SB-OTV (Ta- lite. ble VII-6), $5.6 M remains to be divided among NASA fees for the other APOs and satellite 4.1.2 GB-OTV Launch to GEO owner's incentive. Table VII-7 shows an APO value of $37.2 M, primarily due to the five point reduction in launch 3.3 Discussion of Spinner APOs insurance (Table VII-2) and the reduced Shut- The combination spinner APO at the Space Sta- tle charges for launch of the satellite without tion may show a small incentive for use after perigee stage. (However, the NASA fee will NASA fees are paid. There are additional bene- include Shuttle charges for launching the GB- ._J_J._ ul_JL" lcu_uul_y___l;_lL;1;A-__ _lm_ 1 versatility to be realized. OTV.) The mission cost savings are based on the

VII- 7 COST (millions of 1985 dollars) Launch Mission Launch Satellite Shuttle Perigee Support Ops. Insure Total Baseline Satellite $64.6 $35.4 $6.9 $1.6 $2.6 $27.8 $138.8

Difference from Baseline (better/) APO APOs at Shuttle Satellite Shuttle Perigee Support Mission Insure Value LEO Retrieval < .1 > .... 1.7 1.3 GB-OTV to GEO 3.9 19.3 6.3 - .8 13.5 37.2 Deploy Appendage .... .3 1.8 1.7 Combination 3.9 19.3 6.3 - 1.0 14.7 38.8

Difference from Baseline (better/) APO APOs at Station Satellite Shuttle Perigee Support Mission Insure Value LEO Retrieval < .7 > .... 1.6 .7 SB-OTV to GEO 3.4 19.3 6.3 - .8 15.6 39.5 Deploy Appendages < 1.0 > - - - .3 1.6 .7 Checkout < 1.8 > .... 1.3 < .4 > Fueling < 1.6 > .... < .4 > <1.6> GEO Retrieval < 1.1 > .... 1.5 < .3 > Combination 2.1 19.3 6.3 - 1.0 17.6 41.2

Table VII-7: APO Values for 3-axis Hybrid Satellite

VII - 8 creased cost due to ground simulations of EVA Hybrid Satellite Launch Cost ($M) activity by NASA. Once appendages are de- Launch to GEO Single Dual ployed, a low thrust perigee stage must be used. GB-OTV fee 7.8 7.8 The primary economic value of this APO Aft carrierfee 2.4 2.4 comes from the projected one point savings in Propellantcost 11.2 14.4 launch insurance. It is clearly beneficial for com- Launch Services 2.7 3.7 mercial satellite launches, particularly of satel- OTV total 24.1 28.3 lites with many or large appendages, to have EVA assisted deployment. The question is Sat. to LEO 16.1 32.2 whether NASA can supply the required EVA ac- GB-OTV to LEO 35.4 35.4 tivity within the relatively small ($1.7 M) value Shuttle total 51.5 68.7 of the APO. Another important benefit may be the allow- Total cost 75.6 97.0 ing of new capability. The constraint of packag- Cost per satellite 75.6 48.5 ing to allow unfolding antennas and solar arrays is removed, and large or numerous appendages can be sent up separately in the Shuttle. This Table VII-8: Launch Costs with GB-OTV factor has not been considered in the Model. deletion of the numerous transfer orbit maneu- 4.1.4 3-Axis Combination at Shuttle vers that are currently used by the business-as- usual scenario. These orbit manuevers are now The combination of appendage deployment with under NASA control and billed as OTV launch a low thrust OTV launch results in an APO services. value of $38.8 M. However, the additional fuel Table VII-8 estimates the costs for use of the required for a low thrust OTV launch will be GB-OTV. Note that there is a $2.4 M service more than the value of the deployment. Thus, unless there are unusual difficulties with de- charge for using the aft cargo carrier (at the end of the external tank) for carrying the GB-OTV. ployment, business-as-tfsual satellites will prefer These charges reflect the fact that the external OTV launch without deployment. tank must be carried into LEO. For a two satellite launch, the OTV charges are $14 M per satel- 4.2 3-Ax|s APOs at Station lite and the Shuttle charges for the GB-OTV are There are six APOs plus one combination that $18 M per satellite for a total $32 M compared to the $37.2 M APO value. $5 M remains for are envisioned for the 3-axis satellite at the Space Station, as shown in Table VII-7. They will be owner's incentive, and this APO may be feasi- discussed in turn. ble. A further two APOs are also discussed but not The OTV charges are based on a 1 G launch analyzed by the Financial Model. They are (1) and would be approximately $2 M higher for the low G launch requiredforsatelliteswith de- assembly at the Space Station, and (2) servicing in GEO by a remote servicing module carried by ployed appendages. the OMV via the OTV.

4.1.3 Deploy Appendages 4.2.1 LEO Retrieval Capability

The deployment of appendages such as solar ar- The retrieval of satellites that have failed to rays, antennas, and equipment booms in Space achieve GEO orbit (on account of perigee motor by EVA will increase reliability of deployment malfunction or other LEO failure) can be carried and reduce the cost of ground-commanded de- out by the OMV and the satellite returned to ployment sequences. However_ there will be ,_'n-

VII - 9 for a triple or even quadruple hybrid satellite Hybrid Satellite Launch Cost ($M) launch. Launch to GEO Single Dual SB-OTV fee 7.8 7.8 4.2.3 Deploy Appendages at Station Propellant cost 20.1 25.4 This APO is more realizable than its Shuttle- Launch Services 2.7 3.7 OTV total 30.6 36.9 based counterpart. EVA time is expected to cost less at the Space Station, or perhaps IVA can Shuttlecost 16.1 32.2 take the place of EVA. There is also less time constraints on operations at the Space Station Cost per satellite 46.7 34.5 than at the Shuttle.

Table VII-9: Launch Costs with SB-OTV 4.2.4 Checkout at Station

This APO is not advantageous as a stand alone in insurance rates. Although more costly than capability. Additional costs would be incurred the baseline system, this APO may be required for use of the Space Station. However, this APO as a condition of insurance. The real value of does offer worthwhile advantages when combined this APO depends on the probability of failure, with other APOs. the cost of the retrieval mission, and on whether there is a satellite repair facility at the Space 4.2.5 Fueling at Station Station. Fueling satellites from the Space Station requires 4.2.2 SB-OTV Launch to GEO added complication in satellite design and gives limited benefits. There is no positive impact on Due to the launch cost saved by not having to insurance rates. carry the OTV up on the Shuttle, the potential savings for the space-based OTV are greater 4.2.6 GEO Retrieval than for the ground-based OTV. The $39.5 M APO value is primarily due to a seven point The reasons for including retrieval capability in insurance benefit (Table VII-2) and savings in GEO satellites, other than for a failure in LEO, Shuttle launch costs for the perigee motor. are not clear. The cost of a retrieval mission is Table VII-9 estimates costs for use of the SB- high and any deployed appendages may be dam- OTV at $18.5 M per satellite for a two satellite aged by an aerobraking manuever. The safety launch. There is a potentially large incentive of status of the satellite is another factor that must $21.5 M for use of the OTV. be considered. Finally, the worth of the retrieved satellite must ho _en _* ...... , 'T"t.,;oADA A grapmc_ 1 • comparison• o[#_1launcnes• is shown in Figures VII-l, 2, and 3 of the 1995 hybrid 3-axis would be implemented only if there potential satellite for the following cases: positive benefit.

1. Business-as-usual scenario; single hybrid 4.2.7 Combination of 3-Axis APOs satellite plus upper stage on Shuttle. The following combination of AP0s is consid- 2. Single satellite on Shuttle and OTV. ered:

0 _ ...... A._11"_. .... O'I-___,.A.I_ __ 31 f_rr_[r O. .L WU _i:tbl¢llll, l¢_ OR OILUbbl_ _ILLI k,./ ,,L Y .

Use of the OTV is more expensive than a conven- • Deploy appendages at Space Station; tional upper stage for a single satellite launch, • Fueling at Space Station; but results in substantial savings for a dual satellite launch. The OTV has sufficient capacity • Checkout at Space Station;

VII- 10 • Retrieval capability.

The $41 M APO value is considerably larger than the $18.5 M fees incurred by use of the SB-OTV for transport to GEO. The fee charged by NASA for checkout and deployment of appendages at the Space Station should be considerably less than the $22 M difference, leaving margin for a 1% improvement in DTRR (equiv- Shuttle $35.4M alent to $12 M APO value). Perigee $ 6.9M Total $42.3M 4.2.8 Assembly at Space Station

FigureVII-l: Business-as-Usual Launch Economic assessment of an assembly APO has not been included due to the range of assembly options and the complete change of design that may be required. Assembly for initial missions does not seem to show major advantages for the baseline-sized satellite. Satellites requiring assembly will generally be unique designs that cannot be used for a wide range of commercial uses.

Shuttle $16.1 M 4.2.9 GEO Servicing OTV fees $10.5 M Servicing APO economic assessments have not OTV fuel $20.1 M been included due to the lack of basis for initial Total $46.7 M servicing missions. Current satellite hardware generally has similar expected lifetimes (7 to 10 Figure VII-2: One Satellite OTV Launch years). Although much equipment has survived space environments past the predicted lifetimes, much has not. The choice of which equipment to be made replaceable is highly subjective. Results obtained from an economic evaluation would also lack proper reference. Initial use of GEO servicing will be dependent on the cost of a servicing mission versus the cost of building new satellites.

Shuttle $32.2 M 4.3 Discussion of 3-Axis APOs OTV fees $11.5 M OTV fuel $25.4 M Use of the SB-OTV gives large positive bene- Total (2) $69.1 M fits, justifying use of the Space Station. Once at Total (1) $34.5 M the Station to use the OTV, other APOs can be profitably accomplished. The advantage that OTV use gives is so dra- Figure VII-3: Two Satellite OTV Launch t matic that a special case of a large (2,200 kg) sate!lite is an_yzed in Subsection VII-6.

VII- 11 5 Retrieval Missions to Earth for repair and then relaunch. These missions require advance planning, training, and The economics of retrieval missions are exam- special equipment, and thus a "new" Shuttle ined for the 1995 spinner and the 1995 Hybrid mission. satellite designs. The dominating costs are those The mission that originally launched the satel- for transportation (retrieval and relaunch) and lite is not prepared for retrieval unless the prepare-insurance. ration cost is included for all launches. This The discussion is divided into missions using would be unrealistic since the historical failure tha Rhllttl_ mn¢] thn_ 11_11_e th_ _n_r_ _t_tlnn ...... 0 .... _ ...... £_b_ Jt_ UJ.IJ._ JI..IL.LLU, InB_.LI.U-JLAL_ £O_LJLtL.L_I._ blLO#b ¢_£_:; JLL_,.Pb Two types of missionsare possible;(1) repair- retrievable. in-spacemissions and (2) retrieval,transport- to-Earth for repair,and relannch. Missions at- 5.2.1 Repair at the Shuttle tempting repairin space may includethe capa- bilityto returnthe satelliteto Earth for repair Retrieval can be attempted for failures that oc- or salvage. cur prior to a perigee firing maneuver. The scenario assumes that the satellite is placed on a 5.1 Insurance of Retrieval Missions cradle (brought up by the Shuttle) in the Shut- tle bay for repair. Some of these failures may be Insuranceis an important factorin the retrieval simple to repair and the satellite can be released decision. Present day satelliteinsurance poli- and the perigee stage fired. ciespay the insurerthe fullvalue of the satel- The repair-at-Shuttle mission has the follow- liteupon loss.The satellite,as with the Palapa ing steps: and Westar satellitesretrievedby the Shuttle, becomes the propertyof the insurancecompany 1. Plan and train for retrieval afterfailure.Thus the insurancecompany makes 2. Shuttle to LEO with cradle the retrievalor salvagedecisionsin thisscenario, and has the problem of disposingof the repaired 3. Retrieve satellite satellite.The retrievaldecisionwillbe based on the economic valueof the repairedsatelliteand 4. Inspect satellite; make repair decision the probabilityof successof the mission. Once the capabilityfor satelliteretrievalex- 5. Repair satellite ists,the satelliteinsurance policy will change 6. Release satellite such that the ownership of the satellitewillre- main with the originalowner while insurance 7. Wait to verify success covers retrievaland repair attempts. The insurance company willagainexercisecontrolover if the repair attempt is unsuccessful, the de- retrieval,repair,relaunchor salvageoperations. cision could be made to retrieve the satellite to The costof insuranceduring retrievalmissions Earth for repair or salvage. (There may be future will continue to be based on the value of the requirements to remove "junk" from orbit or at satelliteand the probabilityof accidentduring minimum place in a space junkyard.) As the cost the mission. The insurance assumptions of this for the retrieval has been spent, the satellite can subsectionon retrievaldo not have a firm ba- be salvaged to Earth at this stage in the mission sis,but areour bestestimateswithout interviews at little extra cost. If repair on Earth is chosen, w_iL ,L_urance brokers. _ _vo_ _vi _ L_ _L_ _eLauLL_L are as sltv_ in the next subsection.

5.2 Shuttle Missions in LEO Table VII-10summarizes the costof the repair mission.The Shuttlecostassumes that in addi- A Shuttle retrieval mission has two possible sce- tionto a repairkitand grapple fixtures,a cradle narios; (1) repair at the Shuttle and (2) return is transportedfrom Earth to LEO so that the

VII- 12 Item Spinner 3-Axis Item Spinner 3-Axis Shuttlerendezvous 29.9 35.4 Shuttle 29.9 35.4 Retrieval/release 2.3 2.3 Retrieval 2.0 2.0 Ground operations .5 .5 Ground operations .5 .5 Mission operations 2.0 2.0 Mission operations 1.0 1.0 Replaced parts __Q 0 Retrieval Total 33.4 38.9 Total Cost ($M) 34.7 40.2 Repair cost > 5.0 > 5.0 Relaunch cost 37.9 46.5 Re-insurance 3.8 4.6 Table VII-10: Repair-at-ShuttleMission Cost Total ($M) > 80.1 > 95.0 satellite can be retrieved if necessary. A repair- only attempt would utilize less space in the Shut- Table VII-II: ShuttleRetrieval,Earth Repair tle and have less Shuttle cost. The capital costs of the replaced parts of the satellite are not in- The insurance cost for the relaunch is figured cluded, but should not be large considering the at 3.2% of the value of the satellite. This is the limited capability for repair at the Shuttle. pro rata risk (plus 0.6 points overhead), based There are no additional insurance costs for this on Table VII-l, of the repeated Shuttle launch. mission. Since the cost of the repair mission is assumed to be paid by the insurance company, the value of the satellite remains the same. 5.3 Missions with Space Station Retrieval missions with the Space Station can 5.2.2 Repair on Earth use the OMV for LEO retrieval and the OTV for GEO retrieval. Repair is made at the Space The repair-on-Earth mission has the following Station if possible; otherwise transport to earth steps: via Shuttle for repair and then relaunch on the 1. Plan and train for retrieval Shuttle is required. There is an economic incentive to repair at the Space Station as around 2. Shuttle to LEO with cradle $80 M in transportation-to-Earth and relaunch 3. Retrieve satellite costs are eliminated. A final option is to retrieve and salvage (one-way return to earth). 4. Return satelliteto earth

- Make repair/salvagedecision 5.3.1 LEO Retrieval & Repair at Station For reasons mentioned in the previous subsec- - (Satelliteundergoes repairand checkout) tion, it does not seem to be feasible to offer re- 5. Relaunch satelliteon latermission trieval on the same Shuttle flight as the launch. However, retrieval from LEO by the OMV and Table VII-t1 summarizes the costof thismis- transport to the Space Station can occur within sion. The capitalcostsof the replacedpartsof a day of failure. It is very important in order the satelliteare not included in the estimated to place the satellite in a protective environment $5 M minimum for handling and checkout.De- and to turn it off. This saves in hardware life- pending on the nature of the failureand design time and avoids the damaging effects of atomic of the satellite,thesecostsmay be much higher. oxygen and a rapidly changing thermal environ- The relaunch cost includesShuttle costs,new ment. perigeemotor, launch support,and mission op- The mission involves the following steps: erationsas per Table VII-5forthe spinnersatel- liteand Table V!!-7 for the 3-a__YSssate!lAte.

VII- 13 Item Spinner 3-Axis Item Spinner 3-Axis OMV fee (capture) 3.0 3.0 OMV use fee 3.0 3.0 Station fees > 2.5 > 2.5 Station fees 2.0 2.0 Ground operations .3 .3 Ground operations .3 .3 Repair kit launch .2 .2 Repair kit launch .2 .2 Ground simulations 1.0 1.0 Ground simulations 1.0 1.0 OMV fee (release) 2.0 2.0 Mission operations 1.0 1.0 r1"____1 4"_. __ /'_rJlLE\ xu_.,,._o_L_Ivl) > 9.0 > 9.0 Shuttle return 2i.i i6.1 Repair > 5.0 > 5.0 Re-launch 37.9 46.5 Table VII-12: LEO Retrieval/Repair at Station Reinsurance 3.8 4.6 Total Cost ($M) > 75.3 > 79.7 2. Inspect satellite; make repair decision

Table VII-13: Leo Retrieval to Station, Return 3. Deliver repair kit to Space Station to Earth for Repair. 4. Repair satellite 5.3.3 GEO Retrieval g_ Repair at Station 5. Relaunch satellite via OMV The geosynchronous retrieval is combined with Table VII-12 summarizes the cost of this mis- another OTV mission to reduce costs. The costs sion. The capital costs of the replaced parts of assume a shared mission with a two satellite the satellite are not included. There are no fur- launch and one other satellite retrieved. It is as- ther insurance costs. sumed that some lightweight adaptor is carried and used to capture the satellite to be retrieved. The retrieval mission is not the same as the OTV 5.3.2 LEO Retrieval $_ Return-to-Earth mission that launches the satellite. for Repair This mission involves the following steps: This mission involves the following steps: 1. Make retrieval decision 1. OMV retrieves satellite to Space Station 2. SB-OTV to GEO 2. Inspect satellite; make return-to-earth decision 3. Re__e_e satellite

3. Return satellite to earth 4. Inspect satellite; make repair decision - (Satellite undergoes repair and checkout) - (Satellite undergoes repair and checkout) 4. Relaunch satellite via Shuttle 5. Relaunch satellite to GEO via SB-OTV Table VII-13 summarizes the cost of this mis- slon. The capital costs of the replaced ,,_*_ ,,_ _'1r_1,.1^...... 'XTTT - ...... 1 A ...... ,_^__=_ v-K_,e cost of this mis- the satellite are not included. sion. The cost of repair and relaunch is variable The costs for the return-to-Earth portion of depending on the state of the satellite. A new this mission would be the same for a satellite perigee and apogee system (including fuel) and retrieved from GEO and brought to the Space Space Station repair costs would exist as a min- Station as for a satellite retrieved from LEO. imum.

VII- 14 I !tem Spinner 3-Axis Launch Cost ($M) SB-OTV use fee 2.0 2.0 Centaur Perigee Single Dual OTV fuel 48.0 48.0 Shuttle Cost 95 95 Ground operations .5 .5 Centaur Cost 60 60 Mission operations 1.0 1.0 Launch Services ..A4 5 Total Cost ($M) 51.5 51.5 Total Cost 159 160 Cost per satellite 159 80

Table VII-14: GEO Retrieval,Repair at Station Mission Costs. OTV Perigee Single Dual Shuttle Cost 20.0 40.0 OTV Use Fee 7.8 7.8 5.4 Summary of Retrieval Missions Propellant Cost 20.6 26.1 Launch Services 2.7 3.7 Considering the totalinsured cost of the 1995 Total Cost 51.1 77.6 satellitesrange from $115 M for the spinner to Cost per satellite 51.1 38.8 $138 M forthe 3-axisdesign,itis worth expen- ditureto fixor salvagea satellite.The repair of satellitesat the Space Station,if possible, Table VII-15: Hectosat Launch Costs is much lesscostlythan ground repairwith its Shuttletransportationchargesand re-insurance 6.1 Launch Costs expense. GEO retrievalfor ground repairdoes not appear to be economicallyfeasible. Table VII-15 gives Centaur and OTV launch costsfor Hectosat. The Centaur G is the only FiguresVIL4 and VII-5illustratethe costele- upper stagethat has enough capacityto trans- ments ofthe differentretrievalmissionsusingthe port the 2,200kg Hectosat from LEO to GEO. Space Stationforthe 1995 hybrid satellite(total capitalcost$138 M). For a failurein LEO, repair In factthe Centaur upper stagehas enough capacityto carrytwo Hectosats. The cost of the at the Space Station costs$9 M versus$95 M Shuttlelaunch is the same for Centaur plus one for transport to ground for repair and relaunch or two satellites;i.e.a fullload forpricingpur- to LEO. For a GEO failure repair at the Space Station costs $85 M for OTV launch and $103 M poses. The costof the Shuttlelaunch usingthe OTV for launch via a new perigee motor transported isgreatlyreduced since,unlikethe Centaur, the from Earth. A GEO retrieval to Earth for repair space-basedOTV isnot carriedup in the Shuttle and relaunch costs a prohibitive $144 M, more foreach use. The Shuttlechargesaredetermined than the value of the satellite. by the length occupied by the satellitein the Shuttle. The major expense is the costof fuelforthe 6 Hectosat Economics OTV, as shown in Table VII-16 and discussedin SubsectionVII-2.3. There is a largeadvantage The capital cost of the large 3-axis satel- in launchingtwo satellitesat once on the OTV. lite design which is called IIectosat for its 100 transponder payload is summarized in Ta- 6.2 Economic Performance ble VIL16. Details of the design were presented in Subsection IV-4. Table VII-16 gives a total cost of $215.4 M for This design exploits the large carrying capac- the Centaur launch versus $149.8 M for the OTV ity of the OTV. Thus the only APO analyzed launch, a difference of $65.6 M in capital expen- is the use of the space-based OTV to transport ditures between the two launch methods. Note Hectosat from the shuttle to #on qtatlnn_rv orbit.

VII - 15 _'_ SPACE

_,' REPAIR _ ....

\\,, \\

Figure VII-4: Schematic of LEO Retrieval Costs

COST (millions of 1985 dollars) Launch Mission Launch Launch Vehicle Satellite Shuttle Perigee Support Ops. Insure Total Centaur G $88.1 $47.5 $32.5 $1.6 $2.6 $43.1 15215.4 ] Space-based OTV $88.i $20.0 $i8.8 $i.6 $1.8 $19.5 15149.8 I

Table VII-16: Capital Cost for 3-axis Large Satellite

VII- 16 GEO

52M" $46M

o,v ko:- o,v

• "_$_'x,'% --. sPACEs'ra-r,oN $3_%

/ ! s

PERIGEE $16M \ NEW$16M

\ I I

_i _ -"

Figure VII-5: Schematic of GEO Retrieval Costs

VII- 17 • Savings in insurance costs (20% nominal Trans. rate). Upper DTRR Price • Increase in satellite cost. Satellite Stage Return $M/yr 3-Axis Ku Pare D2 18.9 1.27 The combination APOs have an additional value Hybrid Ford 21.9 1.27 due to the fact that some of the same satellite Hectosat Centaur 23.0 1.04 equipment is required for different APOs. Hectosat SB-0TV 23.0 .78 A preliminary evaluation of a modular satellite that would be assembled at the Space Station Table VII-17: Hectosat Economic Performance and be capable of being serviced in GEO is given in Subsection VIII-4. and 3-axis satellites. The charge for the OTV launch has been included in the cost as per Ta- ble VII-15. The economics of the baseline case, the dual satellite launch via Centaur G, is reported in Subsection V-6 and in the nine pages of Table B- 4 in Appendix B. Table VII-17 compares the economic performance of the OTV launch with the baseline Centaur perigee stage. A premium of 1% DTRR return is allowed for Hectosat versus the average 1995 return due to its large size and consequent large risk for the owner and operator. In order to keep the DTRR return at the desired level, a 18% reduction in transponder price is required for the baseline Centaur launched satellite, and a further 25% for a total of 39% reduction for the OTV-launched Hectosat. These dramatic transponder price reductions give a significant financial advantage to large satellites launched with the OTV.

7 Conclusion

Use of the space-based OTV for transport of two or more 3-axis satellites from LEO to GEO is the high value APO that can make commercial satellite operations with the Space Station a reality. Once at the Space Station, other APOs of marginal value but important to the particular mission can be done. Table VII-!8 summarizes the economic value of the individual APOs for the spinner and 3- axis satellites. The major reasons for economic value as follows:

* Savings in STS launch costs due to decrease in mass.

VII- 18 Spinner Satellite($115 M) 3-Axis Satellite ($139 M) APOs at Shuttle Value $M Major Reasons Value $M Major Reasons Capability for LEO Retrieval yes 1.1 Insurance-1% yes 1.3 Insurance -1% GB-OTV from LEO to GTO yes 12.5 Insurance-2% - - LEO-GEO better GB-OTV from LEO to GEO - - Spinnerdesign yes 37.2 Insurance -5% Deploy appendages no - Spinnerdesign yes 1.7 Insurance -1% 3-Axis Combination yes 38.8 STS cost/ins. -6%

Spinner Satellite ($115 M) a-Axis Satellite ($139 M) APOs at Space Station Value SM Major Reasons Value $M Major Reasons Capability for LEO Retrieval yes .5 Insurance -1% yes .7 Insurance -1% SB-OTV from LEO to GTO yes 13.0 Insurance -2% - - LEO-GEO better SB-OTV from LEO to GEO - - Spinner design yes 39.5 Insurance -5% Deploy satellite appendages no - Spinner design no - Sat. cost increase Checkout of satellite no - Spinner design no - Sat. cost increase Add fuel to satellite no - Spinner design no - Sat. cost increase Capability for GEO Retrieval no - Sat. cost increase yes 1.3 Insurance -1% Spinner Combination yes 15.9 STS cost/Ins. -6% 3-Axis Combination yes 41.2 STS cost/Ins. -9%

Table VII-18: Summary of APO Economics

VII- 19 Section VIII SPACE STATION SCENARIOS

1 Introduction a 3-axisdesign in order to best utilizethe capa- bilitiesof the Space Station. This section presents the Task 4 results de- The assembly/servicingscenario requires a scribing three communications satellite system completely new satellitedesign which will not operating scenarios implementing different com- evolveuntilthe Space Station is in orbit. Its binations of APOs. The economic performance IOC (initialoperationalcapability)is unlikely of these scenarios is evaluated and compared to to be 1995 but ratherthe year 2000. the baseline performance. Finally the sensitivity of the results to different insurance and launch cost assumptions is analyzed. 2 Spinner Satellite Scenario The following scenarios are chosen for evaluation: 2.1 Description

The following APOs are utilized with the 1995 • Spinner satellite scenario: spinner satellite design: - Checkout at Station • Checkout .atStation - Fueling at Station - Space-based OTV to GTO • Fueling at Station - Retrieval capability from GEO • Space-based OTV to GTO • 3-axis satellite scenario: • Retrievalcapabilityfrom GEO - Deploy appendages at Station The checkout APO requiresadditionalsatel- - Checkout at Station litecostsand Space Station costsin time and - Fueling at Station equipment. A hypothesizedone pointreduction in insurance givesthis APO value when com- - Space-based OTV to GEO . bined with an OTV launch. Operationally,itis - Retrieval capability from GEO requiredto verifythe health statusof a satellite afteroperationsat the Space Station. Thus it • Assembly/servicing scenario: is likelythat checkout willbe required by the - Assemble satellite at Station insurancecompany. Even though fuelinghas no benefitfor the - Checkout at Station spinnersatellitedesign,itisincludedin thissce- - Fueling at Station narioinorderto satisfypossibleoperationalcon- - Space-based OTV to GEO straintson storingor moving fueledsatellitesat - Service satellite in GEO the Space Station. (The volume of the spinner designissuch that carryingthesatelliteempty to The spinner satellite APO scenario is not eco- LEO does not save money fora shuttlelaunch.) nomically attractive but is included for com- The space-based OTV is used to carry the pleteness. It is our belief that satellites will have satelliteto GTO (geostationarytransferorbit).

VIII- 1 The reusable OTV is more reliable (3 point reduction in insurance) and avoids the necessity of Cost ($M 1985) carrying an expendable rocket motor from earth Capital Station to LEO. The OTV fuel must still be carried from Expenditure Baseline Scenario Earth to LEO, and the OTV economics are dom- Satellite 54.3 56.9 inated by the cost of this fuel. Since appendages STS Launch 29.9 21.1 are not deployed, the OTV can use high thrust. Perigee stage 3.8 .7 Launch support 1.6 1.6 Retrieval capability from GEO (or GTO) re- Mission ops. 2.6 2.3 quires additional mass on the satellite for sup- Insurance 23.0 13._____5 port during retrieval operations. A one point Total 115.1 96.1 reduction in insurance rate is hypothesized for OMV/OTV - 10.3 this capability. Future insurance policies may Station support - 3.0 require this capability. As analyzed in Subsec- Insurance - 2.2 tion VII-5, there may be substantial benefits if Total 115.1 111.6 the satellite falls in a manner that is retrievable.

Table VIII-l: Spinner Scenario Economics

3 3-A.xis Satellite Scenario 2.2 Economic Evaluation 3.1 Description

Table VHI-1 gives a comparison of the cap- The following APOs are utilized with the 1995 ital expenditures for the spinner scenario with hybrid 3-axis satellite design: the Space Station compared to the baseline spin- • Deploy appendages at Station ner scenario (dmcribed in Subsections IV-3.1, V- 3, and Table ]3-1 of Appendix B). A total in- • Checkout at Station surance benefit of 6 points (a rate change from • Fueling at Station 20% to 14%) is hypothesized this scenario. The APO value methodology of Section VH is not • Space-based OTV to GEO used. The cost of the OTV launch is obtained from Table VII-5, and the Space Station support • Retrieval capability from GEO costs for handling, checkout, and fueling are es- The deployment of appendages APO requires timated. Space Station support costs for IVA and possi- Launch insurance is 20% for the baseline case bly EVA operations. The satellite cost is also increased, but is offset by a hypothesized one.point and 14% for the Space Station scenario. Insur- decrease in insurance rates. Once appendages ance appears twice in the table, first for the up- are deployed, the OTV must be used in low per group of capital expenditures and second for thrust (0.1 G) mode for transport to GEO. the lower group. The checkout APO requires additional satel- The result is a $3.5 M savings for the sce- lite costs and Space Station costs in time and nario versus the baseline satellite. The Fi- equipment. A hypothesized one point reduction nancial Model indicates this corresponds to a in insurance gives this APO value when com- 0.2 point increase in the rate-of-return (DTRR) bined with an OTV launch. Operationally, it is from 18.9_ for the baseline to 19.1_ for the spin- required to verify the health status of a satellite ner scenario with the Space Station. Considering after operations at the Space Station. Thus it the uncertainties in the inputs to this calcula- is likely that checkout will be required by the tion, this scenario has marginal value. insurance company.

VIII - 2 Fueling increasessatellitecost and has no insurance benefit.Itisincludedin thisscenarioin Cost ($M 1985) order to satisfypossibleoperationM constraints Capital Station on storing or moving fueled satellitesat the Expenditure Baseline Scenario Space Station. Satellite 64.6 62.5 The space-based OTV is used to carry the STS Launch 35.4 16.1 satelliteto GEO (geostationaryorbit)in a low Perigee stage 6.9 .6 thrustmode (0.1G). The reusableOTV ismore Launch support 1.6 1.6 reliable(7 point reduction in insurance) and Mission ops. 2.6 1.6 avoids the necessityof carryingan expendable Insurance 27.8 10.2 rocketmotor from earthto LEO. The OTV fuel Total 138.8 92.6 must stillbe carriedfrom Earth to LEO, and OMV/OTV - 18.5 the OTV economics are dominated by the cost Stationsupport - 3.5 of thisfuel. Insurance - 2.7 Retrievalcapabilityfrom GEO requiresaddi- Total 138.8 117.3 tionalmass on the satellitefor support during retrievaloperations. A one point reductionin Table VIII-2: 3-Axis ScenarioEconomics insurancerateishypothesizedforthiscapability. Future insurancepoliciesmay requirethiscapa- bility.As analyzed in Subsection VII-5,there rated intoa redesignedsatellite: may be sustantialbenefitsifthe satellitefailsin Assemble satelliteat Space Station a manner that isretrievable. • • Checkout at Space Station 3.2 Economic Evaluation • Fueling at Space Station Table VHI-2 givesa comparison of the capi- talexpendituresforthe 3-axisscenariowith the • Space-based OTV to GEO Space Station compared to the baseline3-axis • Service satellite in GEO scenario (as described in Subsections IV-3.3, V- 5, and Table B-3 of Appendix B). A total insur- In order to be servicedin orbitby an OMV ance benefit of 9 points (a rate change from 20_ plus servicerfrontend, the satellitemust be de- to 11%) is hypothesized this scenario. The cost signed in a differentmanner. The concept is of the OTV launch is obtained from Table VII-7. to have a satellitedesignwith modules that are Space Station support costs for handling, deploy- replacedduring servicing.This leadsto a less ment, checkout, and fueling are estimated. highlyintegratedsatellitedesignthat consistsof The result is a $21.5 M savings for the sce- piecesthat can be transportedseparatelyand nario using the Space Station versus the base- then assembled at the Space Station.Thus the line case. The Financial Model indicates this concept of servicinga satelliteleadsto the po- corresponds to a 1.4 point increase in the rate- tentialfor assembly. of-return (DTRR) from 21.9% for the baseline to The servicingmission is planned to occur af- 23.3_ for the 3-axis scenario with the Space Sta- ternine years and to resultin extensionof the tion. This indicates substantial economic value. satellitelifeby another nine years. The following are examples of items would be replacedby the servicingmission: 4 Assembly/Servicing Scenario • Batteries 4.1 Description • Momentum wheels The following APOs are utilized with the 1995 hybrid 3-axis satellite payload that is incorpo- • Station-keeping fuel

VIII- 3 transponder prices. Other Model assumptions Cost ($M 1985) are as described in Section II. Baseline Capital Scenario The initial capital expenditure is $10 M more Expenditure 1st or 2nd 1st 2nd but the second launch is $45 M less than the Satellite 62.5 68.9 34.8 baseline approach. A simple way to evaluate the STSLaunch 16.1 15.4 8.0 economics is to consider that $10 M was spent Perigee stage .6 .6 .3 9 years earlier in order to save (or earn) $45 M. Launch support 1.6 1.6 .5 This is a return of 18.2% per year, which is less 1_1";@@;_,,, ,.,,-,o _.],,_ 4-t._ _ej *)0"L- --A- At -_A. .... t'r_m'r-*v,_'t e______" J..v 1.6 1.6 vu._l.& VA&_. 4_qU'.V/U A_,b_-'_A-A_bULA'JLI _U JL J[_JL-'k) AO- r" q.)'_,XJ._ Insurance 10.2 10.9 5.6 (10 yr) satellite scenario. The use of the Finan- Total 92.6 99.0 50.8 cial Model for 18 yr scenarios gives 21.8% return OMV/OTV 18.5 19.9 16.5 for the baseline (two successive 9 yr satellites) Station support 3.5 5.0 3.0 versus 21.0% rate-of-return for the 18 yr assem- Insurance 2.7 3.1 2.4 bly/servicing scenario. Total 117.3 127.0 72.7 The conclusion is that the economics of the assembly/servicing scenario are less favorable than launching two successive conventional satellites Table VHI-3: Assembly/Servicing Economics with the OTV. However, our satellite costs derived using Price H are based on a very prelimi- • Thermal control panels nary design of a assemblable, servicable satellite. We recommend that more work be done on de- • Transponder subsystem sign of such a satellite. In particular, relaxation of constraints on compactness may lead to sub- The modular satellite design would be 10% heav- stantial savings in integration and test costs. ier than the baseline satellite of the same capacity. The servicing mission would replace 40% of the mass of the modular satellite. 5 Sensitivity Analysis

5.1 Introduction 4.2 Economic Evaluation The major uncertainties in the scenario eco- Table VIII-3 gives a comparison of the capital nomics compared to the baseline case lie in the expenditures for an 18 year assembly/servicing areas of insurance costs and launch costs. Using scenario with the Space Station compared to a the 3-axis scenario (Table VIII-2) as an exam- baseline scenario with two successive hybrid 3- ple, these costs are varied through a reasonable axis satellite launches each having a nine year lifetime. The baseline scenario uses the 1995 3- range and their effect on economic performance assessed. axis hybrid satellite with 9 year lifetime and sce- A major study assumption is that changes in nario per Subsection VIII-3. It is assumed that satellite mass due to APOs are not used to alter the second satellite has the same cost as the first. The insurance rate is assumed to be the same the payload (i.e. number of transponders). The 3-axis scenario results in a 31 kg mass savings (11%) for assembly/servicing scenario as for the for the satellite. The effect on economic per- baseline case. OTV launch costs are based on the formance of adding this mass in transponders is same assumptions as Table VII-7. Space Station calculated. support costs for the initial assembly and subsequent servicing mission are estimated. 5.2 Launch Insurance An important assumption of the 18-year Fi- nancial Model is that revenues are not inflated Launch insuranceconsiderationsare discussed while costs are inflated at 4% per year. This was at length in Subsection VII-2.2. The critical done to reflect the long term trend of decreasing point isthe difference,ifany,between the Space

VIII - 4 Cost ($M 1985) I 3*Axis Satellite ($CostM) Rate-of-return(_) Cost Change Baseline Scenario Delta Originalcase 138.8 117.3 21.5 Scenario (11_) 117.3 23.3 OTV plus 50% 138.8 127.7 II.I Baseline (20_) 138.8 21.9 Scenario (20_) 130.5 22.5 STS minus 50% 116.7 108.3 8.4 OTV minus 50% 138.8 106.9 31.9

Table VIII-4: Influenceof InsuranceRate STS/OTV -50% 116.7 97.9 18.8

Stationscenarioand the baselinecase insurance Table VIII-5: Influence of Launch Costs rate. The scenariosassume a 6 point and a 9 point differencerespectivelyforthe spinnerand The mass changes involved are relatively 3-axisscenarios. small, _40 kg, and fall within the mass Ifitisassumed thereisno differencein insur- margin of the satellite. ance ratesdue to the scenarios,the cost of the spinnerscenarioincreasesby $8.3M to $119.9M, Sincethe capacityofthe OTV (contraryto versus$115.1M forthe baseline.The 3-axissce- conventionalupper stages)is much larger nario increasesin costby $13.2 M to $130.5 M, than the mass of a singlesatellite,thereis versus$138.8 M forthe baseline. no upper limiton how much the payload The conclusionisthat without insuranceben- can be increased. efitsthe spinner scenariois definitelynot vi- able.The 3-axisscenariocontinuesto show ben- There isa problem in how to use the mass efits,although reduced greatlyfrom $21.5 M to savings to enhance the payload. More $8.5 M. Table VIII-4 summarizes the satellite transponders may imply more antennas. cost and rate-of-return(DTRR) changes forthe Alternately,more power could be supplied 3-axisscenariowith 20% insurancerate. to each transponder or smallerbandwidth transponders could be used. The resultis a specializedpayload to exploitthe APO 5.3 Launch Costs scenario. Table VHI-5 summarizes the effects of some substantial changes in launch charges on system The spinner and 3-axissatellitedesignsused costs. The baseline and 3-axis scenario costs are in thescenariosare describedin Tables IV-7 and compared for each launch cost assumption. The IV-10 respectively.The satellitemass changes scenario continues to show value regardless of the for the combination APOs used in the scenar- launch cost change. The economics are very sen- ios are +41.9 kg for the spinner (Table VI-1) sitive to changes in OTV costs. The assumption and -31 kg forthe 3-axis(Table VI-2). The ef- of STS charges being reduced by 50% also has a fecton economic performance of usingthe 31 kg large negative effect on scenario economics. mass saved by the 3-axisscenarioto increasethe satellitepayload willbe calculated. The 31 kg of mass correspondsto the addition 5.4 Use of Mass Savings to Increase of two more Ku-band transponders,taking into Payload account the mass increaseof thepower and ther- The methodology for determining APO eco- mal subsystems as wellas the payload. Using the nomic value as set forth in Subsection VII-2 Price H model, the satellitecost increasedfrom states that the "satellite payload is not altered _ $62.47 M to $65.8 M forthe 3-axisscenario.It in response to satellite mass changes due to the isassumed that the two added transpondersare APOs. The reasons behind this assumption are the same as the existing33 W, 36 MHz band- as follows: width Ku-band transponders.

VIII- 5 The FinancialModel isused to calculatethe economic performance of the enhanced satellite. The resultisa 23.4_ rate-of-return(DTRR) versus the baselinescenario23.3% return. This is an insignificantimprovement, but analysishas been based on a relativelycrude analysis.Itdoes point out, however, that use of the OTV APO may have more value than is reported by this

6 Conclusions

The spinnerscenariohas a smallnominal value with the hypothesizedcosts,but is sensitiveto changes in insuranceand launch costs.This scenarioisjudged to be not economicallyviable. The 3-axisscenarioshows substantialvalue which continuesto be positiveunder worst case insurance and launch cost assumptions. This scenarioisjudged to be economicallyviable. The assembly/servicingscenario has equal valueto two successivelaunchesof the 3-axisscenario.Consideringour relativelycrude analysis of the satellitedesign,we believethis scenario has promise of better performance and should be analyzed in more detail.

VIII - 6 Section IX

SPACE STATION REQUIREMENTS

1 Introduction The storage bay should be large enough to accommodate up to four 1995 satellite designs This sectionpresentsthe Task 4 resultson the for storage and an additional area for servic- functionaland technicalrequirements imposed ing. A 10 m x 10 m x 20 m volume should on the Space Station by the implementation of be sufficient. The bay should be enclosed for the scenariosof SectionVIII.The requirements micrometeorite and passive thermal protection are presentedin the followingcategories: which can be augmented by internal satellite • Space Station hardware requirements thermal systems. In addition, standard power and communications ports should be available • OMV requirements so that satellites can use Space Station power and can be monitored from inside the manned • OTV requirements modules. Power consumption is expected to be • Operations and policy in the range of 10 W to 400 W per satellite and data rates are low (1200 b/s). The servicing/storage bay should be located 2 Space Station Requirements near the OTV facility and other transportation nodes for the Shuttle and OMV. Since the mo- Space Station hardware requirements are dis- bile remote manipulator system (MRMS) used cussed forthe followingitems: for satellite transfer moves slowly, the time of • Servicingand storagebay transfer becomes a concern for the power, thermal, and telemetry systems. Increasing satellite • Automated transferfacilities batteries for this procedure should be avoided. • Fuelingfacilities Another issue is the torque noise during satellite transfer, which may affect other operations requiring a stable environment. (Torque noise is 2.1 Servicing and Storage Bay mechanical vibration and oscillation caused by The primary requirement on the Space Station use of the MRMS.) is the inclusion of a servicing/storage bay in the IOC design. An early servicing bay would be 2.2 Automated Transfer Facilities used for unscheduled retrieval missions where a perigee motor or ELV upper stage fails, leav- A universal retention system should be devel- ing the satellite in an orbit not accessible to the oped to reduce the required hardware weight on OMV. satellite systems, and allow automated docking The economic and environmental advantages and release. of retrieval missions to the Space Station jus- Automated systems such as the MRMS (mo- tify the initial inclusion of this area. The servic- bile remote manipulator system) are needed to ing/storage bay would later be used for storage transfer satellites and equipment to and from the of satellites prior to using the OTV and for stor- Shuttle, OTV, OMV, and storage/servicing bay. ing and assembling small satellites. Systems with a high level of articulation and co_-_-

IX-1 trol are desired to reduce demand for EVA ac- ning of the OTV will assist in maintaining the tivity such as deployments and connections. thermal environment of the satellites. The OTV should be capable of maintaining accelerations of 0.1 G or less to allow appendage 2.3 Fueling Facilities deployment at the Space Station. This feature Fueling facilities may be required at the Space would also be required for large communications Station. Although there is no economic advan- antennas and platforms not covered in this study. tage for fueling at the Space Station, other fac- There should be at least two OTVs in order to

_.J_ _JU,_,_L _ _.SJ_L_bbJL_ _¢bULJ_LI_.J.L DGkL_b._ J.JLlCb_ .IL_4_I,.LIJL_. lb, _ a,e,.,nJ._ _L,.n J._bl.l_V_ Ob ILL_,J.IUILt,.bZUIJL111,_ %) J. Y bU blL_ as may APOs such as assembly. The issues sur- Space Station for repair. An OTV based at the rounding fueling should be examined in depth Space Station is preferred to the ground-based before placing requirements on the Space Sta- alternative in order to respond more rapidly to tion. an emergency retrieval.

3 OMV Requirements 5 Operations and Policy

There are other requirements that the satellite The initial use of the OMV is as a space tug to retrieve stranded satellites from LEO as well as communications industry places on the Space Station infrastructure beyond hardware or scar- transfer cargo from ELVs to the Space Station. ring needs. It is important that scheduled use This requires space-basing of an OMV in order to be available for unscheduled events such as of the Space Station, OMV, or OTV not be in- terrupted. Many of the APOs using the Space emergency retrieval. Station will have no alternative if the service is The OMV would need to be attachable to a delayed due to higher priority government mis- servicing device such as the Smart Front End for sions. The Space Station should adopt a set of GEO servicing. This combination should have operations and policies that insure its users a the capability of servicing several satellites on high degree of reliability. each mission. Methods for changing out modules The procedures required on the ground for should be standardized and tested in LEO prior Space Station safety should become streamlined to use in GEO. without hindering the determination of safeness. There should be at least two OMVs in order Present NASA safety requirements for the Shut- to be able to retrieve a malfunctioning OMV to tle require a large amount of paperwork and ad- the Space Station for repair. ditional test time prior to launch. The safety requirements for the Station should be studied 4 OTV Requirements far in advance so that an efficient safety regula- tion program can be utilized. The OTV offers the largest economic advantage Space Station policies should be devised so of the APOs evaluated in Section VII. The re- that termination of services will not occur without sufficient lead time to allow satellite man- quirements placed on the OTV by this study are within the scope of the capabilities required by ufacturers to phase Space Station APOs out of the initial OTV studies. Several satellites must their designs. Reduction of services due to safety be launched at once in order for the relatively or accidents should not be placed only on the commercial users. large capacity 0TV to be economical. This requires a multiple payload carrier (MPC) which should use a standard retention system compat- 6 Summary of Requirements ible with the Space Station servicing bay. The OTV should provide power and telemetry The requirements on the Space Station and its links to the satellite while in transit. Slow spin- associated equipment are summarized in the fol-

IX - 2 lowingsubsections: • Simultaneous fueling of bipropellants, pressurant. 1. Spinner and 3-axis scenario requirements • To be determined flow rate for propellants. 2. Assembly/servicing scenario requirements • Automatic emergency shutdown of fueling 3. LEO retrieval/repair scenario requirements system.

6.1 Spinner/3-axis Scenarios • Telemetry data link for satellite check, fueling data. Provide Storage Area for Satellites

• Provide satellite storage area for 4 satellites. Provide Checkout Facilities for Satellites

• Physical protection from contamination, • Monitor electrical and physical parameters meteorites, etc. of satellite.

• All areas accessibleto MRMS and EVA. • Connect OTV to MPC while docked.

• Standard retentionsystems forsatellites. • Telemetry links to all satellite stations.

• Provide standard voltdc power to satellites; • Equipment for monitoring, trouble shoot- tbd (to be determined) W. ing, and analysis. • Telemetry data linksforthermal and health checks. Provide MRMS with Satellite Handling Capability

Provide Servicing Area for Satellites • Accessible to all satellite areas. • Standard NASA toolsavailablewith back- • Standard grapple fixtures. ups. • Smooth, automatic transfer of satellite. • Proximity to storagearea to facilitatetrans- port. Provide Deployment Capability for Satel- • Physical protection from contamination, lites on OTV meteorites,etc. Provide Space-Based OTV • Accessibleto MRMS, fullRMS capability • Low thrust capability, 0.1 G." at assembly site. • Multiple payload carrier (MPC) with stan- • Standard retentionsystems forsatellites. dard volt dc power. • Provide standard voltdc power to satellites; tbd W. • Slow spin capability for thermal control of satellites. • Telemetry data linksforsatellitechecks. 6.2 Assembly/Servicing Scenario Provide Fueling Capability Assembly Requirements • Low costpropell£ntsand pressurantsavail- able. • Requirements as listed for 3-axis satellite scenario. • Storage areasforpropellants,pressurants. • Storage for unassembled modules with ther- • Fueling quick disconnects. mal control.

IX-3 • Provide standard toolsfor satelliteassem- bly.

• Advanced checkout capabilities.

Servicing Requirements

• "Smart" servicer capable of replacing snap- on modules.

0n-orbit fueling capability.

Multiple mission capability in order to service several satellites during one trip to GE0.

6.3 LEO Retrieval/Repair Scenario Requirements

• 0MV based at Station with quick reaction (1 day) for emergency retrieval (standardized retrieval system with minimum impact on satellite design).

• Storage/servicing area with passive thermal control and micrometeorite protection.

• Standard volt dc power supply outlet for satellite, tbd W.

• Standardized communications port to determine satellite health and monitor thermal status.

• Provide EVA capability to attempt minor satellite repair in servicing bay.

IX-4 Section X

RECOMMENDATIONS

1 Introduction the Shuttle. The logistics and cost of returning, refurbishing, and relaunching an OTV have not This section presentsthe recommendations of been determined. A fueling system of a space- the study on the followingissues: based OTV could possibly be simplified by using • Desirabilityof space-basedOTV ground launched tanks that could be "snapped" into the OTV in space. This concept could de- • Study of retrievalmissions crease the cost of launching and retrieving the entire OTV, and may be more cost effective than • Study of modular satellitedesign scavenging systems with long term space-based fueling depots. • Technology development The final OTV issue is the cost comparison be- • Purpose of Space Station tween space-based and ground-based operation. The obvious advantage of space-basing is that the OTV structure does not need to be carried 2 Need for Space-Based OTV from Earth to LEO for each mission. As shown in the sensitivity analysis of Subsection VIII-5 The space-based OTV is recommended rather and discussion of launch costs in Subsection VII- than a ground-based OTV for several reasons. 2.3, economics are very sensitive to launch cost Most important is minimization of possible assumptions. Perhaps future reduction in launch scheduling problems. Operations based at the costs will make this point academic. A careful Space Station such as deployment and assembly analysis of OTV costs is needed. would need to be scheduled simultaneously with The feasibility of many APOs may be im- the ground launch of the ground-based OTV. pacted adversely by use a ground-based OTV Delays occurring on the ground (for example, due to operational constraints. due to weather) could disrupt schedules at the Station due the necessity for preparing and pro- tecting multiple satellite. Conversely, satellite 3 Study of Retrieval Missions operation delays at the Station could delay the ground launch. The ground-based OTV, if fu- The economics of retrieval missions is discussed eled, requires a large amount of power to prevent in Subsection VII-5. There can be substantial cryogenic boil-off losses. benefits in retrieval missions and we see this to Another reason for recommending a space- be a natural function of the Space Station from based OTV is risk. Requiring a ground launch its position as a "gateway to space" and trans- for every OTV launch adds risk to the system portation node. which could affect the insurance advantage asso- We recommend that NASA sponsor a study ciated with the OTV. of the economics of retrieval missions and the A concern raised by this study is the opera- influence of retrieval on the insurance industry. tional aspect of interfacing a ground-based OTV The goals of this study would be to more accu- with the Station and a return vehicle such as rately demonstrate the value of retrievability for

X-1 the satellite and to more closely define the op- • OTV with low thrust and based in space erational aspects of retrievability on the Space Station and the satellite. • RF interfaces for assemblable satellite Involvement of insurance company represen- • Telerobotics for IVA operations and servic- tatives in the study is desirable, along with a ing methodology to assess financial risk (defined as the standard deviation in the rate-of-return) for different retrieval scenarios. 7 Purpose of Space Station

We see the highest use of the first Space Station 4 Study of Modular Satellite as a transportation node with associated stag- Design ing and assembly areas. Some requirements like safety are of continuing concern, but the inap- A modular satellite design is required for imple- propriate placing of instruments or experiments mentation of assembly and servicing scenarios. on the initial Station that place further difficult We recommend that NASA sponsor a study in requirements is to be avoided. this area in order to stimulate the satellite man- The value of the Space Station as transporta- ufacturing industry to consider these designs. tion node will vanish if it is too difficult to use. A future NASA or government satellite should The commercial sector will not use something then incorporate a requirement for serviceability that places addition financial risks on the opera- and/or assembly in order to demonstrate feasi- tions, such as time delays in on-orbit operation. bility. For instance, a one month delay in IOC is equivalent to 0.4% rate-of-return (DTRR) or $5 M initial cost. 5 Study of ELV Use

NASA has recently said that commercial launches will be phased out of the Shuttle program. Expendable Launch Vehicles (ELVs) will need to be used for transport from Earth to LEO (near the Space Station), instead of using the Shuttle as assumed in this study. There are potential impacts on launch costs and risks, on the APOs, and on the requirements placed on the ELV system. A study is needed to determine the effect that launching commercial communications satellites to LEO on ELVs would have on the APOs, and the requirements placed on the ELVs. The ELV system needs to be designed to supply regular and reliable transportation form Earth to Space Station in order to facilitate the APOs.

6 Technology Developments

The following technology developments are recommended:

• Modular satellite designs

X-2 NASA Contract No. NAS3°24253

COMMUNICATIONS SATELLITE SYSTEMS OPERATIONS WITH THE SPACE STATION

APPENDICES for TECHNICAL REPORT

Prepared by FORD AEROSPACE & COMMUNICATIONS CORPORATION WDL Division

Spacecraft Systems Operation Advanced Systems Department 3939 Fabian Way Palo Alto CA 94303

Prepared for NASA, Lewis Research Center Cleveland, OH 44135 LIST of TABLES

A-I.1 Financial Analysis (1985 C-band System) ...... A - 2 A-1.2 Income Statement (1985 C-band System) ...... A-3 A-1.3 Balance Sheet (1985 C-band System) ...... A-4 A-1.4 Sources and Uses of Funds Statement (1985 C-band System) ...... A-5 A-1.5a Revenue Assumptions - C-band Transponders (1985 C-band System) . . A-6 A-1.6 Capital Expenditure and General Assumptions (1985 C-band System) . . A-6 A-1.7 Operating Expenditure Assumptions (1985 C-band System) ...... A-8 A-1.8 Financing Assumptions (1985 C-band System) ...... A-9 A-1.9 Tax Assumptions (1985 C-band System) ...... A- 10 A-2.1 Financial Analysis (1985 Ku-band System) ...... A- 11 A-2.2 Income Statement (1985 Ku-band System) ...... A-12 A-2.3 Balance Sheet (1985 Ku-band System) ...... A- 13 A-2.4 Sources and Uses of Funds Statement (1985 Ku-band System) ...... A- 14 A-2.Sb Revenue Assumptions - Ku-band Transponders (1985 Ku-band) ..... A- 15 A-2.6 Capital Expenditure and General Assumptions (1985 Ku-band) ...... A-16 A-2.7 Operating Expenditure Assumptions (1985 Ku-band System) ...... A-17 A-2.8 Financing Assumptions (1985 Ku-band System) ...... A- 18 A-2.9 Tax Assumptions (1985 Ku-band System) ...... A- 19 A-3.1 Financial Analysis (1985 Hybrid System) ...... A-20 A-3.2 Income Statement (1985 Hybrid System) ...... A-21 A-3.3 Balance Sheet (1985 Hybrid System) ...... A-22 A-3.4 Sources and Uses of Funds Statement (1985 Hybrid System) ...... A-23 A-3.5a Revenue Assumptions - C-band Transponders (1985 Hybrid) ...... A- 24 A-3.Sb Revenue Assumptions - Ku-band Transponders (1985 Hybrid) ...... A-25 A-3.5c Revenue Assumptions - Other Transponders (1985 Hybrid) ...... A- 26 A-3.6 Capital Expenditure and General Assumptions (1985 Hybrid) ...... A-27 A-3.7 Operating Expenditure Assumptions (1985 Hybrid System) ...... A-28 A-3.8 Financing Assumptions (1985 Hybrid Satellite System) ...... A-29 A-3.9 Tax Assumptions (1985 Hybrid System) ...... A-30

A-i LIST of TABLES (Continued)

B-I.1 Financial Analysis (1995 Spinner System) ...... B-2 B-1.2 Income Statement (1995 Spinner System) ...... B-3 B-1.3 Balance Sheet (1995 Spinner System) ...... B- 4 B-1.4 Sources and Uses of Funds Statement (1995 Spinner System) ...... B- 5 B-1.Sb Revenue Assumptions - Ku-band Transponders (1995 Spinner System) . B- 6 B-1.6 Capital Expenditure and General Assumptions (1995 Spinner System) . . B-6 B-1.7 Operating Expenditure Assumptions (1995 Spinner System) ...... B- 8 ]3-1.8 Financing Assumptions (1995 Spinner System) ...... B- 9 B-1.9 Tax Assumptions (1995 Spinner System) ...... B- 10 ]3-2.1 Financial Analysis (1995 3-Axis System) ...... B- 11 B-2.2 Income Statement (1995 3-Axis System) ...... B- 12 B-2.3 Balance Sheet (1995 3-Axis System) ...... B- 13 B--2.4 Sources and Uses of Funds Statement (1995 3-Axis System) ...... B- 14 ]3-2.5b Revenue Assumptions - Ku-band Transponders (1995 3-Axis) ...... B- 15 B-2.6 Capital Expenditure and General Assumptions (1995 3-Axis) ...... B- 16 B-2.7 Operating Expenditure Assumptions (1995 3-Axis System) ...... B- 17 B-2.8 Financing Assumptions (1995 3-Axis System) ...... B- 18 B-2.9 Tax Assumptions (1995 3-Axis System) ...... B- 19 B-3.1 Financial Analysis (1995 Hybrid System) ...... B- 20 B---3.2 Income Statement (1995 Hybrid System) ...... B- 21 B---3.3 Balance Sheet (1995 Hybrid System) ...... B- 22 B-3.4 Sources and Uses of Funds Statement (1995 Hybrid System) ...... B- 23 B--3.5a Revenue Assumptions - C-band Transponders (1995 Hybrid) ...... B- 24 B--3.Sb Revenue Assumptions - Ku-band Transponders (1995 Hybrid) ...... B- 25 B-3.5c Revenue Assumptions - Other Transponders (1995 Hybrid) ...... B- 26 B---3.6 Capital Expenditure and General Assumptions (1995 Hybrid) ...... B- 27 B-3.7 Operating Expenditure Assumptions (1995 Hybrid System) ...... B- 28 B-3.8 Financing Assumptions (1995 Hybrid Satellite System) ...... B- 29 B-3.9 Tax Assumptions (1995 Hybrid System) ...... B- 30

A - ii Appendix A FINANCIAL MODEL OUTPUT 1985 SATELLITE SYSTEMS

This appendix contains the Financial Model results for the three 1985 satellite designs described in Section II:

1. C-band, spinner satellite (Table II-2) 2. Ku-band, 3-axis satellite (Table II-3) 3. Hybrid (C and Ku-bands), 3-axis satellite (Table II-4)

These resultsarediscussedin SubsectionII-5,and consistof nine pages of output foreach satellitesystem.

The output pages for each satellite contain the following information which is described in Subsection II-4.5.1:

Page 1: Financial analysis Page 2: Income statement (cash basis) Page 3: Balance sheet (cash basis} Page 4: Sources and uses of funds statement Page 5: Revenue assumptions (three separate pages for C-band, Ku-band, and Other transponders} Page 6: Capital expenditure and general assumptions Page 7: Operating expenditure assumptions Page 8: Financing assumptions Page 9: Tax assumptions

The tables are numbered A-I.1, A-1.2, ... A-1.9 for the C-band system, Table A-2.1 etc. for the Ku-band system, and Table A-3.1 etc. for the hybrid satellite system.

The bottom line _result" is the dual terminal rate of return which is presented on page 1. Pages 5a, b, c show the revenues-from lease of the transponders. Page 6 shows the breakdown of costs for the satellite.

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! . o Appendix B

FINANCIAL MODEL OUTPUT 1995 SATELLITE SYSTEMS

This appendix contains the Financial Model results for the four 1995 satellite designs described in Section IV:

1. Ku-band, spin-stabilized satellite (Table IV-7) 2. Ku-band, 3-axis satellite (Table IV-9) 3. Hybrid (C and Ku-bands), 3-axis satellite (Table IV-10) 4. Large Ku-band, 3-axis satellite (Table IV-13)

These resultsare discussedin Section V, and consistof nine pages of output for each satellitesystem.

The output pagesforeach satellitecontainthe followinginformationwhich isdescribed in SubsectionII-4.5.1:

Page 1: Financialanalysis Page 2: Income statement (cashbasis} Page 3: Balance sheet (cashbasis) Page 4: Sources and uses of funds statement Page 5: Revenue assumptions (three separatepages for C-band, Ku-band, and Other transponders) Page 6: Capital expenditureand generalassumptions Page 7: Operating expenditureassumptions Page 8: Financing assumptions Page 9: Tax assumptions

The tablesare numbered B-1.1,B-1.2,...B-1.9 forthe Ku-band spinnersystem, Table B- 2.1 etc. for the Ku-band 3-axissystem, Table B-3.1etc. for the hybrid satellitesystem_ and Table B-4.1 etc.forthe largeKu-band 3-axissystem.

The bottom line "result" is the dual terminal rate of return which is presented on page 1. Pages 5a, b, c show the revenues from lease of the transponders. Page 6 shows the breakdown of costs for the satellite.

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E • ;_i_i-_l _'_"_""_i,' I i i i t , i I ° t ! -- i _. i: - il. ! - i._.i - t t t 41 t i I I i I 1. Re_rt No. 2. Government Accession No. 3. R_lpient's Catal_ No. CR 179527

4. Title and Subtitle 5. Re_rt Date February 1987 Communications Satellite Systems with the

Space Station. Volume II - Technical Report 6. Performing Organization Code

480-43-02 7. Author(s) 8. Performing Organization Report I_o.

K. Price, J. Dixon and C. Weyandt 10. Work Unit NO.

g. P_ormtng O_i_tlon Name _d A_mss Ford Aerospace & Communications Corporation 11. Contract or Grant No. 3939 Fabian Way NAS3-24253 Palo Alto CA 94303 13. Type of Report and Period Covered Final 12. Sponeonng Agency Name and A_mee July 1985 - Sept. 1986 NASA, Lewis Research Center 21,000 Brookpark Road 14. Sgonlodng Agency Code Cleveland Ohio 44135

15. Supplementa_ Notes NASA Contract Technical Manager: Mr. Steven M. Stevenson. One other volume prepared: Volume I, Executive Summary.

16. Abstract

See .c,:.t_- [..

17. Key Words (Suggested by Author(s)) 18, Distribution Statement Communications satellite. Space station. Space-based activities. General Release Orbital transfer vehicle. Econon]iGs of space station. ig. 3eCurity Cias,,if. (of this report) 20. Security Classif. (of this page) 21. No. of pages 22. Price" Unclassified Unclassified

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