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REVISIONS LTR DESCRIPTION DATE APPROVAL REV Original Release of Sea Launch User's 26 March Tom Alexander New Guide 1996 System Engineering A This document supersedes the Sea Launch July 1998 Dennis D. Smith User's Guide in its entirety - Rev New, Chief Engineer March 26, 1996. B This document supersedes the Sea Launch July 2000 Marc Nance User's Guide in its entirety - Rev A, July, Mission Design and 1998. Integration C This document supersedes the Sea Launch February Marc Nance User's Guide in its entirety - Rev B, July, 2003 Mission Design and 2000 and incorporates ADRN 01 Integration D This document supersedes the Sea Launch ______John Steinmeyer User's Guide in its entirety - Rev C, __ Director, Mission February, 2003. Integration
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ACTIVITY RECORD PAGE Section Revision Level Revision Type Numbers
1-12 Appendix D Administrative Major A, B
Appendix C D New Section
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TABLE OF CONTENTS
ABBREVIATIONS AND ACRONYMS...... xiii 1.0 INTRODUCTION...... 1-1 1.1 Overview of the Sea Launch System...... 1-1 1.2 Sea Launch Organization...... 1-5 2.0 VEHICLE DESCRIPTION...... 2-1 2.1 Zenit Stage 1...... 2-3 2.2 Zenit Stage 2...... 2-7 2.3 Block DM-SL—Upper Stage...... 2-8 2.4 Payload Unit...... 2-9 3.0 PERFORMANCE...... 3-1 3.1 Launch Site...... 3-2 3.2 Ascent Trajectory...... 3-3 3.3 Payload Capability...... 3-7 3.4 Coast Phase Attitude...... 3-21 3.5 Spacecraft Separation...... 3-21 4.0 MISSION INTEGRATION...... 4-1 4.1 Management Structure...... 4-1 4.2 Mission Analysis and Operations Planning ...... 4-3 4.3 Mission Documentation...... 4-5 5.0 SPACECRAFT ENVIRONMENTS ...... 5-1 5.1 Structural Loads...... 5-2 5.2 Random Vibration...... 5-6 5.3 Acoustics...... 5-7 5.4 Shock...... 5-10 5.5 Electromagnetic Environment ...... 5-16 5.6 Spacecraft Thermal and Humidity Environments...... 5-26 5.7 Payload Unit Ascent Venting...... 5-32 5.8 Contamination...... 5-35 6.0 SPACECRAFT DESIGN CONSIDERATIONS...... 6-1 6.1 Mass and Center-of-Gravity Limits...... 6-3 6.2 Modal Frequencies...... 6-7 6.3 Electromagnetic Compatibility ...... 6-7 6.4 Spacecraft Design Verification Requirements...... 6-10 6.5 Horizontal Handling...... 6-14 6.6 Pressurized Systems...... 6-15 6.7 Ordnance Systems...... 6-15 6.8 Multiple Manifests and Secondary Payloads...... 6-15 7.0 INTERFACES...... 7-1 7.1 Mechanical Interfaces...... 7-1 7.2 Electrical Interfaces ...... 7-11 8.0 SPACECRAFT INTEGRATION AND LAUNCH OPERATIONS ...... 8-1 8.1 Phase I⎯Spacecraft Processing in the Payload Processing Facility ...... 8-5
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8.2 Phase II⎯Mating to the Launch Vehicle and Integrated Testing on the Assembly and Command Ship...... 8-14 8.3 Phase III⎯Transit and Launch Operations on the Launch Platform...... 8-17 8.4 Flight Operations...... 8-26 9.0 HOME PORT FACILITIES AND SEA LAUNCH VESSELS ...... 9-1 9.1 Home Port Facilities ...... 9-5 9.2 Assembly and Command Ship...... 9-10 9.3 Launch Platform...... 9-21 10.0 OPERATIONS SUPPORT SYSTEMS...... 10-1 10.1 Mission Support Systems...... 10-1 10.2 Telemetry Assets...... 10-6 11.0 SAFETY...... 11-1 11.1 Safety Integration Approach...... 11-1 11.2 Safety Assessment Phases...... 11-2 11.3 Launch Licensing and Flight Safety ...... 11-7 11.4 Safety of Operations ...... 11-9 12.0 MISSION ASSURANCE...... 12-1
APPENDIX A. CUSTOMER INPUTS ...... A-1
APPENDIX B. SEA LAUNCH STANDARD OFFERINGS AND NONSTANDARD SERVICE OPTIONS...... B-1 Standard-Offering: Hardware and Software Services ...... B-1 Standard-Offering Launch Services...... B-5 Standard-Offering Facilities and Support Services ...... B-9 Nonstandard Options ...... B-15
APPENDIX C. WEATHER HISTORY AT THE LAUNCH SITE ...... C-1 Station 51028 ...... C-1 Historical Weather Data...... C-1
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LIST OF FIGURES
Figure 1-1. Generic Integration Timeline...... 1-4 Figure 1-2. Sea Launch Partner Locations ...... 1-5 Figure 1-3. Payload Fairing ...... 1-7 Figure 1-4. Block DM-SL...... 1-8 Figure 1-5. Zenit 2 Launch Vehicle...... 1-9 Figure 1-6. Launch Platform and Assembly and Command Ship...... 1-10
Figure 2-1. Zenit-3SL Launch Vehicle...... 2-1 Figure 2-2. Stage 1, Zenit-3SL ...... 2-3 Figure 2-3. RD-171M Stage 1 Engine...... 2-4 Figure 2-4. Zenit Stage 1 and Stage 2 Configuration ...... 2-6 Figure 2-5. Stage 2, Zenit-3SL ...... 2-7 Figure 2-6. Block DM-SL...... 2-8 Figure 2-7. Sea Launch Payload Unit...... 2-10 Figure 2-8. Payload Fairing ...... 2-11 Figure 2-9. Payload Unit Interface Skirt...... 2-12
Figure 3-1. Launch Site Location ...... 3-2 Figure 3-2. Typical Flight Profile—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission ...... 3-5 Figure 3-3. Typical Groundtrack—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission ...... 3-6 Figure 3-4. Geosynchronous Transfer Orbit Injection Schemes ...... 3-9 Figure 3-5. Geosynchronous Transfer Orbit Payload Capability⎯Synchronous Apogee ...... 3-10 Figure 3-6. Geosynchronous Transfer Orbit Payload Capability⎯Sub/Supersynchronous Apogees...... 3-12 Figure 3-7. Circular Orbit Payload Capability ...... 3-13 Figure 3-8. Elliptical Orbit Payload Capability...... 3-15 Figure 3-9. Molniya and Tundra Transfer Orbit Payload Capability ...... 3-17 Figure 3-10. High-Energy and Earth Escape Payload Capability ...... 3-19
Figure 4-1. Management Structure...... 4-1 Figure 4-2. Mission Documentation and Related Products...... 4-5
Figure 5-1. Transportation and Handling Limit Load Factors ...... 5-3 Figure 5-2. Integrated Launch Vehicle Horizontal...... 5-3 Figure 5-3. Integrated Launch Vehicle Vertical...... 5-3 Figure 5-4. Typical Quasi-Static Load Factors for Flight ...... 5-5 Figure 5-5. Sinusoidal Vibration at the Spacecraft Interface ...... 5-6 Figure 5-6. Random Vibration Environment During Flight...... 5-7 Figure 5-7. Acoustics During Flight⎯1/3-Octave Band Sound Pressure Levels ...... 5-9 Figure 5-8. Acoustics During Flight⎯1-Octave Band Sound Pressure Levels ...... 5-10
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Figure 5-9. Payload Shock Environment 4.0 cm Above the Separation Plane Using SCA1194VS Spacecraft Adapter, Radial Axis...... 5-12 Figure 5-10. Payload Shock Environment 4.0 cm Above the Separation Plane Using SCA1194VS Spacecraft Adapter, Axial Axis ...... 5-12 Figure 5-11. Payload Shock Environment 12.7 cm Above the Separation Plane Using SCA1666 Spacecraft Adapter...... 5-14 Figure 5-12. Payload Shock Environment 12.7 cm Above the Separation Plane Using SCA702+ and SCA702+GEM Spacecraft Adapters ...... 5-15 Figure 5-13. Typical Electromagnetic Environment in Payload Processing Facility⎯Encapsulation Cell, 15 kHz to 18 GHz...... 5-17 Figure 5-14. Typical Measured Electromagnetic Environment at the Sea Launch Home Port Pier, 15 kHz to 18 GHz ...... 5-18 Figure 5-15. Typical Measured Electromagnetic Environment at the Sea Launch Home Port Pier, 18 GHz to 40 GHz ...... 5-19 Figure 5-16. Predicted In-Flight Electromagnetic Environment at the LV/SC Interface...... 5-20 Figure 5-17. Payload Unit Air-Conditioning Venting Scheme ...... 5-28 Figure 5-18. Maximum Internal Payload Fairing Surface Temperatures...... 5-31 Figure 5-19. Payload Unit Ascent Venting Scheme...... 5-32 Figure 5-20. Payload Fairing Internal Pressure During Ascent...... 5-33 Figure 5-21. Maximum Ascent Depressurization (pressure decay rate vs. pressure) ...... 5-34
Figure 6-1. Zenit-3SL Coordinate System...... 6-2 Figure 6-2. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA1194VS...... 6-6 Figure 6-3. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA1666 Spacecraft Adapter...... 6-6 Figure 6-4. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA702+ Spacecraft Adapter...... 6-7 Figure 6-5. Maximum Spacecraft Intentional Radiated Emissions at the Launch Vehicle/Spacecraft Interface...... 6-8 Figure 6-6. Lightning and Weather Services Used to Determine or Forecast Lightning During Launch Campaign...... 6-9 Figure 6-7. Electrical and Mechanical Mates...... 6-14
Figure 7-1. Spacecraft Static Envelope ...... 7-1 Figure 7-2. Spacecraft Static Envelope SCA1194VS Adapter...... 7-2 Figure 7-3. Spacecraft Static Envelope SCA1666 Adapter...... 7-2 Figure 7-4. Spacecraft Static Envelope SCA702+ Adapter ...... 7-3 Figure 7-5. Spacecraft Envelope SCA937S Adapter...... 7-3 Figure 7-6. Payload Fairing Locations for Access Doors...... 7-5 Figure 7-7. SCA1666 Spacecraft Adapter...... 7-7 Figure 7-8. SCA702+ Spacecraft Adapter...... 7-8 Figure 7-9. SCA1194VS Spacecraft Adapter...... 7-10 Figure 7-10. SCA937S Spacecraft Adapter...... 7-11 Figure 7-11. Umbilical Configurations During Launch Processing Flow...... 7-15 Figure 7-12. Re-Radiation of Spacecraft Signal Through Fairing ...... 7-18
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Figure 7-13. Spacecraft Radio Frequency Signal Between Vessels...... 7-19 Figure 7-14. Sea Launch External Communications Network...... 7-20
Figure 8-1. Spacecraft Processing Flow ...... 8-2 Figure 8-2. Nominal 55-Day Processing Schedule...... 8-3 Figure 8-3. Nominal Campaign Timeline...... 8-4 Figure 8-4. Payload Processing Facility at Home Port...... 8-5 Figure 8-5. Fit Check on Adapter Support Stand ...... 8-6 Figure 8-6. Strongback Assembly Vertical Mating Position...... 8-8 Figure 8-7. Strongback Assembly Horizontal Raised Position ...... 8-8 Figure 8-8. Strongback Assembly Rotating to Horizontal With Pathfinder Structure...... 8-9 Figure 8-9. Encapsulated Payload Transportation Mechanical Equipment With Encapsulated Spacecraft ...... 8-10 Figure 8-10. Scaffolding Around Encapsulated Spacecraft ...... 8-11 Figure 8-11. Access to Encapsulated Spacecraft...... 8-11 Figure 8-12. Payload Unit Exiting the Payload Processing Facility ...... 8-12 Figure 8-13. Payload Unit in Transit to the Assembly and Command Ship ...... 8-13 Figure 8-14. Payload Unit on Stern Ramp of the Assembly and Command Ship...... 8-13 Figure 8-15. Payload Unit Mating to the Zenit-3SL...... 8-14 Figure 8-16. Integrated Launch Vehicle in the Assembly and Command Ship ...... 8-15 Figure 8-17. Integrated Launch Vehicle Ready for Lifting...... 8-16 Figure 8-18. Lifting of Integrated Launch Vehicle ...... 8-16 Figure 8-19. Launch Platform Sailing for the Equator...... 8-17 Figure 8-20. Assembly and Command Ship Sailing for the Equator ...... 8-18 Figure 8-21. Zenit-3SL Transporter/Erector...... 8-19 Figure 8-22. Launch Operations at the Launch Site...... 8-21 Figure 8-23. Assembly and Command Ship and Launch Platform Connected by Link Bridge...... 8-22 Figure 8-24. Liftoff from the Launch Platform ...... 8-24 Figure 8-25. Launch Control Center on the Assembly and Command Ship...... 8-25
Figure 9-1. Sea Launch Home Port Complex...... 9-2 Figure 9-2. Sea Launch Vessels...... 9-3 Figure 9-3. Home Port Location and Vicinity...... 9-4 Figure 9-4. Payload Processing Facility—Interior and Exterior Views...... 9-6 Figure 9-5. Building 3, Sea Launch Office Facility ...... 9-9 Figure 9-6. Assembly and Command Ship...... 9-11 Figure 9-7. Launch Vehicle Processing—Main Deck...... 9-12 Figure 9-8. Launch Vehicle Transfer from Assembly and Command Ship to Launch Platform at Home Port ...... 9-12 Figure 9-9. Launch Control Center—Shelter Deck...... 9-13 Figure 9-10. Launch Control Center...... 9-13 Figure 9-11. View of Launch Control Center Consoles and Front Wall (Boeing/Customer Side of Room) ...... 9-14 Figure 9-12. Customer Work Area on First Bridge Deck ...... 9-15 Figure 9-13. Assembly and Command Ship Satcom Radome ...... 9-16
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Figure 9-14. Launch Platform Satcom Radome ...... 9-17 Figure 9-15. Spacecraft Customer Accommodations – First and Second Bridge Decks ...... 9-18 Figure 9-16. Assembly and Command Ship Customer Cabin Accommodations ...... 9-18 Figure 9-17. Main Dining Hall ...... 9-19 Figure 9-18. Assembly and Command Ship Facilities ...... 9-20 Figure 9-19. Launch Platform During Final Countdown ...... 9-22 Figure 9-20. Spacecraft Contractor Work Areas on Launch Platform, 38500 Deck Level ...... 9-23
Figure 10-1. Weather Office...... 10-2 Figure 10-2. On-Site Meteorological Buoy...... 10-3 Figure 10-3. Sea Launch Bell 230 Helicopter ...... 10-4 Figure 10-4. Optical Tracker ...... 10-5 Figure 10-5. Proton Tracking Antenna...... 10-6 Figure 10-6. S-Band System Radome ...... 10-7 Figure 10-7. Nuku Hiva Ground Station ...... 10-8
Figure 12-1. Sea Launch Mission Assurance ...... 12-2 Figure 12-2. Sea Launch Hardware Certification and Acceptance Process ...... 12-4
Figure C-1. Historical Air Temperatures ...... C-2 Figure C-2. Historical Wind Gusts...... C-3 Figure C-3. Historical Atmospheric Pressure at Sea Level...... C-4 Figure C-4. Historical Water Temperatures...... C-5 Figure C-5. Historical Average Wind Speed...... C-6
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LIST OF TABLES
Table 2-1. Zenit Specifications ...... 2-5 Table 2-2. Block DM-SL Specifications...... 2-9
Table 3-1. Typical Flight Timeline—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission ...... 3-4 Table 3-2. Maximum Geosynchronous Transfer Orbit Payload Capability...... 3-8 Table 3-3. Geosynchronous Transfer Orbit Payload Capability⎯Synchronous Apogee...... 3-11 Table 3-4. Geosynchronous Transfer Orbit Payload Capability⎯ Sub/Supersynchronous Apogees...... 3-12 Table 3-5. Circular Orbit Payload Capability ...... 3-14 Table 3-6. Elliptical Orbit Payload Capability...... 3-16 Table 3-7. Molniya and Tundra Transfer Orbit Payload Capability ...... 3-18 Table 3-8. High-Energy and Earth Escape Payload Capability ...... 3-20 Table 3-9. Typical Geosynchronous Transfer Orbit Injection Accuracy (2.3σ)...... 3-22
Table 4-1. Mission Team Roles and Responsibilities ...... 4-2
Table 5-1. Spacecraft Fatigue Environment for Transportation and Handling...... 5-4 Table 5-2. Motions During Transit to Launch Site ...... 5-4 Table 5-3. Acoustics During Flight⎯1/3-Octave Band Sound Pressure Levels ...... 5-8 Table 5-4. Acoustics During Flight⎯1-Octave Band Sound Pressure Levels ...... 5-9 Table 5-5. Payload Shock Environment 4.0 cm Above the Separation Plane Using SCA1194VS Spacecraft Adapter...... 5-13 Table 5-6. Payload Shock Environment 12.7 cm Above the Separation Plane Using SCA1666 Spacecraft Adapter ...... 5-14 Table 5-7. Payload Shock Environment 12.7 cm Above the Separation Plane Using for SCA702+ and SCA702+GEM Spacecraft Adapters...... 5-16 Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies ...... 5-20 Table 5-9. Payload Processing Facility Thermal Environments ...... 5-26 Table 5-10. Payload Unit Environmental Control—Air-Conditioning Inlet Conditions ...... 5-29
Table 6-1. Expected Spacecraft Mass and Center-of-Gravity Limits ...... 6-4 Table 6-2. Factors of Safety and Test Options...... 6-10 Table 6-3. Sine Vibration Test Levels and Duration...... 6-13 Table 6-4. Spacecraft Acoustic Margins and Test Duration ...... 6-13
Table 7-1. Umbilical Configurations...... 7-13 Table 7-2. In-Flight Telemetry, Command, and Separation Breakwire Interfaces...... 7-21
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ABBREVIATIONS AND ACRONYMS
ACS assembly and command ship ADRN Advance Document Revision Notice AST U.S. Department of Transportation, Office of the Associate Administrator for Commercial Space Transportation BCSC Boeing Commercial Space Company BR9DM Zenit-3SL telemetry system C3 velocity squared at infinity CAD computer aided drawing CCAM contamination and collision avoidance maneuver CCTV closed-circuit television CDR critical design review CG center of gravity CIS Commonwealth of Independent States CLA coupled loads analysis CPSP customer prelaunch safety package DBTM Design Bureau of Transport Machinery EDC effective date of contract EGSE electrical ground support equipment EMC electromagnetic compatibility EMI electromagnetic interference EPTME encapsulated payload transportation mechanical equipment FAA Federal Aviation Administration FMH free molecular heating FPR flight performance reserve GEO geosynchronous Earth orbit GORR ground operations readiness review GOWG ground operations working group GSE ground support equipment GTO geosynchronous transfer orbit ha height of apogee hp height of perigee HP Home Port IC instrument compartment ICD interface control document I/F interface ILV integrated launch vehicle IPT integrated product team IRD interface requirements document IS interface skirt ISPS interface skirt and payload structure ITRR ILV transfer readiness review
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KNSG (automated countdown validation) K.O. kickoff LAN local area network LCC launch control center LEO low earth orbit LOX liquid oxygen LP launch platform LRR launch readiness review LSA Launch Services Agreement LSE launch support equipment LV launch vehicle MD Mission Director MEM mission equipment manager MRR mission readiness review NASA National Aeronautics and Space Administration NDBC national data buoy center NRZL non-return to zero level OASPL overall sound pressure level OSHA Occupational Safety and Health Administration PBX private branch exchange PCM pulse control modulated PLA payload accommodations PDR preliminary design review PLF payload fairing PLS payload structure PLU payload unit PPF payload processing facility PS payload structure PSM process safety management RF radio frequency RH relative humidity RMP risk management plan RR readiness review SBA strongback assembly SC spacecraft SCA spacecraft adapter SCC spacecraft contractor SCR Spacecraft Control Room SCRFSS Spacecraft Radio Frequency Support System SL Sea Launch SLS Sea Launch System STA station TDRSS Tracking Data Relay Satellite System
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TIM technical interchange meeting TM telemetry VTC video teleconferencing WX weather
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1.0 INTRODUCTION
Purpose The purpose of the Sea Launch User’s Guide is to describe, on a general basis, the Sea Launch system and associated services. This document is the starting point for understanding how the capabilities of the Sea Launch system and spacecraft integration process come together in delivering satellites to desired orbits. Further information is readily available by contacting Sea Launch directly.
1.1 Overview of the Sea Launch System
What the System The Sea Launch system includes: Includes • The Zenit-3SL launch vehicle • The Launch Platform (LP) • The Assembly and Command Ship (ACS) • Spacecraft and launch vehicle processing facilities • Tracking and communications assets Sea Launch Standard Offerings and Options (Appendix B) describes the hardware, launch services, facilities, support services, and optional services provided by the Sea Launch Company, L.L.C. The Launch Platform and the Assembly and Command Ship are located at Sea Launch Home Port, in the Port of Long Beach in Southern California. The Payload Processing Facility (PPF), which accommodates a state-of-the-art, full range of spacecraft electrical and mechanical processing, testing and fueling capabilities, is also based at Sea Launch Home Port.
REV D HPD-10041 1-1 Sea Launch’s The unique attributes of the Sea Launch system offer satellite operators a Unique Attributes multitude of benefits when considering launch vehicle providers. and Advantages Beginning with a foundation based upon the use of the robust and flight- proven Zenit-3SL integrated launch vehicle (ILV) and its associated
launch support systems, combined with the expertise and experience of our people, Sea Launch’s objective is to utilize these attributes to assure mission success for our customers. Sea Launch advantages include: • Mission Assurance and Reliability • Led by Boeing, the Sea Launch Mission Assurance organization represents the blending of the best practices of U.S. companies, European companies, and companies from Russia and Ukraine • Utilizes an integrated set of systems and processes that ensures mission success and ever-increasing product reliability • The demonstrated Zenit-3SL flight-history is comparable to the new-generation of heavy-lift vehicles • Built-in Schedule Assurance • The use of a dedicated Zenit-3SL Launch Vehicle that is free of co-passenger constraints • Launching from an equatorial launch site with no significant weather-related restrictions • Optimized Performance • Sea Launch provides an optimized mission design, based on full vehicle capability to ensure maximum performance to GTO • Lift capability exceeding 6,000 kilograms to GTO • Industry’s most direct route to GTO • From our equatorial launch site, Sea Launch offers simplified mission profiles with shorter mission durations, fewer staging events and correspondingly less risk • Increased on-orbit maneuvering lifetime (OML) or greater spacecraft mass to orbit when compared to other, non-equatorial, launch sites • Experienced Personnel • Extensive integration experience with all major spacecraft platforms affords Sea Launch’s customers strong confidence for mission success. • A staff and management team consisting of highly knowledgeable technical and managerial personnel with long-term industry experience. • Direct interaction with English speaking technical specialists facilitates efficient integration and verification of compliance to spacecraft requirements.
1-2 HPD-10041 REV D Mission Timeline The first launch of a new spacecraft type (i.e., a spacecraft never integrated by Sea Launch) can be accomplished within eighteen months after contract signature. Repeat missions of spacecraft previously integrated by Sea Launch can be completed within twelve months after contract signature. Sea Launch is capable of reducing the time required for integrating and launching spacecraft. Deviations from the standard eighteen-month and twelve-month flows may be accommodated on a case-by-case basis, recognizing factors such as, but not limited to, previous flight experience. Some generic timelines for missions are shown in Figure 1-1. For a two- cycle mission, a Preliminary Design Review (PDR) and a subsequent Critical Design Review (CDR) are necessary to demonstrate compliance with spacecraft requirements. Both reviews are typically conducted near the end of the Preliminary and Final Design Cycles. For a single-cycle mission, only a CDR is conducted. For more information on these reviews, refer to Section 4. Similarly, the standard mission cycle time is fifty-five days between launches. Sea Launch has committed to reducing launch cycle time with a commensurate reduction in spacecraft processing time. Companies considering the Sea Launch system should review the initial information that will be required by Sea Launch (Appendix A) and be prepared to deliver that information as soon as possible after contract signature. Sea Launch will use this information to begin the integration process.
REV D HPD-10041 1-3 Months L-21 L-20 L-19 L-18 L-17 L-16 L-15 L-14 L-13 L-12 L-11 L-10 L-9 L-8 L-7 L-6 L-5 L-4 L-3 L-2 L-1
EDC/Cust K.O. Customer Data 18-month mission from EDC/ Delivery Preliminary Design Cycle 15-month integration plan Final Design Cycle Flight Task/Simulation Campaign
EDC/Cust K.O. 15-month mission from EDC/ Customer Data Delivery 12-month integration plan Final Design Cycle Flight Task/Simulation Campaign
EDC/Cust K.O. Customer Data Delivery 10.5-month mission from EDC/ Final Design Cycle 9-month integration plan Flight Task/Simulation Campaign
EDC/Cust K.O. Customer Data Delivery 8-month mission from EDC/ Final Design Cycle 7-month integration plan Flight Task/Simulation Campaign
SHG0681874-073.1 Figure 1-1. Generic Integration Timeline
1-4 HPD-10041 REV D 1.2 Sea Launch Organization
Overview Sea Launch Limited Partnership is an exempt Cayman Islands entity that acts through its sole general partner, Sea Launch Company L.L.C. which is a Delaware limited liability company. The limited partners of Sea Launch Limited Partnership consist of: • Boeing Commercial Space Company (BCSC) • RSC Energia • SDO Yuzhnoye/PO Yuzhmash • Aker Partner locations and operating centers are shown in Figure 1-2. Each partner is also a supplier/contractor to the company, capitalizing on the strengths of these industry leaders. The Boeing, Energia, Yuzhnoye/Yuzhmash, and Aker team of contractors has developed an innovative and successfully proven approach to establishing Sea Launch as a reliable, cost-effective and flexible commercial launch service provider.
Aker, Oslo RCS Energia, Moscow BCSC, Seattle Boeing SDO Yuzhnoye/ Corporate PO Yuzhmash, Sea Launch, Dnepropetrovsk Long Beach
Launch Site
SHG0681874-017 Figure 1-2. Sea Launch Partner Locations
REV D HPD-10041 1-5 Sea Launch The Sea Launch Company, headquartered in Long Beach, California, is Company populated by selected representatives of the partner companies. Charged with managing the overall program, Sea Launch responsibilities include: • Overall operations management • Business management • Supplier management • Marketing and sales • Customer relations • Communications • Insurance and financing • Chief Systems Engineering Office • Mission Management Sea Launch is organized to ensure mission success by having segment managers responsible for the hardware acquisition, acceptance, and operational support of each program element.
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The Boeing The Boeing Company brings expertise in space systems integration, a Company reputation for product integrity and dedication to customer service. Boeing responsibilities include: • Designing and manufacturing the payload fairing (Figure 1-3) and adapter • Managing Home Port operations • Integrating the spacecraft with the payload accommodations and the Sea Launch vehicle • Performing mission analysis and analytical integration • Leading mission operations • Securing launch licensing • Providing range services
265807J6-251/R1 Figure 1-3. Payload Fairing
REV D HPD-10041 1-7 RSC Energia RSC Energia, the premier Russian space company, brings its well-known experience in space exploration and launch system integration to the Sea Launch System. Energia is a joint stock company established under the laws of the Russian Federation, with its principal place of business in Korolev (six kilometers northeast of Moscow), Russia. Energia has overall responsibility for: • Manufacturing the Block DM-SL upper stage (Figure 1-4) • Developing and operating the automated ground support equipment • Integrating the Block DM-SL with the Zenit 2S and launch support equipment • Managing the launch support equipment on board both the Launch Platform and the Assembly and Command Ship • Developing flight design documentation for the flight of the upper stage • Performing launch operations and range services
265807J6-252 Figure 1-4. Block DM-SL
1-8 HPD-10041 REV D SDO SDO Yuzhnoye and PO Yuzhmash, the leading Ukrainian aerospace Yuzhnoye/PO companies, are established under the laws of Ukraine, with their Yuzhmash principal place of business in Dnepropetrovsk, Ukraine. They are also contractors to the Sea Launch Company, L.L.C. As a principal designer of ballistic missiles and launch vehicles, they have conducted hundreds of successful launches over the past four decades. SDO Yuzhnoye designs and PO Yuzhmash manufactures the Zenit two- stage launch vehicle derived from the heritage Zenit 2 rocket illustrated in Figure 1-5, which comprises the first two stages of the Sea Launch vehicle. Yuzhnoye/Yuzhmash are responsible for: • Producing the first and second stages of the Zenit 2S launch vehicle • Integrating the launch vehicle flight hardware • Developing flight design documentation for launch, with respect to the first two stages • Supporting Zenit processing and launch operations Several significant configuration modifications were made to allow the basic Zenit design to meet Sea Launch’s unique requirements. For example, the base of the launch vehicle was stiffened to provide greater structural strength required on a dynamic launch platform.
265807J6-253 Figure 1-5. Zenit 2 Launch Vehicle
REV D HPD-10041 1-9 Aker Aker, based in Oslo, Norway, is one of Europe’s largest ship builders, with extensive experience in developing offshore oil platforms designed for operations in the harsh North Sea environment. During development, Aker (formerly Kvaerner) was responsible for: • Designing and building the Assembly and Command Ship (Figure 1- 6) • Designing and reconstructing the Launch Platform • Integrating the marine elements Barber Moss Ship Management is responsible for marine operations and maintenance of both vessels.
SHG0681874-001 Figure 1-6. Launch Platform and Assembly and Command Ship
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2.0 VEHICLE DESCRIPTION
Overview The Zenit-3SL is a liquid propellant launch vehicle capable of transporting spacecraft to a variety of orbits. Figure 2-1 shows the Zenit- 3SL principal components: • Zenit Stage 1 • Zenit Stage 2 • Block DM-SL upper stage • Payload unit
Payload Block Stage 2 Stage 1 Unit DM-SL 10.4 m 32.9 m 11.39 m 4.9 m (34.0 ft) (108.0 ft) (37.4 ft) (16.1 ft)
4.15 m 3.7 m 3.9 m 3.9 m (13.6 ft) (12.1 ft) (12.8 ft) (12.8 ft) 2P341004.1 Figure 2-1. Zenit-3SL Launch Vehicle
Design The original two-stage Zenit was designed by SDO Yuzhnoye to reconstitute military satellite constellations quickly. The design emphasizes robustness, ease of operation, and fast reaction times. The result is a highly automated launch capability using a minimum complement of launch personnel.
Flight History The Zenit launch vehicle was first launched in 1985 from Baikonur Cosmodrome in Kazakhstan. As of the time of publishing, the Zenit 2 and Zenit 2SLB vehicles have successfully completed 32 missions in 37 launch attempts. The Zenit first-stage booster also served as a strap-on for the Energia launch vehicle and logged an additional eight successes in two flights in this capacity. The Block DM-SL is a modification of the existing Block DM developed by RSC Energia. It has a long and successful history as the fourth stage of the Proton launch vehicle.
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With the addition of an upgraded avionics and guidance suite, the Block DM superseded the Block D upper stage in 1974 and has since flown 239 successful missions as of the time of publishing. Primary missions have been geosynchronous transfer orbit; geosynchronous Earth orbit; and interplanetary, including missions to Halley’s Comet, Venus and Mars. Significant configuration differences between the Zenit-3SL and its predecessor are: • New navigation system • Next-generation flight computer • Stiffened first stage • Fueling lines modified for third stage In addition, several enhancements have been made to the Zenit-3SL to upgrade the performance capability from 5,250 kilograms to over 6,000 kilograms to GTO. These changes fall into three categories: • Energy increase (includes Block DM-SL engine extended nozzle, second-stage main engine thrust increase from 87 tons force to 93 tons force, and the addition of propellant for all three stages) • Mass reductions across all three stages and payload accommodations • Optimization of the propellant usage in all three stages The current flight history of the Zenit-3SL is shown on the Sea Launch web page at http://www.sea-launch.com/history.htm.
Payload Payload accommodations were developed by The Boeing Company Accommodations specifically for the Sea Launch Company. The design is based on launch experience that extends back over three decades. The payload accommodations design protects the spacecraft from transportation, handling, launch and flight environments. The design features: • Streamlined operational flows • Flexibility to accommodate varied spacecraft models
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2.1 Zenit Stage 1
Stage 1 Primary The Zenit Stage 1 primary structure is aluminum with integrally Structure machined stiffeners (Figure 2-2). It is based on the same engine and core fuselage that was also used for the four strap-on boosters for the Russian Energia heavy-lift Buran launch vehicle. There have been two such Energia launches, both successful.
265807J6-231 Figure 2-2. Stage 1, Zenit-3SL
RD-171M Engine The RD-171M engine, which powers Stage 1, burns liquid oxygen and kerosene (Figure 2-3). It provides an impressive 740,000 kgf (1.6 million lbf) of thrust at liftoff and is one of the most powerful engines in the world. The four thrust chambers are fed by a single, vertically mounted turbopump, powered by two gas generators feeding hot oxidizer-rich gas to a single turbine. Four differentially gimbaled thrust nozzles provide directional control during Stage 1 flight.
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SHG0681874-019 Figure 2-3. RD-171M Stage 1 Engine
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Overall Zenit specifications and performance parameters are shown in Table 2-1. Specifications and Stage 1 and Stage 2 configurations are shown in Figure 2-4. Configurations Table 2-1. Zenit Specifications Zenit Stage 1 Stage 2 Length 32.9 m (108 ft) 10.4 m (34 ft) Diameter 3.9 m (12.8 ft) 3.9 m (12.8 ft) Weight (fueled) 354,582 kg 90,757 kg (781,719 lb) (200,085 lb) Thrust (sea level) 740,000 kgf Not applicable (1.63 million lbf) Thrust (vacuum) 806,400 kgf main - 93,000 kgf (1.78 million lbf) (205,026 lbf); vernier - 8,100 kgf (17,900 lbf) Fuel (kerosene) 88,768 kg (195,700 22,832 kg lb) (50,336 lb) Oxidizer (LOX) 233,512 kg (514,805 58,908 kg lb) (129,869 lb)
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Block DM-SL interface plane
Liquid oxygen tank
Kerosene tank Stage 2 Solid-propellant 10.4 m (34 ft) separation thrusters (4) Interstage frame Separation plane
Electric/pneumatic connector plate Main Steering thruster engine 3.9 m nozzle (12.8 ft)
Liquid oxygen tank
Stage 1 Kerosene tank 32.9 m (108 ft)
Solid- propellant separation thrusters (4)
Single turbopump Solid- Main propellant engine separation View A-A nozzle (4) thrusters (4)
A A 265807J6-257 Figure 2-4. Zenit Stage 1 and Stage 2 Configuration
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2.2 Zenit Stage 2
Stage 2 Primary The Zenit second stage also uses integral stiffened aluminum Structure construction (Figure 2-5). Stage 2 propellants are liquid oxygen and kerosene. The lower kerosene tank is toroid-shaped, and the liquid oxygen tank is a domed cylinder. The stage is powered by a single- nozzle RD-120 engine, which produces 93,000 kgf (205,026 lbf) of thrust in flight.
265807J6-258 Figure 2-5. Stage 2, Zenit-3SL
Three-Axis The RD-8 vernier engine mounted in the aft end of Stage 2 provides Control three-axis attitude control. The RD-8 uses the same propellants as the single fixed-nozzle RD-120, with one turbopump feeding four gimbaling thrusters. The RD-8 produces 8100 kgf (17,900 lbf) of thrust in flight.
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2.3 Block DM-SL—Upper Stage
Configuration The Block DM-SL is a restartable upper stage capable of igniting up to five times during a mission. Avionics are housed in a sealed toroidal equipment bay at the front end of the upper stage. An interstage cylinder of aluminum skin and stringer construction encloses the Block DM-SL. The configuration of the Block DM-SL is shown in Figure 2-6.
Propulsion Propulsive capability for the upper stage is provided by the 11D58M engine, which operates on a combination of liquid oxygen and kerosene. The kerosene is contained in a toroidal tank that encircles the main engine turbopump. The spherical liquid oxygen tank is located above the fuel tank. The 11D58M has a single gimbaling nozzle that provides directional control during propulsive phases.
PLF Avionics/ interface equipment plane bay Middle adapter separation LOX tank plane
Upper adapter
ø3.7 m (ø 12.1 ft) 5.6 m (18.4 ft) Kerosene tank Middle adapter (interstage) 11D58M engine Stage II separation Attitude plane control Lower engines adapter Block DM-SL (without interstage) (two places)
ø 3.9 m (ø 12.8 ft) SHG0681874-020 Figure 2-6. Block DM-SL
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Attitude Control Three-axis stabilization of the Block DM-SL during coast periods is provided by two attitude control/ullage engines. Each engine has four nozzles that are grouped in clusters on either side of the main engine nozzle. The attitude control system uses the hypergolic propellant nitrogen tetroxide (N2O4) and unsymmetrical dimethylhydrazine (UDMH). Overall specifications for the Block DM-SL are shown in Table 2-2. Table 2-2. Block DM-SL Specifications Length1 5.6 m (18.4 ft) Diameter (primary) 3.7 m (12.1 ft) Weight (fueled)2 19,711 kg (43,455 lb) Thrust (vacuum) 8103 kg (17,864 lb) Fuel (kerosene) 4560 kg (10,053 lb) Oxidizer (LOX) 11,290 kg (24,890 lb) Note 1: The bottom of the payload unit interface skirt overlays the top of the Block DM-SL by 1.1 m (3.6 ft). Note 2: Includes the Block DM-SL lower and middle adapters.
2.4 Payload Unit
Components and The payload unit consists of the spacecraft integrated with the payload Construction accommodations. The payload accommodations consist of the spacecraft adapter, payload structure, interface skirt, payload fairing, and flight avionics and instrumentation. Figure 2-7 shows the arrangement of the major payload accommodations components. The construction is aluminum honeycomb and graphite/epoxy for the fairing, nosecap and payload support structure. The payload fairing is 11.39 m (37.4 ft) long and 4.15 m (13.6 ft) in diameter. These elements are integrated in a class 100,000 clean facility during ground processing in the PPF at Home Port. The standard spacecraft adapters are procured from the industry standard supplier, Saab Space AB.
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Spacecraft adapter
Payload structure
Interface skirt
Nosecap (composite)
Payload fairing (graphite/epoxy composite)
Longitudinal separation system
11.39 m 3.9 m (37.4 ft) (12.8 ft)
4.15 m (13.6 ft)
Spacecraft I/F plane Spacecraft adapter (aluminum) Payload support structure Closeout membrane
Block DM-SL I/F plane Circumferential separation system
Interface skirt (aluminum) Block DM-SL (partial view) SHG0681874-077- Figure 2-7. Sea Launch Payload Unit
Payload Fairing The payload fairing (Figure 2-8) provides environmental protection for spacecraft from encapsulation through launch and ascent. It can accommodate a wide range of payload designs.
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265807J6-261 Figure 2-8. Payload Fairing
Access Access to the spacecraft before rollout to the launch pad can be gained Characteristics through doors in the payload fairing. The standard offering includes two payload fairing access doors approximately 610 mm (2 ft) in diameter. These doors are located on opposite sides of the payload fairing longitudinal separation plane and at least 17 deg from the separation plane. Within certain payload fairing structural constraints, variations in the number, location, and size of the doors can be accommodated (Figure 7- 5).
Cooling Air From encapsulation to launch, conditioned air is provided to the payload fairing volume. The cooling air flows from an entry on the side of the payload fairing to the aft end, where it exits through one-way valves on the payload structure. Conditioned air capabilities are described in detail in Section 5.
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Thermal External thermal insulation protects the payload fairing structure and Insulation limits the interior payload fairing surface temperatures. Payload fairing jettison is constrained to a time sufficient to ensure that the maximum dispersed free molecular heating never exceeds spacecraft requirements. The customer can tailor the time of payload fairing jettison (and associated maximum free molecular heating). The thermal environments are described in Section 5.
Interface Skirt The interface skirt joins the payload fairing to the Block DM-SL and also supports the spacecraft loads through the spacecraft adapter and payload structure. The interface skirt is constructed of aluminum with integral stiffeners (Figure 2-9). The interface skirt is 0.81 m (2.6 ft) long and accommodates the transition from a 3.7 m (12.1 ft) diameter on the Block DM-SL to a 4.15 m (13.6 ft) diameter on the payload fairing.
265807J6-262 Figure 2-9. Payload Unit Interface Skirt
Spacecraft The spacecraft adapter mechanical interface with the spacecraft can be Adapters either a bolted or a typical clampband (Marmon-type) interface. The specific spacecraft interface information for the bolted and clampband based adapters is provided in Section 7. Spacecraft separation is achieved by redundant initiation, which causes the separation of the retaining bolts or clampbands.
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Payload Structure The payload structure acts as a physical support for the spacecraft during horizontal and vertical operations and functions as an environmental barrier. It is a graphite/epoxy composite structure. One-way valves in the payload structure allow airflow out of the payload unit and prevent backflow of air from the Block DM-SL cavity. The closeout membrane, located at the upper diameter of the payload structure, provides a barrier that limits thermal and contamination exchange between the Block DM-SL and the spacecraft.
Future 5-m Sea Launch is assessing customer interest in a 5-m diameter payload Capability unit. If sufficient interest exists, Sea Launch will consider offering this capability to our customers. The Sea Launch 5-m design would provide a payload static envelope 4.57 m in diameter. The design will provide encapsulated operations similar to our baseline payload unit operations, as well as similar avionics and structural interfaces.
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3.0 PERFORMANCE
Overview The Zenit-3SL can deliver spacecraft to a wide variety of orbits, includ- ing geosynchronous transfer, medium Earth, highly elliptical, and Earth escape orbits. The data presented in this section assumes a generic set of spacecraft requirements. Please contact Sea Launch for a performance quote for specific mission requirements. With a standard flight performance reserve, the maximum geosynchro- nous transfer orbit (GTO) payload capability with one or two Block DM- SL main engine burns is 6160 kg and 6090 kg, respectively. A reduced flight performance reserve option is also available, which provides 55 kg of additional payload capability and increased dispersions on the orbital parameters. Characteristics of performance are covered in Sections 3.1 through 3.5, including: • Launch site • Ascent trajectory • Payload capability • Coast phase attitude • Spacecraft separation
Trajectory Design Trajectory design is part of the overall mission analysis process de- Process scribed in Section 4. Boeing Commercial Space Company (BCSC) flight design engineers work with each spacecraft customer to optimize the mission design and coordinate all customer requirements with the Rus- sian and Ukrainian partners. The Sea Launch Russian and Ukrainian partners generate the final trajectory and perform software and hardware simulations to verify the mission design. Finally, BCSC verifies that all customer requirements are satisfied.
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3.1 Launch Site
Site Location Figure 3-1 shows the location of the Sea Launch launch site. Located on the equator at 154° W in the Pacific Ocean, this site was chosen based on a number of factors, including stage impact zones, weather, and vessel transit times. Sea Launch may select an alternate launch site for inclined missions considering range safety factors (e.g., a population center lo- cated along the ascent groundtrack).
30° N
20° N
10° N Launch Site 0° N, 154° W
0° N
10° S
160° W 140° W 120° W 100° W 80° W SHG0681874-021
Figure 3-1. Launch Site Location
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Performance The equatorial launch site enables the launch vehicle to efficiently de- Benefits liver a geostationary spacecraft to a zero degree orbital inclination. This saves the spacecraft from having to perform a plane change maneuver, typically resulting in a longer on-orbit maneuvering lifetime (OML). For a typical geostationary spacecraft, the circularized GEO spacecraft mass is 10% higher when launched from the equator compared to a launch originating from Cape Canaveral. For missions requiring non-zero inclinations, a wide variety of orbital inclinations are possible, due to the lack of population centers downrange of the launch site.
3.2 Ascent Trajectory
Stage 1 Flight The Zenit Stage 1 engine provides the thrust for the first 2.5 min of flight. A 10-second vertical rise is followed by a roll to the appropriate mission launch azimuth. The pitch profile is designed to minimize the lateral loads during the periods of highest dynamic pressure. When the axial acceleration reaches 4.0 g’s, the engine is gradually throttled to 50%, which is then maintained until Stage 1 separation.
Stage 2 Flight The Zenit Stage 2 main and vernier engines provide the thrust from 2.5 to 8.5 min after liftoff. The vernier engine ignites prior to Stage 1 sepa- ration and Stage 2 main engine ignition. The payload fairing is jetti- soned based on a free molecular heating constraint, and typically occurs three to four minutes after liftoff. Prior to Stage 2 main engine shut- down, the engine is throttled to 80%. After main engine shutdown, the vernier engine operates for another 1.3 min until Stage 2 separation.
Block DM-SL Immediately after Stage 2 separation, the middle adapter surrounding the Flight Block DM-SL is jettisoned. Block DM-SL main engine ignition occurs ten seconds later. Depending on the mission requirements, the Block DM-SL performs one or more main engine burns. Most missions utilize two burns of the Block DM-SL, as this usually provides a higher perigee, and may also be used to target additional orbital parameters (e.g, apogee longitude, argu- ment of perigee). For missions with multiple burns, the initial burn es- tablishes a stable parking orbit. The parking orbit coast duration is at least thirty minutes to allow sufficient cooling prior to engine restart.
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Flight Timeline Table 3-1 lists the times when the main flight events occur for a two- burn 6090-kg geosynchronous transfer orbit mission.
Table 3-1. Typical Flight Timeline—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission Time, min:sec Event 0:00 Stage 1 engine ignition 0:04 Liftoff 0:11 Begin pitchover 1:06 Maximum dynamic pressure 1:56 Maximum axial acceleration 2:15 Stage 1 engine throttled to 50% 2:25 Stage 2 vernier engine ignition 2:27 Stage 1 engine shutdown 2:30 Stage 1 separation 2:33 Stage 2 main engine ignition 3:51 Payload fairing jettison 7:13 Stage 2 main engine shutdown 8:31 Stage 2 vernier engine shutdown 8:31 Stage 2 separation 8:32 Block DM-SL middle adapter jettison 8:41 Block DM-SL main engine ignition 1 13:12 Block DM-SL main engine shutdown 1 43:30 Block DM-SL main engine ignition 2 50:35 Block DM-SL main engine shutdown 2 61:15 Spacecraft separation
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Flight Profile Figure 3-2 shows a flight profile with key trajectory events and parameters labeled for a two-burn, 6090-kg geosynchronous transfer orbit mission.
Apogee • t = 6:03:08 • h = 35786 km • hp = 219 km • ha = 35786 km Block DM-SL shutdown 1 Stage 2 separation • t = 13:12 Spacecraft separation • t = 8:31 • h = 181 km • h = 180 km • t = 1:01:15 • hp = 180 km • h = 2095 km • hp = -2266 km Fairing jettison • ha = 207 km • h = 183 km • t = 3:51 a • h = 118 km Block DM-SL shutdown 2 • Dispersed FMH < 1,135 W/m2 • t = 50:35 • h = 334 km • hp = 214 km Stage 1 separation • ha = 35963 km • t = 2:30 • h = 69 km
Max acceleration • t = 1:56 • Accel = 4.0 g Stage 2 impact Max Q • Range = 4490 km • t = 1:06 • Q = 5400 kgf/m2 Fairing impact Liftoff • Range = 1050 km • t = 0:04 • Accel = 1.6 g Stage 1 impact • Range = 770 km
SHG0681874-022 Figure 3-2. Typical Flight Profile—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission
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Groundtrack Figure 3-3 shows the Zenit-3SL ascent groundtrack and indicates the location of the major events for a two-burn, 6090-kg geosynchronous transfer orbit mission.
60 N
30 N
0 Stage 1 & 2 Block DM-SL Block DM-SL Spacecraft Apogee Ascent Burn 1 burn 2 separation Latitude, deg Latitude,
30 S
60 S
150 W 120 W 90 W 60 W 30 W 0 30 E 60 E 90 E 120 E 150 E Longitude, deg SHG0681874-023 Figure 3-3. Typical Groundtrack—Two-Burn, 6090-kg Geosynchronous Transfer Orbit Mission
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3.3 Payload Capability
Ground Rules This section presents the payload capability for a wide variety of mis- sions, based on the following set of generic ground rules: • Payload capability is defined as the separated spacecraft mass, and assumes an 1194, 1666, or 937 spacecraft adapter. For a 702 adapter, subtract 15 kg from the payload capability values in this section. • Sufficient flight performance reserves are maintained to reach the specified orbit with a confidence level of 2.7σ. • At the time of fairing jettison, the free molecular heating is less than 1,135 W/m2 with a confidence level of 3σ. • Orbital heights are specified with respect to an Earth radius of 6,378 km. • For elliptical orbits (including GTO and Molniya/Tundra transfer), the orbital parameters correspond to the osculating orbital elements at first apogee passage. • For circular orbits and Earth escape, the orbital parameters corre- spond to the osculating orbital elements at Block DM-SL main en- gine shutdown. • The maximum separated spacecraft mass is assumed to be 7300 kg based upon the Zenit-3SL structural capability specified in Section 6.1. Refer to this section for more details on the structural capability of the Zenit-3SL using current SC adapters and associated payload support structure. • For missions with multiple Block DM-SL engine burns, the mini- mum parking orbit perigee height is 180 km. For missions with a sin- gle Block DM-SL burn, the minimum final orbit perigee height is 200 km. • Missions requiring longer durations (over 2 hr at spacecraft separa- tion) or higher altitudes (over 15,000 km at spacecraft separation) may require additional analyses due to battery life and radiation con- siderations.
Reduced Flight All of the data in this section corresponds to a standard flight perform- Performance ance reserve on the Block DM-SL. A reduced flight performance reserve Reserve Option option is also available, which provides 55 kg of additional payload ca- pability and increased dispersions on the orbital parameters. Please con- tact Sea Launch for more detailed information.
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Geosynchronous The Zenit-3SL can deliver large satellites to a variety of geosynchronous Transfer Orbit transfer orbits. For a mission with one Block DM-SL burn, the maxi- mum payload capability is 6160 kg. For a mission with two Block DM- SL burns, the maximum payload capability is 6090 kg. The orbital pa- rameters for these cases are shown in Table 3-2.
Table 3-2. Maximum Geosynchronous Transfer Orbit Payload Capability One Block Two Block Orbital parameter DM-SL burn DM-SL burns Perigee height 200 km 220 km Apogee height 35786 km 35786 km Inclination 0 deg 0 deg Maximum 6160 kg 6090 kg spacecraft mass For lighter spacecraft, the excess Zenit-3SL performance may be used to raise the perigee height. Either one or two Block DM-SL burns may be used but, for most cases, two burns provide the maximum perigee. A one-burn mission is optimal only for the heaviest spacecraft, and may be selected if other spacecraft requirements do not necessitate the use of two burns.
For two-burn missions, there are two types of GTO injection schemes: “perigee inject” and “post-perigee inject.” These have different intermedi- ate parking orbits and placement of the second burn, which are illustrated in Figure 3-4. Perigee inject is optimal for the lower perigees and post- perigee inject is optimal for the higher perigees, with the crossover occur- ring near a perigee height of 2000 km. Injection at apogee is generally not desirable due to a minimum propellant constraint at the start of a Block DM-SL main engine burn, plus it greatly extends the mission duration.
Figure 3-5 and Table 3-3 show the geosynchronous transfer orbit payload capability for a synchronous apogee. Figure 3-6 and Table 3-4 show the geosynchronous transfer orbit payload capability for subsynchronous and supersynchronous apogees.
All of the geosynchronous transfer orbit data in this section assumes an orbital inclination of 0 deg. A small non-zero inclination can be accom- modated with minimal performance loss. Additional constraints (e.g., apogee longitude, argument of perigee, right ascension of the ascending node) can also be accommodated, but the performance loss depends on the specifics of the mission.
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One Block DM-SL Burn Ascent + DM burn Hp > 200 km
GTO
Two Burn Perigee Inject DM burn 2 Hp = 220 - 2000 km Ascent + DM burn 1
GTO
Ascent + DM burn 1
Two Burn Post-Perigee Inject Hp > 2000 km
DM burn 2 GTO
SHG0681874-024 Figure 3-4. Geosynchronous Transfer Orbit Injection Schemes
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Apogee height = 35786 km Inclination = 0 deg 6000
5000 Two Block DM-SL Burns Payload capability, kg capability, Payload 4000 One Block DM-SL Burn
2000 4000 6000 8000 Perigee height, km SHG0681874-025 Figure 3-5. Geosynchronous Transfer Orbit Payload Capability⎯Synchronous Apogee
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Table 3-3. Geosynchronous Transfer Orbit Payload Capability⎯Synchronous Apogee
Perigee Spacecraft Payload capability, kg height, ΔV to GEO, One Block Two Block km mps DM-SL burn DM-SL burns 200 1477 6160 ⎯ 220 1475 6123 6090 250 1472 6055 6072 300 1466 5940 6033 400 1456 5704 5953 500 1446 5463 5873 600 1436 5218 5795 700 1426 4970 5720 800 1416 4718 5645 900 1406 4463 5574 1000 1397 4206 5504 1250 1373 3568 5339 1500 1349 2957 5186 1750 1327 2370 5048 2000 1304 ⎯ 4921 2100 1296 ⎯ 4887 2250 1283 ⎯ 4869 2500 1261 — 4795 3000 1220 ⎯ 4614 3500 1180 ⎯ 4455 4000 1142 ⎯ 4314 4500 1106 ⎯ 4187 5000 1070 ⎯ 4072 6000 1003 ⎯ 3873 7000 941 ⎯ 3704 8000 882 ⎯ 3559 9000 827 ⎯ 3431 10000 775 ⎯ 3319
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Inclination = 0 deg Two Block DM-SL Burns Perigee height, km
220 6000 500
1000 5500 Payload capability, kg capability, Payload
5000 2000
32000 34000 36000 38000 40000 Apogee height, km SHG0681874-026
Figure 3-6. Geosynchronous Transfer Orbit Payload Capability⎯Sub/Supersynchronous Apogees
Table 3-4. Geosynchronous Transfer Orbit Payload Capability⎯ Sub/Supersynchronous Apogees* Payload capability, kg Apogee Perigee ht Perigee ht Perigee ht Perigee ht height, km 220 km 500 km 1000 km 2000 km 30000 6414 6192 5818 5227
32000 6292 6071 5699 5111
34000 6182 5962 5592 5006
35786 6090 5873 5504 4921
38000 5991 5773 5406 4825
40000 5907 5690 5324 4746
42000 5831 5615 5250 4673 *Two Block DM-SL burns.
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Circular and The Zenit-3SL can deliver satellites to a variety of circular and elliptical Elliptical Orbits orbits. Figure 3-7 and Table 3-5 show the payload capability for a range of circular orbit heights and inclinations. Figure 3-8 and Table 3-6 show the payload capability for a range of elliptical orbit apogees and inclinations.
Two Block DM-SL Burns 7000
6000 Orbital inclination, deg
0
5000 45
90 4000 Payload capability, kg capability, Payload
135 3000
5000 10000 15000 Circular orbit height, km SHG0681874-027 Figure 3-7. Circular Orbit Payload Capability
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Table 3-5. Circular Orbit Payload Capability* Payload capability, kg Height, Inclination, Inclination, Inclination, Inclination, km 0 deg 45 deg 90 deg 135 deg 1000 7300 7300 7300 5466 2000 7300 7300 6617 4712 3000 7300 7300 5851 4061 4000 7300 6990 5203 3510 5000 6949 6334 4656 3045 6000 6347 5776 4192 2651 7000 5829 5297 3795 2315 8000 5377 4882 3455 2027 10000 4631 4203 2903 1561 12000 4041 3675 2480 1206 14000 3565 3255 2147 ⎯ 16000 3173 2914 1882 ⎯ 18000 2846 2634 1668 ⎯ 20000 2568 2398 1494 ⎯ 22000 2329 2197 1350 ⎯ 24000 2123 2024 1231 ⎯ 26000 1942 1873 1132 ⎯ 28000 1782 1738 1048 ⎯ 30000 1641 1616 ⎯ ⎯ *Two Block DM-SL burns.
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Perigee height = 200 km 7000 One Block DM-SL Burn
Orbital inclination, deg 0 6000 45
5000 90
Payload capability, kg capability, Payload 135 4000
3000
10000 20000 30000 40000 50000 60000 Apogee height, km SHG0681874-028
Figure 3-8. Elliptical Orbit Payload Capability
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Table 3-6. Elliptical Orbit Payload Capability*
Apogee Payload capability, kg height, Inclination, Inclination, Inclination, Inclination, km 0 deg 45 deg 90 deg 135 deg 10000 7300 7300 6854 5552 12500 7300 7300 6341 5102 15000 7300 7300 5944 4756 17500 7300 7022 5627 4481 20000 7300 6716 5367 4258 22500 7085 6462 5152 4072 25000 6854 6248 4970 3915 27500 6657 6064 4814 3781 30000 6485 5906 4679 3665 32500 6333 5767 4562 3563 35000 6200 5646 4458 3474 37500 6080 5538 4366 3395 40000 5974 5441 4284 3324 42500 5878 5354 4211 3261 45000 5791 5275 4144 3203 47500 5712 5203 4084 3151 50000 5640 5137 4029 3104 52500 5574 5076 3978 3060 55000 5513 5021 3932 3020 57500 5457 4969 3889 2983 60000 5406 4922 3850 2949 65000 5314 4836 3779 2888 70000 5233 4762 3718 2835 *One Block DM-SL burn, perigee height 200 km.
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Molniya and Figure 3-9 and Table 3-7 show the Zenit-3SL payload capability for Tundra Transfer transfer to 12-hr Molniya and 24-hr Tundra orbits. All cases have an or- Orbits bital inclination of 63.4 deg and argument of perigee of 270 deg.
Inclination = 63.4 deg Argument of Perigee = 270 deg 5000 Apogee height, km Two Block DM-SL Burns
35000 4500 40000
45000 4000 50000
3500 Payload capability,kg 3000
2500
1000 2000 3000 4000 5000 Perigee height, km SHG0681874-074 Figure 3-9. Molniya and Tundra Transfer Orbit Payload Capability
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Table 3-7. Molniya and Tundra Transfer Orbit Payload Capability*
Perigee Payload capability, kg height, Apogee ht, Apogee ht, Apogee ht, Apogee ht, km 35000 km 40000 km 45000 km 50000 km 213 5135 4940 4782 4652 300 5089 4894 4737 4608 400 5020 4825 4668 4539 500 4950 4756 4600 4471 600 4882 4689 4533 4405 700 4817 4624 4469 4341 800 4755 4563 4407 4280 900 4695 4504 4349 4222 1000 4639 4448 4293 4166 1250 4509 4319 4165 4039 1500 4381 4203 4053 3927 1750 4229 4068 3932 3816 2000 4102 3918 3795 3689 2500 3893 3711 3563 3441 3000 3711 3531 3385 3265 3500 3551 3374 3229 3110 4000 3411 3235 3092 2973 4500 3287 3113 2971 2854 5000 3177 3004 2863 2747 6000 2992 2822 2683 2568 7000 2843 2675 2538 2424 8000 2722 2556 2421 2289 9000 2623 2459 2325 2213 10000 2540 2379 2247 2136 *Two Block DM-SL burns, inclination 63.4º, argument of perigee 270º
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High-Energy Figure 3-10 and Table 3-8 show the Zenit-3SL payload capability to high Orbits and Earth energy orbits and Earth escape. These are presented as a function of C3 Escape (velocity squared at infinity) and declination of the outgoing asymptote.
6000 Perigee height = 200 km One Block DM -SL Burn
5000 Declination, deg
0
4000 30
3000 Payload capability,kg
2000
-15 -10 -5 0 5 10 15 20 25
2 2 C3, km /sec SHG0681874-030 Figure 3-10. High-Energy and Earth Escape Payload Capability
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Table 3-8. High-Energy and Earth Escape Payload Capability* Payload capability, kg 2 2 C3, km /s Declination = 0 deg Declination = 30 deg –18 6403 6132 –16 6114 5855 –14 5832 5585 –12 5557 5322 –10 5292 5066 –8 5035 4818 –6 4788 4579 −4 4550 4348 –2 4320 4125 0 4098 3911 2 3885 3704 4 3679 3504 6 3480 3312 8 3289 3126 10 3104 2947 12 2926 2774 14 2753 2607 16 2588 2446 18 2427 2290 20 2273 2140 22 2123 1995 24 1979 1854 26 1840 1719 28 1705 1587 30 1574 1461 *One Block DM-SL burn, perigee height 200 km
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Low Earth Orbit In order to maximize the payload capability to low Earth orbit (LEO), a two-stage version of the Sea Launch System could be developed. This could deliver as much as 15,000 kg to LEO, or 12,000 kg to the Interna- tional Space Station orbit. Higher LEO altitudes can be achieved by ex- tending the duration of the Stage 2 vernier engine burn. Since develop- ment work would be required for a two-stage vehicle, contract lead times would be much longer than those set forth in this document. Please con- tact Sea Launch to discuss this option. The maximum LEO capability of the three-stage Zenit-3SL is 7300 kg, due to structural limitations.
3.4 Coast Phase Attitude
Three-Axis The Block DM-SL can accommodate three-axis attitude pointing (includ- Pointing ing sun pointing) while coasting between Block DM-SL main engine burns. Attitude accuracy depends heavily on spacecraft mass properties but, for preliminary purposes, may be assumed to be ±3 deg (2.3σ) in all three axes.
Roll Maneuvers For missions with lengthy coasts between Block DM-SL main engine burns, continuous rolls for spacecraft thermal control and battery charg- ing are supported. For a roll rate of 5 deg/s, the roll-axis pointing accu- racy is ±4 deg, and the attitude rate accuracy is ±0.2 deg/s in all three axes (2.3σ). For preliminary planning purposes, it may be assumed that 40 min of each coast phase is reserved for other Block DM-SL attitude maneuvers (e.g., reorientation to/from the roll attitude).
3.5 Spacecraft Separation
Separation Spacecraft separation typically occurs 10 to 11 min after the final Block Timeline DM-SL main engine shutdown. This allows for engine cooling and reori- entation to the required spacecraft separation attitude. Several hours after separation, the Block DM-SL performs a contamina- tion and collision avoidance maneuver (CCAM), which prevents future contact with the spacecraft. The Block DM-SL then vents all residual propellant and gases and depletes any remaining charge in its batteries.
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Injection The injection accuracy at spacecraft separation is a function of the injec- Accuracy tion orbit and the mission flight duration. Table 3-9 shows the injection accuracy (2.3σ) for a typical geosynchronous transfer orbit mission using both standard and reduced Block DM-SL flight performance reserves. Table 3-9. Typical Geosynchronous Transfer Orbit Injection Accuracy (2.3σ) Injection accuracy Orbital parameter Standard FPR Reduced FPR Perigee height ± 10 km ± 10 km Apogee height ± 80 km +80/-1000 km Inclination ± 0.25 deg ± 0.25 deg
Separation The Zenit-3SL can provide the following spacecraft separation modes: Capabilities • Transverse spin—Two sets of springs with differing energies are used to impart a relative separation velocity as well as the desired angular velocity about a specified spacecraft axis. • Longitudinal spin—Before the separation event, the Block DM-SL rotates to achieve the desired angular velocity about the lon- gitudinal axis. Uniform separation springs impart relative separation velocity to the spacecraft. • Three-axis stabilized—Uniform separation springs impart relative separation velocity while the Block DM-SL minimizes pre-separation rotational rates about each axis. The tolerances of the separation parameters associated with each of the different spacecraft separation events heavily depend on the spacecraft mass properties and their associated uncertainties. Sea Launch works closely with the spacecraft customer to understand and define separation requirements that ensure mission success.
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4.0 MISSION INTEGRATION
Overview This section details the mission integration process, the mission management structure and the roles and responsibilities of mission personnel. It also defines baseline and mission-specific documentation. The following characteristics of mission integration are covered in Sections 4.1 through 4.3: • Management structure • Mission analysis and operations planning • Mission documentation
4.1 Management Structure
To provide the customer with an efficient mission integration and launch operations process, Sea Launch provides direct lines of contact between the customer and the Sea Launch organization. Figure 4-1 illustrates this management structure for both the planning phase and the launch campaign.
Sea Launch/Boeing Sea Launch partner Spacecraft Mission Integration Team mission integration customer team
Spacecraft PLU Mission manager interface control document Mission integrator interface control document • Spacecraft • LV to PLU interface information • Mission design • Mission details • Flight software • Operational development requirements • Requirements • Requirement verification verification BCSC technical specialists Sea Launch/Boeing Flight design, operations team coupled loads, shock, thermal, Spacecraft campaign document vibration, acoustics, contamination, • SC to PLU interface electrical interfaces • Mission-specific operational requirements • Operational constraints Mission Integration SHG0681874-011.1 Figure 4-1. Management Structure
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Mission Manager For each mission, Sea Launch assigns a mission manager to be and Team responsible for all mission-related customer support through launch and post-launch reporting. The mission manager leads a mission team comprised of analysts, operations and contracts personnel. The mission team is responsible for all mission analyses, operations planning and integration activities.
Mission Team The mission manager acts as the primary point of contact with the Roles and customer throughout the integration and launch process. When the Responsibilities spacecraft arrives at Home Port, the focus shifts from analytical integration to the physical integration of the spacecraft to the launch vehicle and the launch campaign, with other mission team members, including the mission director, taking a more active role. Table 4-1 shows mission team roles and responsibilities. Table 4-1. Mission Team Roles and Responsibilities Title Roles and Responsibilities Customer focal point • Works closely with the mission manager and the payload integration manager • Has a direct line of contact with the mission director during operations • Is empowered to make relevant commitments on behalf of the customer Sea Launch Mission • Primary customer interface during payload Manager integration and mission analysis • Leads the mission team Sea Launch Mission • Coordinates BCSC analytical and integration Integrator activities • Creates mission specific documentation • Coordinates Yuzhnoye and Energia analytical and integration activities Sea Launch Mission • Directs the operations phase and the launch Director campaign Sea Launch Payload • Primary operations customer interface Integration Manager • Responsible for planning and integrating all spacecraft operations at Home Port, and on board the Assembly and Command Ship and Launch Platform
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4.2 Mission Analysis and Operations Planning
Mission Analysis Mission analysis begins following a contractual commitment between Sea Launch and the customer and continues until submittal of the final post-launch report to the customer. Mission analysis is conducted by Boeing, Energia and Yuzhnoye specialists. The mission analysis process assures compatibility of the spacecraft with the launch vehicle and with the overall Sea Launch system. These tasks encompass: • Interface requirements development and documentation • Requirements verification planning • Requirements verification • Performance, interface, and environmental analyses • Flight trajectory design and flight software development
Mission Specialists Boeing, Energia, and Yuzhnoye mission analysis personnel include specialists in: • Systems engineering • Electrical and mechanical interface design • Structural loads and dynamics analysis • Mass properties • Thermal analysis • Electromagnetic interference and compatibility • Communications • Venting analysis • Trajectory and orbit design • Spacecraft separation systems • Contamination control Boeing mission analysis personnel conduct the analyses required to verify compatibility of the spacecraft with the Zenit-3SL launch vehicle and the Sea Launch system. Energia mission analysis personnel conduct the analyses required to develop and implement the Block-DM flight data and integrated vehicle flight software. Yuzhnoye mission analysis personnel conduct the analyses required to develop and implement the Zenit booster flight data and flight software.
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Mission Design The Preliminary Design Review (PDR) is an Interface Control Document Reviews (ICD) review held early in the integration cycle (for first-of-type SC) to provide a detailed evaluation of the mission requirements documented in the preliminary spacecraft to Sea Launch ICD. The purpose of the meeting is to assure that the ICD requirements are understood and that there is a plan established for meeting those requirements. The Mission Manager conducts this review. All requirements from the preliminary ICD will be addressed along with the preliminary design, analyses and planning that ensures requirements will be met. The Critical Design Review (CDR) is an ICD review to provide a detailed review of the mission requirements documented in the final spacecraft to Sea Launch ICD. The purpose of the meeting is show that the ICD requirements have either been satisfied or that that there is a plan in place for satisfying the requirements. The Mission Manager conducts this review.
Operations Operations planning focuses on tailoring the baseline operations process Planning and products to accommodate each specific mission. The mission- specific spacecraft campaign documentation is used to document all of the planning and products. These operations products include: • Operations flows • Timelines and checklists • Operations policies • Detailed procedures Operations planning also addresses mission-specific requirements for personnel training and certification, readiness reviews and rehearsals.
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4.3 Mission Documentation
Overview Sea Launch has developed baseline documentation for a generic mission, which is used as a template for each mission. This proven process assures consistency and repeatability across missions. This generic documentation includes: • Mission analysis documentation that defines the spacecraft interfaces and the initial mission integration process • Operations plans and products that define and support the actual mis- sion conduct Mission-specific documentation is also defined in this section. Figure 4-2 illustrates the documentation set that supports each mission.
Generic Integration Analysis and Operations
Sea Launch SC/SLS ICD Analysis plan Safety Plan Facilities Handbook Safety SC ICD Contamination Regulations verification control plan Manual plan EMI/EMC SL facilities control plan handbook for customers
SL facilities spacecraft cable diagrams Customer input Generic • IRD/draft ICD • Spacecraft schedule Mission • Spacecraft operations specific plan
Mission integration Launch operations SC/SLS ICD schedule schedule
Mission specific SC ICD Readiness analysis reports verification Review packages • Coupled loads matrix • Venting Operational • Clearance Verification documentation task sheets • SC separation • Campaign schedules • PLF separation • Launch commit criteria • Thermal • Launch definition document • EMI/EMC • SC test requirements • Contamination • SC operations • SC campaign requirements document Mission design • Master countdown procedure reviews packages • Training material Training • Certifications Postflight report • Rehearsals SHG0681874-018.2 Figure 4-2. Mission Documentation and Related Products
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Integration The Statement of Work, which is part of the Launch Service Agreement, Documentation defines integration inputs, analyses and deliverables to be provided by Sea Launch and the customer. At the beginning of the integration process, the customer will receive the generic documentation used for spacecraft integration on all Sea Launch missions. These documents allow the customer to understand the generic integration process. This documentation includes: • Mission integration schedule • Contamination control plan • Electromagnetic interference and electromagnetic control plan • Sea Launch System Safety Plan and the Sea Launch Safety Regula- tions Manual If the customer provides an interface requirements document, it will be the basis for generating the mission-specific spacecraft interface control document. The interface control document allows the customer to define unique interfaces between the spacecraft and the Sea Launch system.
Spacecraft/Sea The spacecraft/Sea Launch system interface control document defines all Launch System spacecraft and spacecraft ground support equipment to Sea Launch Interface Control environmental, orbital and operational interfaces. It also contains the Document mission-unique facility interfaces. (Standard facility interfaces are described in the Sea Launch Facilities Handbook)
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Mission-Specific The customer-provided deliverables (Appendix A) and/or spacecraft Documentation IRD, as well as the generic integration schedule template are used to define the specific documentation for each mission. A generic mission integration schedule is shown in Figure 1-1. The mission-specific integration documentation consists of: • Unique mission integration schedule • Mission-specific spacecraft/Sea Launch system interface control doc- ument • Verification matrix • Mission design and analysis reports
The mission-specific operations documentation consists of: • Spacecraft Test Requirements Document • Spacecraft Operations Requirements Document • Spacecraft Campaign Document • Tailored operations procedures • Launch Commit Criteria • Master Countdown Procedure • Training materials Following the mission, a post-launch report is generated for the cus- tomer.
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Mission The Tier I mission integration schedule will be developed after a contract Integration is signed resulting in a customer spacecraft launch commitment on the Schedule Sea Launch manifest. The Tier I integration schedule is produced by the mission integrator and coordinated between the mission manager and the customer. This schedule includes: • Spacecraft integration milestones • Data deliverables (spacecraft models and analysis reports) • Reviews • Integration activities • Schedule of customer reviews • Hardware need dates • Spacecraft integration milestones A separate schedule for the launch campaign will be provided by the payload integration manager. This schedule delineates: • General operations flow • Overall training and rehearsal requirements • Integrated schedules for processing and launch operations • Key Home Port, Assembly and Command Ship, and Launch Platform events
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Analysis The Statement of Work lists the analysis plan for the specific mission. This includes: • Each of the specialty areas where analyses are performed • Inputs required from the customer • Tasks to be accomplished • Specific deliverables to be provided to the customer The plan covers analyses in the areas of: • Thermal • Shock and vibration • Acoustic • Flight design • Radio frequency links • Separation clearances • Contamination • Electromagnetic interference and • Loads electromagnetic compatibility • Venting
Spacecraft A spacecraft verification matrix defines the process by which the inter- Verification face control document requirements and functions are verified before and Matrix during launch processing. The matrix is included in the interface control document.
Operations The generic operations documentation available to the customer for Documentation understanding the operations details consists of: • Sea Launch Facilities Handbook • Sea Launch safety documentation • Training, certification and rehearsal documentation • Sea Launch security documentation
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Operations The operations planning process defines the detailed plans and processes Planning for all Sea Launch operations activities, which include: • Operations phases • Rehearsals • Concepts and activities for • Readiness reviews each phase • Management team • Operations products • Support facilities and • Overall control processes capabilities • Personnel training and certifi- cation processes
Sea Launch The Sea Launch Facilities Handbook describes those portions of the Sea Facilities Launch Home Port, Payload Processing Facility, Assembly and Handbook Command Ship and Launch Platform required by the customer during spacecraft processing and launch operations. Any modifications made to standard facility interfaces will be included in the mission-specific spacecraft/Sea Launch system interface control document.
4-10 HPD-10041 REV D 5.0 SPACECRAFT ENVIRONMENTS
Overview This section describes the major environments the spacecraft is exposed to from the time of its arrival at Home Port until its separation from the launch vehicle, including: • Transportation and handling at sea • Thermal • Structural loads • Humidity • Random vibration • Pressure • Acoustics • Venting • Shock • Contamination • Electromagnetic radiation
Ground and Flight For each of the above environments, the levels are generally categorized Environments by time of occurrence. The levels designated as “ground transportation and handling” cover the period of time from the arrival of the spacecraft at Home Port until Stage 1 ignition. The levels designated as “flight” cover the period of time from Stage 1 ignition (liftoff) through spacecraft separation from the launch vehicle, and end with the completion of the Block DM-SL contamination and collision avoidance maneuver (CCAM). Although the Sea Launch system has a unique processing flow, every effort has been made to ensure that the environments experienced by the spacecraft will be less than or comparable to those provided by other launch systems. More information on environments that apply to ground support equipment is located in the Sea Launch Facilities Handbook. Please contact your Sea Launch representative for more detailed information.
REV D HPD-10041 5-1 5.1 Structural Loads
Overview The structural loading environments on all spacecraft primary and secondary structure are examined for: • Ground transportation and handling • Flight • Spacecraft sinusoidal vibration testing Compliance requirements of the spacecraft to these environments are discussed in Section 6.
Quasi-Static Load The Sea Launch system will not induce loads on the spacecraft that Factors for exceed the load factor envelope shown in Figure 5-1 during spacecraft Ground Handling ground and sea transportation, handling or processing. and The load factor levels are applied quasi-statically at the spacecraft center Transportation of gravity. The (x,y,z) load factor components are contained within a sphere with radius 0.5 g centered about the 1.0 g contribution due to gravity. The load factor envelope in Figure 5-1 includes all contributions generated from the Launch Platform motion (angles, rates, and accelerations) during ocean transport. In particular, angle changes of the spacecraft as a result of the pitch and roll of the Launch Platform are included. The quasi-static load factors in Figure 5-1 also determine the maximum change in angle of a fluid’s free surface in a tank to be 30 deg for spacecraft slosh frequencies ≥0.25 Hz. The load factors in Figure 5-1 envelop all ground transportation and handling events from arrival at Home Port to Stage 1 ignition and apply to all orientations of the spacecraft. The orientation of the spacecraft when the Zenit-3SL is horizontal or vertical is shown in Figures 5-2 and 5-3, respectively.
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Vertical g
1.5 R 0.5 g
1.0
0.5
-0.5 Horizontal g 0.5 265807J6-249 Figure 5-1. Transportation and Handling Limit Load Factors
Vertical
Horizontal
Horizontal 265807J6-265 Figure 5-2. Integrated Launch Vehicle Horizontal
Vertical
Horizontal Horizontal
265807J6-266 Figure 5-3. Integrated Launch Vehicle Vertical
REV D HPD-10041 5-3 Fatigue The fatigue loads during ocean transit are usually less than those experienced during rail transport at other launch sites. Table 5-1 depicts fatigue loads on the spacecraft for Sea Launch, covering ocean transit and all ground handling operations at Home Port.
Table 5-1. Spacecraft Fatigue Environment for Transportation and Handling Load, g 0.39 0.37 0.33 0.29 0.25 0.22 0.18 0.14 0.10 Number of 1 5 21 92 387 1,569 6,086 21,842 1.0E+08 cycles Notes: • Vertical loads are in addition to gravity: g (vertical) = 1.0 ± load (g). • Environment covers three trips (1.5 round trips) between Home Port and the equatorial launch site. • Loads act uniformly along the length of the spacecraft. • The dynamic loads act in all directions simultaneously.
Pitch and Roll Table 5-2 depicts the maximum pitch and roll motions during transport between Home Port and the launch site. Launch Platform motions apply to spacecraft and ground support equipment while Assembly and Command Ship motions apply only to ground support equipment. The quasi-static load factors for ocean transport in Figure 5-1 account for Launch Platform motions. In particular, the effect of the Launch Platform pitch and roll motions on spacecraft fluid slosh is taken into account in Figure 5-1.
Table 5-2. Motions During Transit to Launch Site Maximum Assembly and Motions Command Ship Launch Platform Pitch 8 deg 8 deg Pitch rate 5.5 deg/s 5.5 deg/s Roll 25 deg 10 deg Roll rate 21 deg/s 7.5 deg/s
5-4 HPD-10041 REV D Coupled Loads Determining the ability of the spacecraft primary and secondary structure Analysis Required to withstand the dynamic loading events from flight requires a coupled to Determine Qua- loads analysis. Sea Launch performs a coupled loads analysis for all si-Static Loads for critical flight events to determine the accelerations, loads and deflections Flight of all primary and secondary structure of the spacecraft caused by the coupled elastic behavior of the Zenit-3SL launch vehicle and the spacecraft. The results determine the interface loads of the spacecraft with the spacecraft adapter. The resulting equivalent spacecraft quasi- static load factors at the spacecraft center of gravity are then determined for a specific spacecraft with given mass and stiffness properties.
Flight Quasi-Static The quasi-static limit load factors for flight that envelop the loads over a Load Factors for wide range of spacecraft are set forth in Figure 5-4. These factors are to Primary Structure be used as a guide only. Limit load factors for spacecraft with specific mass and stiffness properties will be defined in the spacecraft interface control document after a coupled loads analysis has been performed. Lateral load factors ranging from 1.5 to 2.5 g have been calculated for spacecraft with unique mass and/or stiffness properties.
Axial load factor, g 5 (–0.7, +4.5) (+0.7, +4.5) 4 3 (–2, 2.5) (+2, +2.5) 2 Lateral load 1 factor, g (–2, 0) (+2, 0) 0 –4 –2 2 4 –1
(–1, –2) –2 (+1, –2) –3 Notes: • Positive axial acceleration results in compression at the SC-LV separation plane. • Guide only. Specific values determined after spacecraft CLA. 265807J6-058R4 Figure 5-4. Typical Quasi-Static Load Factors for Flight
REV D HPD-10041 5-5 Sinusoidal The low-frequency sinusoidal vibration environment generated at the Vibration During spacecraft separation plane during launch and flight will not exceed that Flight defined in Figure 5-5. A coupled loads analysis performed by Sea Launch determines the minimum sinusoidal vibration environment for all flight events. The results of the coupled loads analysis determine the maximum notching that can be used during the spacecraft sinusoidal vibration testing.
1.2 1.0 0.8
0.6 Axial and lateral 0.6 (5-100 Hz) 0.4
0.2 Axial and lateral g acceleration, base 0 0 20 40 60 80 100 Frequency, Hz 265807J6-059R2 Figure 5-5. Sinusoidal Vibration at the Spacecraft Interface
5.2 Random Vibration
Random Vibration The random vibration environment for flight measured at the spacecraft for Components interface is depicted in Figure 5-6. Near Spacecraft Maximum values occur during liftoff and are closely correlated with the Interface acoustic environment. The environment in Figure 5-6 applies to components within 0.5 m from separation plane along any structural path. This environment is not to be applied to the complete spacecraft as a rigid base excitation.
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0.1
0.01 6.8 grms Acceleration power spectral density spectral power Acceleration 0.001 10 100 1000 10000 Frequency, Hz 265807J6-269 Figure 5-6. Random Vibration Environment During Flight 5.3 Acoustics
Space Average The space average sound pressure levels internal to the payload unit are Sound Pressure depicted in Table 5-3 (Figure 5-7) for 1/3-octave band center frequencies Levels and Table 5-4 (Figure 5-8) for the equivalent 1-octave center band frequencies. The values for the “Reference SC” apply to a spacecraft with an equivalent radius that result in a spacecraft-to-payload fairing gap of 43 cm (1.5 ft). The values for “Typical SC” apply to a larger, rectangular spacecraft that fits within the static envelope. Higher acoustic levels are generally expected for spacecraft that exceed the static envelope.
Fill Effects to be The equivalent radius of the spacecraft is defined as the radius of a circle Considered with a cross-sectional area equal to the spacecraft cross-sectional area. The envelope of the internal maximum expected sound pressure levels also can be provided after clocking and fill effects for a specific spacecraft are determined. Section 6 defines the compliance requirements of the spacecraft to the internal acoustic environment.
REV D HPD-10041 5-7 Table 5-3. Acoustics During Flight⎯1/3-Octave Band Sound Pressure Levels Sound pressure levels, dB Typical SC con- 1/3-octave band center Reference SC tained within static frequency, Hz with 43 cm gap envelope 31.5 119 123 40 121 123 50 123 124 63 125 125 80 127 127 100 129 129 125 130 130 160 131 131 200 133 133 250 134 134 315 133 133 400 131 131 500 129 129 630 127 127 800 125 125 1,000 122 122 1,250 121 121 1,600 120 120 2,000 119 119 2,500 118 118 3,150 117 117 4,000 115 115 5,000 114 114 6,300 113 113 8,000 111 111 10,000 110 110 OASPL 141.6 141.7 • Reference SC has 43 cm clearance between SC and the payload fairing • Reference: 2 x 10-5 Pa (2.9 x 10-9 psi)
5-8 HPD-10041 REV D Table 5-4. Acoustics During Flight⎯1-Octave Band Sound Pressure Levels Sound pressure levels, dB Typical SC con- 1-octave band center Reference SC tained within static frequency, Hz with 43 cm gap envelope 31.5 125.2 127.6 63 130.1 130.3 125 134.8 134.8 250 138.1 138.1 500 134.1 134.1 1,000 127.8 127.8 2,000 123.8 123.8 4,000 120.3 120.3 8,000 116.3 116.3 OASPL 141.6 141.7 • Reference SC has 43 cm clearance between SC and the payload fairing • Reference: 2 x 10-5 Pa (2.9 x 10-9 psi)
140
Typical SC 135 Reference SC
130
125
120
115 (Ref:micropascal) 20
110 1/3 Octave Band Sound Pressure Level - dB 105
100 10 100 1000 10000 1/3 Octave Band Center Frequency – Hz SHG0681874-031 Figure 5-7. Acoustics During Flight⎯1/3-Octave Band Sound Pressure Levels
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135 Typical SC Reference SC
130
125
120
115
110 1 Octave Band SPL (dB) Ref. 20 micropascal
105
100 10 100 1000 10000 1 Octave Band Center Frequency (Hz) SHG0681874-032 Figure 5-8. Acoustics During Flight⎯1-Octave Band Sound Pressure Levels
5.4 Shock
Overview Shock environments are provided for the three spacecraft adapter types: • The SCA1194VS: has a diameter of 1194 mm and a Marmon clamp separation system. • The SCA1666: has a diameter of 1666 mm and a Marmon clamp separation system. • The SCA702+ and SCA702+GEM: have diameters of 1663 mm and four-bolt separation systems. The maximum shock environments for these systems are consistent with industry standards. Check with Sea Launch for environments associated with other spacecraft adapters.
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SCA1194VS Sea Launch, together with Saab Space, has developed a low-shock, 1194-mm spacecraft separation system (designated the SCA1194VS) for the Zenit-3SL. In addition to reduced shock environment, the SCA1194VS provides the ability to accommodate spacecraft weighing up to 8 metric tons by using a 60-kN clampband tension (for structural capability of the Zenit-3SL launch vehicle, see Section 6.1). The only significant difference for the SCA1194VS from the previous Saab 1194- mm adapter is the use of a fast-acting clampband opening device in place of bolt cutters to release the clampband tension, which results in a lower spacecraft shock environment. The shock spectra shown in Figures 5-9 and 5-10 and Table 5-5 describe the maximum expected shock environment in the radial and axial directions for a spacecraft adapter having a diameter of 1194 mm and a low-shock Marmon clamp separation system. The shock spectrum is defined for a location 4.0 cm from the separation plane on the spacecraft side of the interface. The shock environment is the maximum expected levels for all shock generated by the integrated launch vehicle, which includes payload fairing separation, second stage staging from Block DM-SL, and spacecraft separation events. For the SCA1194VS, the shock levels from spacecraft separation are lower than those from payload fairing separation and second stage staging from Block DM-SL shock events. Spacecraft compatibility to the radial and axial shock levels from all shock events must be considered and evaluated by Sea Launch. The shock spectrum corresponds to a tension of 40 kN to 60 kN in the spacecraft adapter clampband. The shock levels are generated using conservative shock transmission losses at the spacecraft interface. Actual shock levels on the spacecraft side of the interface may be lower using actual transmissions losses unique to the spacecraft bus. The actual losses for a spacecraft bus are determined during matchmate testing (see Section 6.4).
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SCA1194VS spacecraft separation shock, radial axis
All other ILV shock events - radial axis Acceleration Shock Response Spectra, gpeak (Q=10) (Q=10) gpeak Spectra, Response Shock Acceleration 1 100 1000 10000 Frequency - Hz
Figure 5-9. Payload Shock Environment 4.0 cm Above the Separation Plane using 1194VS Spacecraft Adapter, Radial Axis
10000
1000
100
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SCA1194VS spacecraft separation shock, axial axis
All other ILV shock events - axial axis Acceleration Shock Response Spectra, gpeak (Q=10) (Q=10) gpeak Spectra, Response Shock Acceleration 1 100 1000 10000 Frequency - Hz
Figure 5-10. Payload Shock Environment 4.0 cm Above the Separation Plane using 1194VS Spacecraft Adapter, Axial Axis
5-12 HPD-10041 REV D Table 5-5. Payload Shock Environment 4.0 cm Above the Separation Plane using 1194VS Spacecraft Adapter Frequency, Hz Spacecraft Separation Shock - Radial Axis, g 200 60 1200 1400 5000 1400 Frequency, Hz All Other ILV Shock Events - Radial Axis, g 100 20 1000 1400 5000 2000 Frequency, Hz Spacecraft Separation Shock - Axial Axis, g 200 20 2000 700 5000 700 Frequency, Hz All Other ILV Shock Events – Axial Axis, g 100 20 1000 1000 5000 2000
SCA1666 The shock spectrum shown in Figure 5-11 and Table 5-6 represents the maximum expected shock environment for the spacecraft adapter with a diameter of 1666 mm and a Marmon clamp separation system. The shock spectrum corresponds to a tension of 30 kN in the spacecraft adapter clampband and is defined for a location 12.7 cm (5 in) from the separation plane on the spacecraft side of the interface.
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100
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All SCA1666 shock events - radial and axial axes Acceleration ShockResponse Spectra, gpeak(Q=10) 1 100 1000 10000 Frequency - Hz
Figure 5-11. Payload Shock Environment 12.7 cm Above the Separation Plane using 1666 Spacecraft Adapter
Table 5-6. Payload Shock Environment 12.7 cm Above the Separation Plane using 1666 Spacecraft Adapter Frequency, Hz All Shock Events - Radial and Axial Axes, g 100 50 800 3000 3000 3000 5000 3500
SCA702+ and The Sea Launch system generated separation shock spectrum using the SCA702+GEM SCA702+ and SCA702+GEM spacecraft adapters will not exceed that defined in Figure 5-12 and Table 5-7, as measured 12.7 cm (5 in) from the separation plane on the spacecraft side of the interface on the spacecraft attachment fittings. Section 6.1 sets forth the strength capabilities of the SCA702+ and SCA702+GEM spacecraft adapters.
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100
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SCA702+ spacecraft separation shock - radial and axial axes
All other ILV shock events - radial and axial axes Acceleration ShockResponse Spectra, gpeak(Q=10) 1 100 1000 10000 Frequency, Hz
Figure 5-12. Payload Shock Environment 12.7 cm Above the Separation Plane using 702+ and 702+GEM Spacecraft Adapters
REV D HPD-10041 5-15 Table 5-7. Payload Shock Environment 12.7 cm Above the Separation Plane using 702+ and 702+GEM Spacecraft Adapters Spacecraft Separation Shock - Radial and Axial Freq axes, g 100 30 2000 1000 5000 3000
Freq All Other ILV Shock Events - Radial and Axial axes, g 100 30 500 424 1000 650 1500 650 3000 1400 5000 1400
5.5 Electromagnetic Environment
Overview The electromagnetic radiation environment experienced by the spacecraft will be less than 10 V/m at frequencies between 15 kHz and 40 GHz. This applies to: • Processing cells in the Payload Processing Facility (Figure 5-13) • The Home Port pier (Figures 5-14 and 5-15) • The launch vehicle assembly compartment on board the Assem- bly and Command Ship • The Launch Platform hangar • The spacecraft while erected on the Launch Platform Note: Some levels shown exceed 10 V/m, but suppression and operational constraints are used to keep all levels impinging on the spacecraft below 10 V/m.
5-16 HPD-10041 REV D 120
Legend: Inside encapsulation cell
100
80 V/m μ
60 Electric field, dB field, Electric 40
20
0 1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+08 1.00E+09 1.00E+10 1.00E+11 Frequency, Hz 265807J6-273 Figure 5-13. Typical Electromagnetic Environment in Payload Processing Facility⎯Encapsulation Cell, 15 kHz to 18 GHz
REV D HPD-10041 5-17 140
120
100 V/m
μ 80
60 Electric field, dB field, Electric
40
20
0 1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+08 1.00E+09 1.00E+10 1.00E+11 Frequency, Hz 265807J6-274 Figure 5-14. Typical Measured Electromagnetic Environment at the Sea Launch Home Port Pier, 15 kHz to 18 GHz
5-18 HPD-10041 REV D 140
120
100 V/m μ 80
60 Electric field, dB field, Electric 40
20
0 10 20 30 40 50 60 70 80 90 100 Frequency, GHz SHG0681874-035 Figure 5-15. Typical Measured Electromagnetic Environment at the Sea Launch Home Port Pier, 18 GHz to 40 GHz
Radiated There are six telemetry transmission systems and two command-receive Environment systems mounted on the launch vehicle, as follows: • The Sirius and Zenit telemetry systems are located on the Zenit second stage, and transmit at 211.3, 219.3, 1010.5, 1018, and 2211.0 MHz. • The Kvant and BR9DM systems are located on the interface skirt. The Kvant system transmits at 922.76 MHz, and the BR9DM transmits at 166 MHz. • The payload unit telemetry system transmits at 2272.5 MHz using several antennas that are mounted on the interface skirt. Figure 5-16 shows the in-flight electromagnetic environment, produced by the launch vehicle transmitters, predicted by analysis to be present at the LV/SC interface.
REV D HPD-10041 5-19 180 Kvant 160 922.76 MHz, Sirius TM System Sirius TM System 140 dBuV/m 1010.5 MHz, 1018 MHz 140 211.3 MHz, 127 dBuV/m 219.3 MHz, 120 126 dBuV/m PLU TM System BR9DM 100 2,272.5 MHz 166.0 MHz, 140 dBuV/m 129 dBuV/m 80
60 Unintentional Radiated Emissions < 70 dBuV/m Electric Field (dBuV/m)
40
20
0 10 100 1000 10000 100000 Frequency (MHz) SHG0681874-036 Figure 5-16. Predicted In-Flight Electromagnetic Environment at the LV/SC Interface
Electromagnetic Table 5-8 shows the electromagnetic environment generated by the Sea Environment Launch transmitters at Home Port and on the Assembly and Command Ship, Launch Platform, and launch vehicle.
Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies Predicted E-Field (dBµV/m) at Frequency PLF Surface Comments on Transmitter (MHz) Location Vessels at HP Vessels at LS Predicted Levels Non- 0.19-0.535 LP 116 116 Worst case levels Directional Beacon MF-HF 1.6-30 ACS 123 89 Worst case levels MF-HF 1.6-30 LP 127 127 Worst case levels VHF/AM 118-137 ACS 103 68 Worst case levels Fixed 1 (heli- copter) VHF/AM 118-137 ACS 104 68 Worst case levels Fixed 2 (heli- copter)
5-20 HPD-10041 REV D Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies (continued) Predicted E-Field (dBµV/m) at Frequency Comments on Transmitter Location PLF Surface (MHz) Predicted Levels Vessels at HP Vessels at LS VHF/AM 118-137 LP 103 103 Worst case levels Fixed (heli- copter) VHF/AM 118-137 LP 101 103 Worst case levels Handheld (helicopter) Homing 121.5 ACS 88 53 Worst case levels transmitter Homing 121.5 LP 86 86 Worst case levels transmitter VHF Lifeboat 150.8- ACS 99 64 Worst case levels 163.6 VHF Lifeboat 150.8- LP 97 97 Worst case levels 163.6 VHF Duplex 155-158 ACS 112 78 Worst case levels (ship-to-ship) VHF 1 (ship- 155-163.6 ACS 113 78 Worst case levels to-ship) VHF 2 (ship- 155-163.6 ACS 113 78 Worst case levels to-ship) VHF 1 (ship- 155-163.6 LP 114 114 Worst case levels to-ship) VHF 2 (ship- 155-163.6 LP 114 114 Worst case levels to-ship) VHF DSC 156.525 ACS 113 78 Worst case levels (continuous) VHF DSC 156.525 LP 114 114 Worst case levels (continuous) VHF AIS 156-162 ACS 82.2 82.2 Worst case levels VHF AIS 156-162 LP 116 116 Worst case levels BR9DM 116 ILV 116 129 Transmit power limited at HP to comply with FCC license Sirius-LV 211.3 ILV 107 126 Transmit power limited at HP to comply with FCC license
REV D HPD-10041 5-21 Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies (continued) Predicted E-Field (dBµV/m) at Frequency Comments on Transmitter Location PLF Surface (MHz) Predicted Levels Vessels at HP Vessels at LS Sirus-LV 219.3 ILV 106 125 Transmit power limited at HP to comply with FCC license EPIRB 406.025 ACS 105 70 Worst case levels EPIRB 406.025 LP 103 103 Worst case levels RM/C/C- 420.8 LP 78 96 Transmit power SVET limited at HP to comply with FCC license RM/C/C- 421 LP 78 96 Transmit power SVET limited at HP to comply with FCC license R/M/C/C- 425.8 ACS 80 68 Transmit power LUCH limited at HP to comply with FCC license R/M/C/C- 426 ACS 80 68 Transmit power LUCH limited at HP to comply with FCC license UHF Port- 430-470 ACS 102 70 Worst case levels ables UHF Remote 433.2- ACS 80 44 Worst case levels Crane OP 433.7 UHF Remote 441.725 LP 95 95 Worst case levels Crane OP UHF Port- 457.525- LP 106 107 Worst case levels ables 470 Boeing Two- 476.85 ACS 98.2 63.8 Worst case levels Way Radio SCRFS over 480-712 LP and 127 93 Worst case levels the water and 826- ACS 890 SCABS over 480-712 ACS and 124 85 Worst case levels the water and 826- LP 890
5-22 HPD-10041 REV D Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies (continued) Predicted E-Field (dBµV/m) at Frequency Comments on Transmitter Location PLF Surface (MHz) Predicted Levels Vessels at HP Vessels at LS Kvant-ACS 768.97 ACS 85 117 Transmit power limited at HP to comply with FCC license Kvant-DM 922.76 ILV 118 127 Transmit power limited at HP to comply with FCC license Sirius-LV 1010.5 ILV 118 127 Worst case levels Sirus-LV 1018 ILV 127 112 Worst case level Inmarsat B1 1626.5- ACS 129 95 Based on an- 1646.5 tenna gain in SC direction Inmarsat B2 1626.5- ACS 131 95 Based on an- 1646.5 tenna gain in SC direction Inmarsat C 1626.5- ACS 118 84 Based on an- 1646.5 tenna gain in SC direction Inmarsat B1 1626.5- LP 135 131 Based on an- 1646.5 tenna gain in SC direction Inmarsat B2 1626.5- LP 135 131 Based on an- 1646.5 tenna gain in SC direction Inmarsat C 1626.5- LP 120 120 Based on an- 1646.5 tenna gain in SC direction Zenit Teleme- 2211.0 ILV 134 134 Antenna gain in try direction of PLF surface factored in PLU Teleme- 2272.5 ILV 134 134 Antenna gain in try direction of PLF surface factored in S-band Bea- 2272.5 LP 116 115 Worst case levels con WiFi over the 2412-2808 LP and 112 110 Worst case levels water ACS
REV D HPD-10041 5-23 Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies (continued) Predicted E-Field (dBµV/m) at Frequency Comments on Transmitter Location PLF Surface (MHz) Predicted Levels Vessels at HP Vessels at LS Helicopter 2475.5 LP and 88 88 Worst case levels Data Link ACS S-band Radar 3005 ACS 133 133 ACS navigation radars turned off when ILV is erect and vessel separation <5 km S-band Radar 3050 LP 133 133 Navigation ra- dars turned off when ILV is erect Helicopter 4200-4400 LP and 81 81 Worst case levels Radar Altime- ACS ter Weather Ra- 5620 ACS 89 98 Sector blanked dar so that ILV not directly illumi- nated SATCOM 6262-6298 ACS 112 73 Antenna gain in direction of erected ILV PLF SATCOM 6262-6298 LP 110 93 Antenna gain in direction of erected ILV PLF LOS Radio 7194.5 ACS 84 86 Antenna gain in T1 direction of erected ILV PLF LOS Radio 7355.5 LP 107 127 Antenna gain in T1 direction of erected ILV PLF LOS Radio 7494.5 ACS 84 87 Antenna gain in T2 direction of erected ILV PLF LOS Radio 7655.5 LP 107 127 Antenna gain in T2 direction of erected ILV PLF Radar Trans- 9200-9500 ACS 94 59 Worst case levels ponder Life- boat
5-24 HPD-10041 REV D Table 5-8. Predicted E-Field Levels From Sea Launch Transmitters at Their Operation Frequencies (continued) Predicted E-Field (dBµV/m) at Frequency Comments on Transmitter Location PLF Surface (MHz) Predicted Levels Vessels at HP Vessels at LS Radar Trans- 9200-9500 LP 92 92 Worst case levels ponder Life- boat Helicopter 9375.0 ACS and 109 109 Worst case levels Radar LP X-band Radar 9410 ACS 136 136 Navigation ra- dars turned off when ILV is erect and separa- tion distance <5 km X-band Radar 9410 LP 136 136 Navigation ra- dars turned off when ILV is erect
REV D HPD-10041 5-25 5.6 Spacecraft Thermal and Humidity Environments
Thermal and Spacecraft thermal and humidity environments are characterized in the Humidity following sections for three primary operational phases: Environments • Spacecraft ground processing and handling before payload fairing Overview encapsulation • Operations after payload fairing encapsulation, including integrated launch vehicle integration, transportation to the launch site, and pre- launch preparations • Liftoff, ascent and in-flight operations
Ground Process- Spacecraft thermal environments are maintained and controlled during ing and Transpor- ground processing, encapsulation, launch vehicle integration, tation Before transportation, and prelaunch preparation phases. Encapsulation During operations phases in the Payload Processing Facility, before encapsulation by the payload fairing, spacecraft thermal environments will be maintained within temperature and relative humidity ranges indicated in Table 5-9. For each phase, temperature and humidity are continuously monitored and alarmed. During propellant loading periods, temperatures are selectable and can be controlled to a variance of 1.7°C (3°F).
Table 5-9. Payload Processing Facility Thermal Environments Relative Location/Phase Temperature Humidity Air lock 18–24°C 30–60% (65–75°F) Spacecraft processing 18–24°C* 30–60% and fueling cell (65–75°F) Encapsulation cell 18–24°C* 30–60% (65–75°F) Propellant storage cells 18–24°C 30–60% (65–75°F) *Selectable by customer, within range
5-26 HPD-10041 REV D Ground Process- Following encapsulation of the spacecraft in the payload fairing, the ing and Transpor- spacecraft environment is maintained with a continuous, conditioned air tation after supply. Air is ducted to the payload fairing thermal conditioning inlet Encapsulation diffuser (Figure 5-17), and exits to the Block DM-SL cavity through vent valves on the payload structure. In addition to the nominal airflow, while the ILV is in the LP hangar, Sea Launch is capable of providing directed air for cooling the spacecraft batteries with a given flow rate and temperature set-point. Modifications to the payload accommodations are required for this non-standard option, and therefore the request for this service must be provided no later than twelve months prior to the then applicable launch schedule Air-conditioning units on the Encapsulated Payload Transportation Mechanical Equipment (EPTME), Assembly and Command Ship, and Launch Platform are dedicated to maintaining spacecraft environments within the payload fairing during processing and handling at Home Port, in the Assembly and Command Ship, and while in transit to the launch site. The Launch Platform integrated launch vehicle air-conditioning system satisfies conditioned air requirements of the payload unit, Block DM-SL, and Zenit on the launch pad during prelaunch phases. The air flow from the Launch Platform integrated launch vehicle air- conditioning system is terminated at approximately L-17 minutes (on retraction of the transporter/erector and separation of the launch vehicle umbilicals). Conditioned airflow to the payload fairing is maintained through the remainder of the countdown (and through an abort recycle operation) by way of the Launch Platform high-pressure payload fairing purge system. Table 5-10 summarizes payload unit environmental control system inlet conditions and control capabilities. All air-conditioning systems maintain positive pressure inside the payload fairing to prevent air-transported contaminants from entering the payload fairing. Air quality is maintained by providing purge air, that has been high efficiency particulate air-filtered to class 5,000 maximum, in accordance with FED-STD-209E.
REV D HPD-10041 5-27 PLF PLS vents flight vents (one-way valves) Block DM flight vents Payload structure Block DM-SL Block DM-SL vents PLF volume IC vent
Spacecraft
Zenit-2SL IC Sidewall IC diffuser Block DM thermostating thermostating inlet PLF thermostating (5500 m3/hr) (3600 m3/hr) inlet (up to 4500 m3/hr)
Note: Drawing not to scale. 265807J6-010R7 Figure 5-17. Payload Unit Air-Conditioning Venting Scheme
5-28 HPD-10041 REV D Table 5-10. Payload Unit Environmental Control—Air-Conditioning Inlet Conditions Inlet temperature Operations phase range (1) RH (2) Flowrate range Home Port processing/ 10–25°C 0–60% 9–38 kg/min transportation (50–77°F) (19.8–84 lb/min) • After PLF encapsulation in PPF • Transfer from PPF to ACS • ACS processing Launch Platform hangar 10–25°C 0–60% 18.1–90 kg/min • Transport from Home Port to launch (50–77°F) (40–198 lb/min) site • At launch site • Transfer from hangar to launch pad Prelaunch operations at launch site 10–25°C 0–60% 18.1–90 kg/min • At launch pad erection to (50–77°F) (40–198 lb/min) vertical • On launch pad to L-17 min Recycle operations • After lowering of ILV, before trans- fer into hangar Prelaunch operations at launch site 10–25°C 0–60% >18 kg/min • After L-17 min to L-0 (50–77°F) (>40 lb/min) • In event of abort, through lowering of ILV until primary air reinstated Continuous air-conditioning is maintained throughout all phases by primary and/or backup air- conditioning systems. Notes: (1) Inlet temperatures are selectable by the customer within the indicated ranges and controlled to within +2°C of setpoint. Inlet temperatures are not selectable during Home Port processing operations and on launch pad after L-17 minutes. (2) Relative humidity controlled to 30–60% if access to spacecraft in payload fairing is required.
REV D HPD-10041 5-29 Flight Phases The Sea Launch payload fairing, described in Section 2, provides protection to the spacecraft from thermal environments during prelaunch, liftoff, and ascent mission phases. Effects of ascent aerodynamic heating on the spacecraft thermal environment inside the fairing are controlled by the insulating characteristics of the fairing. Ablative coatings on external payload fairing surfaces, acoustic damping material on internal surfaces, and the low conductivity of the graphite composite/aluminum honeycomb fairing structure combine to minimize heat transfer to the spacecraft from heated external surfaces during ascent. Figure 5-18 illustrates predicted internal payload fairing surface temperatures during the ascent period. Actual flight measurements are well within the predicted values. Predicted peak temperatures of acoustic blanket inner surfaces (which comprise 62% of the payload fairing area) remain below 85°C. At the time of fairing jettison, more than 90% of the fairing inner surface area is less than 95°C, 98% is under 120°C, and 2% is in the 120 to 140°C range. Materials comprising payload fairing inner surfaces generally have surface emittances ranging from ε=0.64 for surfaces covered with acoustic damping material to ε=0.83 for non- insulated inner walls. The Sea Launch system accommodates user-specified limitations on aerodynamic heat loads during periods after payload fairing jettison as constraints on the design of the ascent trajectory and at the time of fairing jettison. Maximum 3σ dispersed free molecular heat loads are calculated for mission-unique launch azimuth, day-of-launch ascent profile, and parking orbit as part of the trajectory design. For the typical spacecraft, fairing separation will occur after the maximum 3σ dispersed free molecular heat rate falls below 1135 W/m2 (360 BTU/h ft2). Payload fairing jettison will generally occur at an altitude of approximately 118,000 m (387,000 ft) at times ranging from 190 to 240 seconds following liftoff. Delaying time of fairing jettison or raising parking orbit perigee altitude to achieve a reduction in free molecular heating will have a negative effect on launch vehicle performance. Spacecraft thermal environments for the period following payload fairing jettison until spacecraft/Block DM-SL separation comprise the combined effects of free molecular heating, incident solar heating, planet-reflected solar heating (albedo), Earth-radiated heating, radiation to space, and heat transfer with the launch vehicle by radiation or conduction. The Block DM-SL flight, as indicated in Section 3.4, is designed to accommodate preferred attitude pointing, continuous rolls, maneuvers, and orientations during coast and pre-separation phases that satisfy spacecraft thermal control, battery charging, solar exposure, and other customer-specified attitude or exposure requirements.
5-30 HPD-10041 REV D 140
Legend: Tmax < 140°C for 100% of PLF area Tmax < 120°C for 98% of PLF area 120 Tmax < 101°C for 94% of PLF area Tmax < 95°C for 90% of PLF area
100
80 Temperature, °C Temperature,
60
PLF jettison range 40
20 0 60 120 180 240 Time from liftoff, sec 265807J6-276 Figure 5-18. Maximum Internal Payload Fairing Surface Temperatures
REV D HPD-10041 5-31 5.7 Payload Unit Ascent Venting
Overview During ascent, the payload unit is vented through one-way valves in the payload structure and through vents that open approximately ten seconds after liftoff. Figure 5-19 shows the venting scheme. The payload unit cavity pressure is dependent on spacecraft displaced volume and trajectory. Figure 5-20 shows a typical payload fairing cavity pressure band for a generic trajectory. The maximum ascent depressurization rate is shown in Figure 5-21. Maximum pressure differential inside the payload fairing at payload fairing jettison will not exceed 193 Pa (0.028 psi).
PLF PLS vents (one-way valves) flight vents (open at T+10 sec) Block DM flight vents Payload structure Block Block DM-SL vents PLF volume DM-SL IC vent
Spacecraft
Zenit-2SL IC Sidewall diffuser Note: Drawing not to scale. 265807J6-011R5 Figure 5-19. Payload Unit Ascent Venting Scheme
5-32 HPD-10041 REV D
120
Predicted Internal Pressure 100 Ambient Pressure
80
60 Pressure, kPa Pressure,
40
20
0 0 20 40 60 80 100 120 Time, sec SHG0681874-037 Figure 5-20. Payload Fairing Internal Pressure During Ascent
REV D HPD-10041 5-33
3.5
3.0
2.5
2.0
(kPa/Second) 1.5
1.0 Maximum Allowed Pressure Decay Rate [-dP/dt] [-dP/dt] Rate Decay Pressure Allowed Maximum
0.5
0.0 0 102030405060708090100 Absolute Pressure (kPa) SHG0681874-038 Figure 5-21. Maximum Ascent Depressurization (pressure decay rate vs. pressure)
5-34 HPD-10041 REV D 5.8 Contamination
Overview Contamination of the spacecraft by the launch vehicle is addressed in all phases of the mission. Contamination control includes the following: • Launch vehicle hardware that comes in contact with the space- craft environment has been designed, manufactured and tested to meet strict contamination control guidelines regarding materials outgassing, cleanability, particle generation, corrosion, and wear. • Integration is performed in controlled (class 100,000) cleanroom environments, including spacecraft stand-alone testing, payload accommodation preparations, and spacecraft encapsulation. • After encapsulation, the cleanliness of the spacecraft environment is ensured by the maintenance of a continuous clean-air purge to the payload fairing. • The design of the payload structure isolates the spacecraft from most on-orbit outgassing sources, such as the flight avionics hardware and the Block DM-SL. • Launch vehicle plume impingement on the spacecraft is mini- mized by physical barriers and flight design restrictions.
Contamination Processing of the payload accommodation is conducted in accordance Control with the environmental controls and operational requirements documented Documents in the Contamination Control Plan for Sea Launch Home Port Operations Through Launch Operations. The Generic Payload Contamination Analysis is a comprehensive analysis of contamination deposition on the spacecraft from all ground processing and launch vehicle sources.
Contamination Ground contamination control begins with manufacture of the payload Control During accommodation and continues through launch. The two halves of the Ground Process- payload fairing undergo final assembly at the factory, and all subsequent ing processing is performed with the fairing halves mated. This unique approach dictates that the payload fairing will receive its final comprehensive cleaning before shipment.
REV D HPD-10041 5-35 Precision Cleaning Precision cleaning of the payload fairing inner surface is performed just Process During before installation of the pre-cleaned acoustic blankets. The payload Ground Process- fairing assembly is then double-bagged and placed in a metal container ing for shipment to Home Port. The interface skirt payload structure is similarly cleaned, double-bagged, and placed in a separate container. Extensive surface sampling has shown that standard cleaning processes will achieve cleanliness levels well below the generic requirements of 11 mg/m2 (1.0 mg/ft2) molecular residue and 0.5% particulate obscuration. Cleanliness verification is performed by close visual inspection. Direct surface sampling is available as a mission option.
Flight Hardware On arrival at Home Port, the payload accommodation hardware is un- Ground Process- bagged in the Payload Processing Facility cleanroom, and a receiving ing Environment inspection is performed. The payload accommodation is inspected and re-cleaned as required to meet strict visual cleanliness criteria before the spacecraft is encapsulated. A class 100,000 cleanroom environment is provided for the spacecraft customer during processing in the Payload Processing Facility. Cleanroom garments, materials, and janitorial services are also provided. If necessary, a mission-specific annex to the Home Port control plan is generated to incorporate any unique spacecraft contamination control requirements.
Purge Air Quality After encapsulation, the payload fairing air purge defines the spacecraft During Ground environment. The purge air is high efficiency particulate air-filtered and Processing delivered at a nominal cleanliness of class 1,000 (in accordance with FED-STD-209E). Purge air is provided continuously from the time the payload unit leaves the Payload Processing Facility through liftoff. In the case of a launch abort, the purge will be maintained during recycle events and subsequent launch attempts. The purge air quality is monitored with an airborne particle counter near the payload fairing inlet. Monitoring is discontinued only when the integrated launch vehicle is hoisted from the Assembly and Command Ship to the Launch Platform, and just before the integrated launch vehicle is erected on the launch pad. Purge air quality has averaged well under class 100 for all missions to date.
5-36 HPD-10041 REV D Helium Exposure Helium concentration in the spacecraft environment is of concern to some customers. The primary source of helium in the Sea Launch system is venting and leakage during Block DM-SL insulation blanket purging before liquid oxygen tanking. The spacecraft is protected from helium exposure by a positive pressure gradient between the payload fairing and the Block DM-SL cavity, which provides a constant flow of air through the payload structure one-way vent valves. In addition to the positive pressure airflow toward the Block DM-SL, an optional gaseous nitrogen (GN2) purge can be provided at the SC interface for further protection. GN2 purge is a non-standard option available from completion of SC mate to the SCA up to removal of the transporter-erector (at L-17 minutes).
Spacecraft Access If access to the spacecraft is required on the Launch Platform, a at Sea temporary air lock may be installed around the payload fairing access door to provide a clean zone for working and to shield the payload fairing opening from debris. Purge air from the payload fairing will provide a clean air environment within the air lock. This is a non- standard option, see Appendix B for more information.
Contamination The potential for spacecraft contamination during liftoff, ascent and During Flight injection into transfer orbit is minimized by proper hardware design and flight design. Flight hardware design considerations include surface cleanability, selection of low-outgassing materials, venting design, and containment of separation system pyrotechnics. Flight design considerations include minimization of thruster plume impingement and limiting exposure to materials outgassing during separation maneuvers.
Payload Unit The payload unit is constructed from low-outgassing materials and Materials screened using the ASTM E595 outgassing test method. Small amounts Outgassing of critical materials may not meet the screening requirement but are accepted based on other factors, such as moderate flight temperature, small surface area, or short-duration use. A full list of payload accommodation nonmetallic materials is provided in the Generic Payload Contamination Analysis.
REV D HPD-10041 5-37 Spacecraft The spacecraft is isolated from the Block DM-SL by a closeout Isolation membrane and one-way vent valves in the payload structure. These vent valves open to allow purge air to flow aft through the payload structure but are spring-loaded to close during flight. This prevents any particulate or molecular contamination from the Block DM-SL cavity from entering the spacecraft volume. The wide diameter of the interface skirt also prevents any line-of-sight view to the spacecraft from external Block DM-SL surfaces during flight.
Separation System Pyrotechnic products of the payload fairing separation system and the spacecraft separation system are contained to prevent contamination of the spacecraft. The payload fairing separation system design incorporates a pressure- containment tube, which expands to break a structural aluminum flange and separate the two halves of the payload fairing. Aluminum particles generated by breaking the flange have been characterized by extensive testing and represent a minimal contamination risk. Spacecraft adapter separation systems consist of standard clampband or separation nut designs, which feature O-ring seal containment of pyrotechnic products and debris containment for frangible bolts.
Retrorocket Zenit-3SL second-stage separation is accomplished by activating four Plumes solid propellant retrorockets located at the aft end of the stage. The retrorockets are canted fifteen degrees outboard. The wide diameter of the interface skirt blocks any direct view from the retrorocket exit plane to the spacecraft and prevents any impingement of high-velocity particles on the spacecraft. The only condensable gaseous species in the retro plume is lead (Pb), which is generated by vaporization of solid Pb particles in the fuel. Extensive plume modeling predicts a deposition of less than 5 Angstroms of Pb on most outboard vertical spacecraft surfaces. Aft facing surfaces may accumulate up to 30 Angstroms of Pb. Thin film deposition tests show that 30 Angstroms of Pb will have minimal effects on most spacecraft optical, electrical, and radio frequency properties. These data are provided in the Generic Contamination Analysis Report.
5-38 HPD-10041 REV D Contamination After spacecraft separation, the Block DM-SL performs a contamination and Collision and collision avoidance maneuver (CCAM) to prevent re-contact with Avoidance the spacecraft and to minimize contamination from outgassing and Maneuver attitude control thruster plume impingement. Several flight design parameters are constrained to limit contamination, including maximum Block DM-SL tipoff angle, and minimum separation distance from the spacecraft before major CCAM events. Major events include long orbit-separation and fuel-depletion burns of the attitude control system, and kerosene fuel tank venting. A total plume impingement limit of 1.E-05 g/cm2 is verified by analysis for each mission-unique separation maneuver.
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5-40 HPD-10041 REV D
6.0 SPACECRAFT DESIGN CONSIDERATIONS
Overview This section contains spacecraft design guides to assist the prospective customer with an initial evaluation of spacecraft compatibility with the Sea Launch system. This includes: • Mass and center-of-gravity limits • Modal frequencies • Electromagnetic compatibility • Spacecraft design verification requirements • Horizontal handling • Pressurized systems • Ordnance systems • Multiple manifests and secondary payloads
Contact Sea Contact Sea Launch if a spacecraft exceeds the design guides described Launch for Unique in the following sections or requires a unique qualification approach. Sea Requirements Launch works with each customer to achieve launch vehicle compatibility. Figure 6-1 shows the coordinate system Sea Launch uses for the Zenit- 3SL and its components.
REV D HPD-10041 6-1
+X PLU Roll
Down during powered flight, up -YPLU during horizontal launch operations
OPLU
Yaw +XUS Pitch
+Z PLU
+YPLU +Z US
+YUS OLV, OUS, OPLU ZLV, ZUS, ZPLU
OUS
YLV, YUS, YPLU +X LV VIEW A
XPLU = XUS – 4000 mm XPLU = XLV – 44652 mm YPLU = YLV = YUS ZPLU = ZLV = ZUS
OLV
+Z LV A
+YLV SHG0681874-039 Figure 6-1. Zenit-3SL Coordinate System
6-2 HPD-10041 REV D
6.1 Mass and Center-of-Gravity Limits
Definitive Several verification activities must be executed to demonstrate Zenit- Compatibility 3SL compatibility with a specific spacecraft having a given mass and Requiring center of gravity. These include: Multiple • Coupled loads analysis Verification Activities • Strength and controllability analyses • Critical clearance analyses to assess volumetric constraints • Performance calculations for the specific mission • Peaking load calculations for spacecraft having a nonuniform inter- face stiffness More information on spacecraft mass is presented in Section 3.3.
Zenit-3SL The Zenit-3SL launch vehicle can structurally accommodate spacecraft Structural with a wide range of mass and center of gravity for each spacecraft Capability adapter offered by Sea Launch. Table 6-1 provides the expected spacecraft mass and center-of-gravity limits for the Zenit-3SL when the lateral load factor on the spacecraft is 2.0 g or less (see Section 5.1 for a discussion of design quasi-static load factors on spacecraft primary structure). Conservative analyses are used to construct the mass and center-of-gravity limits (Table 6-1). Spacecraft with a higher center of gravity for a given mass may be acceptable with further analysis. Definitive compatibility of a spacecraft with the Zenit-3SL system requires all verification activities outlined above. The Zenit-3SL system is currently certified to a maximum SC mass and CG of 6100 kg and 2.0 m, respectively. Additional activities above and beyond the normal mission specific verification activities are required to certify the Zenit- 3SL system when the spacecraft mass and/or CG exceeds these limits. Please contact Sea Launch for a mission-specific Zenit-3SL compatibility assessment when the spacecraft mass and/or CG exceeds the certification limits, or its mass and CG exceeds the limits in Table 6-1.
REV D HPD-10041 6-3
Table 6-1. Expected Spacecraft Mass and Center-of-Gravity Limits Spacecraft CG, m SC mass, SCA1666 with SCA702+ (and kg SCA1194VS 0% peaking upgraded SCA1666) ≤4600 > 2.5 > 2.5 4700 2.49 2.49 > 2.5 4800 2.43 2.43 4900 2.38 2.38 2.49 5000 2.34 2.34 2.44 5100 2.29 2.29 2.39 5200 2.25 2.25 2.35 5300 2.20 2.20 2.30 5400 2.16 2.16 2.26 5500 2.12 2.12 2.22 5600 2.09 2.09 2.18 5700 2.05 2.05 2.14 5800 2.01 2.01 2.10 5900 1.98 1.98 2.07 6000 1.95 1.95 2.03 6100 1.91 1.91 2.00 6200 1.88 1.88 1.97 6300 1.85 1.85 1.94 6400 1.83 1.83 1.91 6500 1.80 1.80 1.88 6600 NA 1.77 1.85 6700 NA 1.74 1.82 6800 NA 1.72 1.79 6900 NA 1.69 1.77 7000 NA 1.67 1.74 7100 NA 1.65 1.72 7200 NA 1.62 1.69 7300 NA 1.60 1.67 Notes: ● Mass and CG limits calculated with a SC lateral load factor equal to 2.0 g. ● The Zenit-3SL system is currently certified to a maximum SC mass and CG of 6100 kg and 2.0 m, respectively. ● Improved capability generally possible using mission specific SC properties. Contact Sea Launch if SC mass or CG exceeds either the Zenit-3SL System Certification Limits or the Zenit-3SL Structural Capability Limits.
6-4 HPD-10041 REV D
Spacecraft Mass Figures 6-2 through 6-4 represent the results in table 6-1 for each of the and Center-of- spacecraft adapters offered. These figures show the expected spacecraft Gravity Limits mass and center-of-gravity limits for each of the adapters. Results are provided based upon the current Zenit-3SL structural capability as well as the current Zenit-3SL system certification limits. Definitive compatibility for any given spacecraft requires that all mission integration activities be performed. Spacecraft with mass and center-of- gravity values outside the designated acceptable region may be accept- able on further analysis. Increased Zenit-3SL structural capability can be realized with strength upgrades to the spacecraft adapters and/or the as- sociated payload support structure for the adapters (see Figure 2-7). The longitudinal center-of-gravity location of the spacecraft in figures 6-2 through 6-4 refers to the spacecraft separation plane. The separation plane for the SCA1666 and SCA702+ spacecraft adapters is at payload unit x-axis station XPLU = 2.10 m (82.7 in). Similarly, the SCA1194VS spacecraft adapter separation plane is at XPLU = 1.97 m (77.6 in). The maximum radial center-of-gravity offset (nominal plus uncertainty) from the spacecraft longitudinal centerline axis should be less than 5.0 cm (2.0 in). Radial offsets exceeding this require further analysis (e.g., radial offsets of 6.0 cm (2.4 in) have been shown to be acceptable on a case-by-case basis). Figure 6-2 provides SC mass and CG limits for the Zenit-3SL using the SCA1194VS adapter with typical load peaking at the SC interface. The SCA1194VS adapter has been qualified for SC up to 8000 kg with a CG at 2.0 m. Figure 6-3 provides SC mass and CG limits for the Zenit-3SL us- ing the SCA1666 adapter. Results are provided for 0% and 30% peaking. If the SCA1666 spacecraft adapter is upgraded to realize the full struc- tural capability of the Zenit-3SL, then higher spacecraft mass is permit- ted, and the penalty imposed on mass and center-of-gravity limits due to load peaking is eliminated. Figure 6-4 provides SC mass and CG limits for the Zenit-3SL using the SCA702+ adapter.
REV D HPD-10041 6-5
8000 Zenit-3SL Structural Capability with SCA1194VS
7000 Current Zenit-3SL System Certication Limits
6000
5000
4000 Notes: 3000 • Mass and CG limits calculated with a SC lateral load factor equal to 2.0 g. Spacecraft mass, kg • Improved capability generally possible using mission specific 2000 SC properties. Contact Sea Launch if SC mass or CG exceed either the Zenit-3SL System Certification Limits or the Zenit-3SL Structural Capability Limits. 1000
0 0.00 0.50 1.00 1.50 2.00 2.50 3.00 SHG0681874-078 Spacecraft CG, m Figure 6-2. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA1194VS 8000
7000
6000
5000
Maximum Expected Zenit-3SL Structural Capability with upgraded SCA1666 4000 Zenit-3SL Structural Capability with SCA1666 and 0% peaking Zenit-3SL Structural Capability with SCA1666 3000 and 30% peaking Spacecraft mass, kg Current Zenit-3SL System Certification Limits
2000
Notes: 1000 • Mass and CG limits calculated with a SC lateral load factor equal to 2.0 g.
0 0.00 0.50 1.00 1.50 2.00 2.50 3.00 Spacecraft CG, m SHG0681874-079 Figure 6-3. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA1666 Spacecraft Adapter
6-6 HPD-10041 REV D
8000 Zenit-3SL Structural Capability with SCA702+ 7000 Current Zeint-3SL System Certification Limits
6000
5000
4000
Spacecraft mass, kg 3000
Notes: • Mass and CG limits calculated with a SC lateral load factor 2000 equal to 2.0 g. • Improved capability generally possible using mission specific SC properties. Contact Sea Launch if SC mass or CG exceed either the Zenit-3SL System Certification Limits or the Zenit-3SL 1000 Structural Capability Limits.
0 0.00 0.50 1.00 1.50 2.00 2.50 3.00 Spacecraft CG, m SHG0681874-080 Figure 6-4. Expected Spacecraft Mass and Center-of-Gravity Limits With SCA702+ Spacecraft Adapter
6.2 Modal Frequencies
Frequency Limits Sea Launch can accommodate spacecraft with longitudinal and lateral fundamental frequencies below those commonly offered in the industry. As a guide, the spacecraft longitudinal fundamental frequency should be ≥20 Hz and its lateral fundamental frequency should be ≥8 Hz.
6.3 Electromagnetic Compatibility
Radiated To avoid electromagnetic interference with launch vehicle systems, Emissions at the spacecraft radiated emissions at the launch vehicle/spacecraft interface Launch Vehicle must not exceed the levels shown in Figure 6-5. Spacecraft radio frequency system characteristics will be required from the spacecraft manufacturer for integrated analysis and for the required Federal Communications Commission special temporary authority license, which is necessary for Home Port radio frequency testing.
REV D HPD-10041 6-7
160
140 140
120
100
80
60 Electric field, dBµV/m
40
20 Kvant receiver 762-776 MHz 0 10 100 1,000 10,000 100,000 Frequency, MHz SHG0681874-040 Figure 6-5. Maximum Spacecraft Intentional Radiated Emissions at the Launch Vehicle/Spacecraft Interface
Lightning Sea Launch uses various sensors, satellite imagery, radar and Protection weather prediction systems (Figure 6-6) (e.g., weather and lightning occurrence predictions) to ensure a lightning-free launch environment. Sea Launch provides its customers with notification of lightning occurrences (e.g., lightning already having occurred within a certain distance) should they wish to de-cable their spacecraft during spacecraft processing, transit or checkout.
6-8 HPD-10041 REV D
Lightning/weather service PPF in Spacecraft transfer PLU ACS on Spacecraft transfer ILV HP in LP on Spacecraft erect and rollout Dry S/C transitto launch site Launch site 1 Global atmospheric service Off-site, real-time warning 2 Boltek lightning monitor (ACS) On-site, real-time warning 3 ACS-02, PLU transfer to ACS Procedure identifying weather/lightning constraints 4 PLU/ILV transfer RR MD readiness review includes RC-03 WX briefing 5 MP-01, task 15, PLU transfer to ACS MD checklist prior to transfer 6 XFR-01, ILV transfer Procedure identifying weather/lightning constraints 7 MP-01, task 18, ILV transfer to LP MD checklist prior to transfer 8 LP-06, PLU launch countdown Procedure identifying weather/lightning constraints 9 MP-01, task 19, ILV dry rollout and erect MD checklist prior to rollout 10 RC-03 WX briefing Procedure performing the weather briefing 11 Boltek lightning monitor (LP) On LP, real-time warning 12 Launch Readiness Review MD readiness review includes RC-03 WX briefing 13 Launch Commit Criteria Jointly agreed-to weather/lightning commit criteria 265807J6-332 Figure 6-6. Lightning and Weather Services Used to Determine or Forecast Lightning During Launch Campaign
Radio Frequency There are very few areas within the Sea Launch segments that present Hazards electromagnetic radiation hazards to personnel, ordnance or fuels. A few high-power radio frequency transmitters can produce radio frequency hazard levels in areas within the antenna’s main beams. These radio frequency hazard locations and conditions are identified in the Sea Launch EMC Control Plan. Safety personnel have identified the radio frequency hazard areas as controlled-access areas. Safety training is discussed in Section 11.
Electromagnetic In order to ensure compatibility with the Sea Launch system, the Interference/ spacecraft contractor will provide component-, subsystem- and Electromagnetic equipment-level electromagnetic interference and electromagnetic Compatibility Test compatibility qualification test plans and test reports. Report
REV D HPD-10041 6-9
6.4 Spacecraft Design Verification Requirements
Flexibility to Meet Design verification requirements for the spacecraft have been engineered Customer Needs and selected by Sea Launch to ensure mission success. These requirements are not rigid and can be modified with Sea Launch approval to meet the verification needs of a spacecraft. In addition, detailed safety design constraints on the spacecraft are contained in the Sea Launch Safety Manual.
Customizable Strength requirements acceptable to Sea Launch are set forth in the Strength following section but can be modified to meet customer needs with Sea Requirements Launch approval.
Factors of Safety Minimum yield and ultimate factors of safety for the spacecraft depend on the testing option chosen and the design heritage of the hardware. The
minimum factors of safety and test levels for several test options are set forth in Table 6-2. Note that the minimum ultimate factor of safety set forth is 1.3. However, for qualification or proof tests, an ultimate factor of safety equal to 1.25 on limit loads is acceptable for structural components having a flight heritage with Sea Launch or when a model uncertainty factor no less than 1.05 is used to determine limit load.
Table 6-2. Factors of Safety and Test Options
Factors of safety Test Test option Yield Ultimate level Test success criteria Qualification test (test of 1.0 1. 3 • 1.0 • No detrimental dedicated article to ultimate loads) deformation at 1.0 • 1.25 • No failure at 1.25 Proto-qualification test (test of 1.25 1.4 1.25 No detrimental article used subsequently for flight or deformation or system test) misalignment Qualification by analysis (test of 1.6 2.0 N/A N/A article not required) Proof test (performed on each flight 1.1 1.3 1.1 No detrimental vehicle) deformation
6-10 HPD-10041 REV D
Test-Verified The spacecraft must have positive margins of safety for yield and Model Required ultimate loads. These loads are generated from limit loads using factors for Final Coupled of safety no less than those displayed in Table 6-2. Limit loads for the Loads Analysis transient flight events are calculated by Sea Launch in a coupled loads analysis and provided to the spacecraft customer. Limit loads for the
quasi-static events are calculated using the Sea Launch quasi-static load factors for ground handling, transportation and flight. The coupled loads analysis limit loads used for verification must be generated using a test- verified model for the spacecraft. Model verification of the spacecraft can be performed by either modal survey or sine test.
Fatigue Factor of Compliance to the Sea Launch fatigue environment is demonstrated by Safety showing positive margin to the fatigue spectrum (Section 5, Table 5-1) after increasing the number of cycles at each load level by a factor of four.
Customizable Test Sea Launch requires that the spacecraft structural capability be Requirements demonstrated by the spacecraft manufacturer. The Sea Launch qualification requirements on the spacecraft are standard industry practice, but can be modified to meet specific spacecraft needs. Structural testing on the spacecraft generally depends on design heritage. Unique qualification tests can be developed by the spacecraft customer to account for design heritage. Sea Launch will work with the spacecraft customer by evaluating customer-proposed testing to support the spacecraft integration process. The following tests are accepted by Sea Launch for demonstrating structural compliance. If the qualification approach that best suits the spacecraft is not discussed here, Sea Launch will evaluate any proposed alternative to ensure spacecraft compatibility.
Modal Survey Test A coupled loads analysis will be performed with a test-verified spacecraft model before flight to verify that the maximum expected limit loads for the spacecraft and the testing performed on the spacecraft structure are in compliance with Table 6-2. Model verification of the spacecraft can be performed by either modal survey or sine test. Repeat- mission spacecraft do not require further testing.
REV D HPD-10041 6-11
Static Loads Test A spacecraft static loads and/or sine vibration test must be performed by the spacecraft manufacturer for spacecraft structure. The extent of the testing depends on the heritage of the spacecraft structure but, ultimately, spacecraft compliance to Table 6-2 must be demonstrated.
Sine Vibration The sine vibration test levels for qualification, proto-qualification, proof Testing and acceptance are presented in Table 6-3. Qualification testing is for a dedicated test article and proto-qualification testing is for the first-flight article. Proof testing is for spacecraft that have not been qualification or proto-qualification tested for first flight and must be proof tested for every flight. Once a spacecraft has been qualification or proto- qualification tested, acceptance sine testing on all future repeat spacecraft need not be performed. Acceptance testing is generally performed to demonstrate workmanship and is an option available to the spacecraft customer for minor spacecraft design changes for spacecraft that have not flown previously on Sea Launch. The extent of testing depends on the heritage of the spacecraft structure and degree of static testing but, ultimately, spacecraft compliance with Table 6-2 must be demonstrated by the spacecraft manufacturer. Notching may be used to prevent excessive loading of the spacecraft structure. However, the resulting sine vibration environment with notching should not be less than a test factor level times the equivalent sine vibration level. This is determined from coupled loads analyses provided by Sea Launch. The test factor levels depend on the test options chosen and are provided in Table 6-3.
6-12 HPD-10041 REV D
Table 6-3. Sine Vibration Test Levels and Duration
Test parameters Proto- SC axis tested Frequency Qualification qualification Proof test Acceptance range test levels levels levels levels Longitudinal 5 to 100 Hz 0.75 g 0.75 g 0.66 g 0.6 g Lateral 5 to 100 Hz 0.75 g 0.75 g 0.66 g 0.6 g Sweep rate 2 octaves/ min 4 octaves/min
Acoustic Testing Qualification of a spacecraft type is required. In place of a qualification test, a proto-qualification test may be run on the first unit. Acoustic testing to proto-qualification or acceptance levels and durations is required for every mission. The acoustic margins and duration for qualification, proto-qualification and acceptance tests are set forth in Table 6-4.
Table 6-4. Spacecraft Acoustic Margins and Test Duration Test Margin Duration, min Qualification +3 dB over acceptance 2 Proto-qualification +3 dB over acceptance 1 Acceptance Maximum expected 1 acoustic level
Shock The spacecraft should be compatible with the shock environment in Qualification Section 5.4 with a 3-dB margin. The shock environment in Section 5.4 represents the maximum expected environment with no added margin. Analysis, similarity and/or test can demonstrate qualification.
REV D HPD-10041 6-13
Adapter For the first mission of a spacecraft type, a matchmate test is required. Matchmate and The test is performed between the spacecraft and the adapter flight Fit-Check Test hardware at the spacecraft manufacturer’s facility. A shock transmission test is performed with the mated spacecraft and adapter. This test defines the shock transmission losses across the interface. Following this test, a spacecraft separation test is conducted. This test is includes firing the spacecraft separation ordnance to determine the shock levels on both sides of the spacecraft interface. Repeat missions incorporate an adapter fit check at the start of processing flow at the Sea Launch Payload Processing Facility. These tests include both electrical and mechanical mates (Figure 6-7).
265807J6-333 Figure 6-7. Electrical and Mechanical Mates
6.5 Horizontal Handling
Processing While Spacecraft systems and procedures must be compatible with operations Encapsulated that place the encapsulated spacecraft in a horizontal, cantilevered attitude for transportation and handling. Access to the spacecraft during this period of operations is through the payload fairing access doors. Door location is described in Section 7.
6-14 HPD-10041 REV D
6.6 Pressurized Systems
Design and Safety Design and verification of spacecraft flight hardware pressurized vessels, Requirements structures and components must be in accordance with recognized aerospace industry design guidelines. These pressurized systems should also be based on the Sea Launch system payload environment. The design must protect the launch system and personnel before launch and protect the launch system during flight from damage due to pressure system failure. Such criteria as operating pressures, stress levels, fracture control, burst factor, leak-before-burst factor, material selection, quality assurance, proof-pressure testing, and effects of processing and handling in the horizontal orientation should be considered. Design details are required to be included in the documentation to support regulatory agency requirements for each launch.
6.7 Ordnance Systems
Recognized Ordnance systems on board spacecraft for operation of propulsion, Standards and separation and mechanical systems must be designed in accordance with Regulations recognized standards and regulations. These systems must preclude inadvertent firing when subjected to Sea Launch–specified shock, vibration, thermal and electromagnetic environments. Ordnance devices must be classified in accordance with applicable federal and state codes and meet U.S. Government regulations for transportation and handling. Design for initiation of ordnance in the system must incorporate more than one action; no single failure may result in ordnance device activation. Use of a safe-and-arm-device is recommended; however, other techniques may be considered with adequate justification. System design and ordnance classification documentation is required to support regulatory agency requirements on the launch mission.
6.8 Multiple Manifests and Secondary Payloads
Overview Sea Launch is currently assessing the requirements necessary to offer launch services for multiple manifests and secondary payloads. These types of missions will be evaluated on a case-by-case basis.
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6-16 HPD-10041 REV D 7.0 INTERFACES
7.1 Mechanical Interfaces
Overview Characteristics of the mechanical interfaces covered in this section include: • Payload fairing • Access doors • Spacecraft adapters
Payload Fairing Configuration for the payload fairing is illustrated in Section 2. The Sea Launch payload fairing accommodates the spacecraft static envelope, as shown in Figure 7-1. Excursions to the defined spacecraft envelope may be accommodated on a case-by-case basis. Figures 7-2, 7-3, 7-4, and 7-5 define the spacecraft envelope available below the separation plane for the 1194VS, 1666, 702+, and 937S spacecraft adapters, respectively.
ø 1300 mm (4 ft, 3 in)
ø 2500 mm (8 ft, 2 in) 2250 mm 9312 mm (7 ft, 5 in) (30 ft 7 in)
ø 3750 mm 5712 mm (12 ft, 4 in) (18 ft 9 in) 702+ – 775 mm (30.51 in) 1666 – 775 mm (30.51 in) 937S – 772 mm (30.4 in) 1194VS – 641 mm (25.24 in)
Separation Plane
Envelope inside SCA will vary Envelope Edge slightly depending on SCA interface SHG0681874-041.1 Figure 7-1. Spacecraft Static Envelope
REV D HPD-10041 7-1 Ø 3750 (147.64) 51 (2.00)
Separation Plane
541 (21.30) 642 (25.27)
653 Ø (25.73) 863 Ø (33.93) 1194 Ø (47.01) 1420 Ø (55.90) mm (in) Ø 2826 (111.25) SHG0681874-006 Figure 7-2. Spacecraft Static Envelope SCA1194VS Adapter
3750 Ø (147.64) 59 (2.32) 450 Separation (17.72) Plane
615 775 (24.21) (30.51)
1120 Ø (44.09) 1374 Ø (54.17) 1666 Ø (65.59) 2322 Ø (91.41) mm 2826 (in) Ø (111.25) SHG0681874-007 Figure 7-3. Spacecraft Static Envelope SCA1666 Adapter
7-2 HPD-10041 REV D Ø 3750 (147.64)
269 Separation (10.60) Plane
775 (30.5)
500 (19.69) 1521 Ø (59.9) 1815 Ø (71.44) mm 2826 (in) Ø (111.25) SHG0681874-008 Figure 7-4. Spacecraft Static Envelope SCA702+ Adapter
3750 Ø (147.6)
338 Separation (13.3) Plane
773 (30.4)
673 (26.5) 580 Ø (22.8) 724 Ø (28.5) 904 Ø (35.6) mm 1130 (in) Ø (52.36) SHG0681874-009 Figure 7-5. Spacecraft Envelope SCA937S Adapter
REV D HPD-10041 7-3 Access Doors The standard Sea Launch payload fairing provides up to two access doors, one door per fairing half, with a diameter of 610 mm (24 in). Figure 7-6 shows the zones on the fairing surface where doors may be located, as follows: • Access doors centered between station 2260.6 mm (89 in) and station 2565.4 mm (101 in) must be defined by L-8 months. • Access doors centered at higher stations must be defined by L-12 months. As an option, the payload fairing can accommodate additional access doors beyond the standard quantity of two.
7-4 HPD-10041 REV D 360 deg 180 deg
331 deg 180 deg Access door opening 209 deg 610.0 [24] diameter 360 deg Door center in this zone must be defined by L-8
Door center in this zone STA 2260.6 [89] must be defined by L-12
STA 2565.4 [101]
601, 702 separation plane STA 5969.0 [235] STA 2099.8 [82.7]
1194 separation plane Flat Pattern of 180-360 Half of Fairing STA 1967.0 [77.4] Note: Door shown in closest location to edge; dimensions in mm [in].
270 deg Zone for center of door
331 deg 209 deg
180 deg 0/360 deg
29 deg 151 deg Longitudinal separation plane
90 deg Zone for center of door Top View of Fairing 110 deg Access door opening 610.0 [24] diameter Door center in this 151 deg zone must be defined by L-12 1194 separation plane STA 1967.0 [77.4] STA 5969.0 [235] 1666, 702 separation plane Door center in this STA 2099.8 [82.7] zone must be defined by L-8 STA 3962.4 [156] STA 2565.4 [101] STA 2489.2 [98] STA 2260.6 [89]
0 deg 29 deg 131 deg 180 deg Flat Pattern of 0-180 Half of Fairing Note: Door shown in closest location to edge, dimensions in mm [in]. 0 deg 180 deg SHG0681874-081 Figure 7-6. Payload Fairing Locations for Access Doors
REV D HPD-10041 7-5 Payload Structure The 45-deg conic design of the Sea Launch payload structure maximizes and Spacecraft the height available to the spacecraft and accommodates the full range of Adapters existing and planned spacecraft adapters. The following sections define the standard spacecraft adapters currently offered. All Sea Launch spacecraft adapters incorporate flight proven spacecraft separation systems developed and manufactured by Saab Space, the industry leader in spacecraft separation systems.
SCA1666 and The SCA1666 has a diameter of 1666 mm (65.59 in) at the spacecraft SCA1666 MVS interface and uses a Marmon clamp separation system. This adapter accommodates the Boeing 601-type spacecraft. The SCA 1666 MVS accommodates the Mitsubishi Electric Company DS2000 type spacecraft. The SCA 1666 MVS incorporates a clampband soft separation system developed by Saab Space and uses the principle of the fast-acting shockless separation nut developed by Starsys Research. Six pushoff springs are used during separation to impart an initial delta velocity with respect to the launch vehicle. The spacecraft adapter also has the capability to impart spacecraft spin about an axis transverse to the launch vehicle longitudinal axis. In principle, this is achieved by limiting the stroke of adjacent springs with respect to the stroke of the opposing springs. In addition to the pushoff springs, there are two electrical disconnects and a clampband impulse that are integral to spacecraft separation dynamics. The major elements of this clampband-type adapter are shown in Figure 7-7. The SCA1666 and the SCA 1666 MVS are designed to accommodate a 6500-kg (14,330 lb) spacecraft with a center of gravity at 1.8 m (70.9 in) with 0% peaking.
7-6 HPD-10041 REV D Separation Springs Umbilical Connection
Clampband Catcher
SAS Structure
Clampband
Vent Holes
SCA—PLS Interface Bolt Cutter SHG0681874-002.1 Figure 7-7. SCA1666 Spacecraft Adapter
SCA702+ and The SCA702+ has a diameter of 1664 mm (65.51 in) at the spacecraft SCA702+GEM interface and is attached to the spacecraft using four bolts. This adapter accommodates the Boeing 702-model spacecraft. The SCA702+ is qualified to accommodate up to a 7300-kg (16,094 lb) spacecraft with a center of gravity at 1.67 m (5.47 ft) and includes the effects of spacecraft flexibility at the interface. The separation system uses four pushoff springs to impart both an initial delta velocity with respect to the launch vehicle and a spacecraft spin rate about an axis transverse to the launch vehicle longitudinal axis. In principle, this relative motion is achieved by limiting the stroke of two adjacent springs with respect to the stroke of the other two opposing springs. In addition to the pushoff springs, there are two electrical disconnects, which are integral to spacecraft separation dynamics (Figure 7-8). The SCA702+GEM is an SCA702+ with minor modifications to account for slightly different geometry at the spacecraft interface.
REV D HPD-10041 7-7 Shear Cone
Separation Spring
Umbilical Connection
Separation Nut
Vent Holes
SCA—PLS Interface SCA Structure
SHG0681874-003 Figure 7-8. SCA702+ Spacecraft Adapter
7-8 HPD-10041 REV D SCA1194VS The SCA1194VS has a diameter of 1194 mm (47 in) at the spacecraft interface and uses a Marmon clamp separation system. The SCA1194VS is designed to accommodate an 8000-kg (17,637 lb) spacecraft with a center of gravity at 2.0 m (6.6 ft) with 10% peaking. The spacecraft separation system uses from eight to fourteen pushoff springs to impart both an initial delta velocity with respect to the launch vehicle and a spacecraft spin rate about an axis transverse to the launch vehicle longitudinal axis. In principle, this relative motion is achieved by limiting the stroke of adjacent springs with respect to the stroke of the opposing springs. In addition to the pushoff springs, there are two electrical disconnects and a clampband impulse that are integral to spacecraft separation dynamics. The major elements of this clampband- type adapter are shown in Figure 7-9. The SCA1194VS (Figure 7-9) incorporates a clampband soft separation system developed by Saab Space and uses the principle of the fast-acting shockless separation nut from Starsys Research. The SCA1194VS consists of a preshaped aluminum band with titanium end fittings. The adapter uses only one band joint with a nominal clampband pre-tension of 60 kN or 40 kN, selectable by the customer. The clampband opening device is a reusable, and thus testable, system consisting of a flywheel housing and an energetically actuated pin puller. The installation of the SCA1194VS is easier and faster than the previous SCA1194 model. The SCA1194VS requires only one band joint and tensioning operation instead of two, as required with the SCA1194. Additionally, the required access zone is on only one side of the spacecraft.
REV D HPD-10041 7-9 Clampband Separation Spring
Umbilical Connection
SCA Structure
Vent Holes Clampband Opening Device SCA—PLS Interface Pin Puller
SHG0681874-004 Figure 7-9. SCA1194VS Spacecraft Adapter
SCA937S The SCA937S has a diameter of 937 mm (36.89 in) at the spacecraft interface and uses a Marmon clamp separation system. The SCA937S is designed to accommodate a 4500-kg (9,920 lb) spacecraft with a center of gravity at 1.5 m (4.9 ft). The spacecraft separation system uses four pushoff springs to impart an initial delta velocity with respect to the launch vehicle. In addition to the pushoff springs, there are two electrical disconnects and a clampband impulse that are integral to spacecraft separation dynamics. The major elements of this clampband-type adapter are shown in Figure 7-10. The SCA937S (Figure 7-10) also incorporates the clampband soft separation system developed by Saab Space and uses the principle of the fast-acting shockless separation nut from Starsys Research. This clampband opening device is a reusable, and thus testable, system consisting of a flywheel housing and energetically actuated pin puller. The SCA937S consists of a preshaped aluminum band with titanium end fittings. The adapter uses only one band joint with a nominal clampband pre-tension of 30 kN or 40 kN.
7-10 HPD-10041 REV D Umbilical Bracket
Clampband Separation Spring
Clampband Catcher
Pin Puller and Clampband Opening
Device Vent Holes GN2 Bracket SCA—PS Interface SCA Structure SHG0681874-005 Figure 7-10. SCA937S Spacecraft Adapter
7.2 Electrical Interfaces
Overview Characteristics of the electrical interfaces are covered in the following sections, including: • Hard-line link (umbilical) • Radio frequency link • External communication • In-flight interfaces • Electrical power
Spacecraft/ The launch vehicle adapter standard configuration accommodates two Adapter In-Flight SAE-AS81703 style connectors (37 pins or 61 pins). Additional required Electrical connectors can be accommodated and will be evaluated on a case-by- Disconnects case basis.
REV D HPD-10041 7-11 Hard-Line Link Spacecraft hard-line link electrical interconnections between the (Umbilical) spacecraft adapter in-flight disconnects and the spacecraft electrical ground support equipment are provided by two functionally similar but physically separate umbilical cable paths: • Interface skirt to the Payload Unit • T-0 (through the Launch Vehicle) Both umbilical paths are available for testing during integration checkout: • During transit to the launch site, only the interface skirt umbilical is available. • On the launch pad (from approximately L-24 hours to L-8 hours), both the T-0 umbilical and the high current connections of the interface skirt umbilical are available. • At approximately L-6 hours, the interface skirt umbilical is disconnected. • The T-0 umbilical remains connected until liftoff. Table 7-1 and Figure 7-11 depict umbilical configurations during spacecraft encapsulation, integrated launch vehicle integration, transit phase and on-pad operations. The voltages are referenced at the Customer Control Room DIN Panel.
7-12 HPD-10041 REV D
Available Table 7-1. Umbilical Configurations Umbilical Functions Available Umbilical Umbilical Cable Paths Functions Signal Cables Power Cables Interface Skirt Interface skirt • 10 twisted • Up to 150 V Umbilical in PPF shielded pairs Availability • Up to 94 A • 10 twisted • 14 circuits at 5 pairs A each • Up to 250 mA • 8 circuits at 3 at 50 VAC or A each 150 VDC Interface skirt in • 10 twisted • Up to 150 V ACS Room 101 shielded pairs • Up to 94 A • 10 twisted • 14 circuits at 5 pairs A each • Up to 250 mA • 8 circuits at 3 at 50 VAC or A each 150 VDC Interface skirt in • 10 twisted • Up to 150 V LP Hangar shielded pairs • Up to110 A • 10 twisted • 22 circuits at 5 pairs up to 250 A each mA at 50 VAC or 150 VDC Interface skirt Signals not • Up to 150 V on LP after available via rollout and interface skirt • Up to 110 A before L-8 hr • 22 circuits at 5 A each Interface skirt Signals not • Power not on LP L-8 hr available via available via to launch interface skirt interface skirt
REV D HPD-10041 7-13
Table 7-1. Umbilical Configurations (continued) Available Umbilical Umbilical Cable Paths Functions Signal Cables Power Cables T-0 Umbilical T-0 in PPF • 10 twisted • Up to 150 V Availability shielded pairs • Up to 58 A • 10 twisted • 17 circuits at pairs 2 A each • Up to 250 mA • 24 circuits at at 50 VAC or 1 A each 150 VDC T-0 in ACS • 10 twisted • Up to 150 V Room 101 shielded pairs • Up to 58 A • 10 twisted • 17 circuits at pairs 2 A each • Up to 250 mA • 24 circuits at at 50 VAC or 1 A each 150 VDC T-0 in LP Signals not • Power not Hangar available via available via T-0 T-0
T-0 on LP after • 10 twisted • Up to 150 V ILV rollout shielded pairs until Launch • Up to 58 A • 10 twisted • 17 circuits at pairs 2 A each • Up to 250 mA • 24 circuits at at 50 VAC or 1 A each 150 VDC
7-14 HPD-10041 REV D Encapsulation ILV Integration
SC SC ILV
IS T-0 IS T-0 umbilical umbilical umbilical umbilical
SC EGSE SC EGSE
Payload Processing Facility Assembly and Command Ship
Transit On Pad
SC
SC ILV
IS T-0 umbilical umbilical ILV
IS T-0 SC EGSE umbilical umbilical
Launch Platform SC EGSE External comm network Launch Platform Line of sight link and External comm network
SC EGSE SC EGSE
Assembly and Command Ship Assembly and Command Ship 265807J6-227 Figure 7-11. Umbilical Configurations During Launch Processing Flow
REV D HPD-10041 7-15 Ground Test Spacecraft hard-line link electrical interconnections between the Cabling (SC EGSE spacecraft electrical ground support equipment and the spacecraft control to SCR wall room connector panels are required. There are two functionally similar panels) but physically separate umbilical cable paths available at these connector panels: • Interface skirt to the Payload Unit • T-0 (through the Launch Vehicle) Both umbilical paths are available for testing during integration checkout: • During transit to the launch site, only the interface skirt path is available. • On the launch pad (from approximately L-24 to L-8 hours), the T-0 umbilical and the high current connections of the interface skirt umbilical are available. • At approximately L-6 hours, the interface skirt umbilical is disconnected. • The T-0 umbilical remains connected until liftoff. Table 7-1 and Figure 7-11 depict umbilical configurations during spacecraft encapsulation, integrated launch vehicle integration, transit phase and on-pad operations.
7-16 HPD-10041 REV D
Spacecraft There are five configurations in which the spacecraft customer verifies Umbilical Test umbilical wiring to their spacecraft simulator: Adapters 1. Interface skirt umbilical path from the Spacecraft Control Room (SCR) in the Payload Processing Facility to the SC/PLA interface. The spacecraft simulator will be attached to the SC/PLA interface using customer-provided test cables. 2. T-0 umbilical path from the SCR in the Payload Processing Facility to the SC/PLA interface. The spacecraft simulator will be attached to the SC/PLA interface using customer-provided test cables. 3. T-0 umbilical path from the SCR in the Assembly and Command Ship to the top of the Block DM-SL. The spacecraft simulator will be attached to the Block DM-SL interface using a combination of Sea Launch-provided breakout boxes and a customer-provided T-0 test cable harness. This T-0 harness simulates the PLA wire harness connection from the SC/PLA interface to the Block DM-SL interface. 4. Interface skirt umbilical path from the SCR in the Launch Platform to the end of the interface skirt ground test cables. The spacecraft simulator will be attached to these cables using a combination of Sea Launch-provided breakout boxes and a customer-provided ISPS test cable harness. This ISPS harness simulates the PLA wire harness connection from the SC/PLA interface to the interface skirt. 5. T-0 umbilical path from the SCR in the Launch Platform to the end of the T-0 umbilical test panel located in Room 212a on the Launch Platform. The spacecraft simulator will be attached to a simulated Block DM-SL interface to SC/PLA interface cable harness using a combination of Sea Launch-provided breakout boxes and customer- provided test cables. This T-0 harness simulates the PLA wire harness connection from the SC/PLA interface to the Block DM-SL interface.
REV D HPD-10041 7-17 Spacecraft Radio Spacecraft RF links are re-radiated between the internal fairing envelope Frequency Links and the external integrated launch vehicle environment using an amplified bi-directional avionics RF system. The Sea Launch term for this re-radiation system is “Rerad.” See Figure 7-12. The Rerad System provides bi-directional RF command and telemetry links between the spacecraft and spacecraft electrical ground support equipment when the spacecraft is located in the Payload Processing Facility. Sea Launch support of spacecraft RF links while the spacecraft is located in the Assembly and Command Ship and in the hanger of the Launch Platform can be evaluated on a case-by-case basis. The Rerad System and the Spacecraft Radio Frequency Support System (SCRFSS) provide bi-directional RF command and telemetry links between the spacecraft and spacecraft electrical ground support equipment when the spacecraft is erected on the Launch Platform in launch configuration. See Figure 7-13. The launch configured links are tested at Home Port prior to departure for the launch site. The Rerad/SCRFSS support spacecraft command and telemetry frequencies between 3 GHz and 18 GHz. Operational capabilities outside this range may be accommodated. Sea Launch support of frequencies between 18 GHz to 40 GHz can be evaluated on request.
Pickup antennas
A 75-mm foam with RF absorber v io ry n t d ic e n s m le e mma T o C
Telemetry and command antennas
Telemetry
Spacecraft Reradiating antennas
Avionics Command
265807J6-228 Figure 7-12. Re-Radiation of Spacecraft Signal Through Fairing
7-18 HPD-10041 REV D Command and telemetry at spacecraft frequencies
Command and telemetry at marine link frequencies
Steering Gear Thrus Room t ST-16 BW Room ST-18 BW ST-20 BW ST-1 5 BW ST-13 BW ST -1 1 BW ST-9 BW ST-1 5 BW ST-5 BW ST-3 BW ST-1 BW
Command and telemetry at spacecraft frequencies
265807J6-229 Figure 7-13. Spacecraft Radio Frequency Signal Between Vessels
External The external communications network provides worldwide connectivity Communications in support of Sea Launch operations, including links between the Network Assembly and Command Ship and the Launch Platform. This network provides serial digital and Ethernet connectivity for spacecraft contractor and customer use. During transit and at the launch site, some links are available by satellite relay. When vessel position permits, the link is available directly by radio frequency line of sight. The capabilities of this network include: • Spacecraft command and telemetry hard-line links use the external communications network between vessels. This is illustrated in Figure 7-14. • Sea Launch provides communications connectivity using a ground station in Brewster, Washington. This connectivity can be established with a customer-procured data line between the Brewster ground station, the spacecraft mission ground station and/or the spacecraft factory. Sea Launch also provides an Internet connection.
REV D HPD-10041 7-19 Intelsat 180°
External Network
Brewster Ground Station C us tom er P ro Lin vid k ed
Spacecraft Internet SC EGSE Customer Facility
Optional Internet Homeport Spacecraft Customer Facility
SC EGSE SC EGSE Line-Of-Sight Digital Network
Assembly andand Launch CommandComand ShipShip Platform SHG0681874-042.1 Figure 7-14. Sea Launch External Communications Network
7-20 HPD-10041 REV D
Available In-flight Sea Launch provides several different types of in-flight interfaces to the Interfaces spacecraft. Table 7-2 provides brief descriptions of the in-flight telemetry, command and separation breakwire interfaces.
Table 7-2. In-Flight Telemetry, Command and Separation Breakwire Interfaces When Interface Description provided Serial PCM The launch vehicle onboard data acquisition system can telemetry accommodate two serial data streams from the spacecraft, up to 2 Kbps combined. This spacecraft data is interleaved within the launch vehicle telemetry downlink. It is recovered and provided to the spacecraft or customer as a part of the postflight data processing function. Serial data PCM codes of NRZL (data and clock/RS422) are accommodated. Other PCM coding schemes are optional and Provided will be evaluated on a case-by-case basis. postflight Analog Four launch vehicle analog telemetry channels are available for telemetry spacecraft use. These channels accommodate data input at ≥400 samples per second. Digital The launch vehicle telemetry system accepts up to eight discrete (bi- telemetry level) inputs from the spacecraft. In-flight The launch vehicle can provide up to four redundant (primary and commands backup) pulse-type commands for spacecraft use. The commands can be configured in an application of 28 V or as contact closures. Pulse duration is 50 ms. Separation The spacecraft is required to provide continuity loops for separation Provided breakwires monitoring by the launch vehicle. At a minimum, one loop must be in real time provided for each separation connector, but two loops per connector are preferred. The launch vehicle will provide continuity loops for use by the spacecraft as defined by the spacecraft separation connector pin allocations.
Electrical Power Sea Launch provides customer electrical ground support equipment with power from uninterruptible power supplies at Home Port and on the ACS and LP. Either 50 or 60 Hz power is available. This launch-critical power is fault tolerant and isolated from other power sources.
Electrical Each Sea Launch facility provides electrically conductive surfaces Grounding (copper plates or threaded studs) for connecting the spacecraft electrical ground support equipment to facility ground.
REV D HPD-10041 7-21
Electrical Bonding Sea Launch provides a spacecraft adapter-to-launch vehicle interface bond of ≤10 mΩ. In addition, the spacecraft-to-spacecraft adapter interface bond should be ≤10 mΩ.
7-22 HPD-10041 REV D 8.0 SPACECRAFT INTEGRATION AND LAUNCH OPERATIONS
Overview Sea Launch optimizes the convenience of West Coast United States spacecraft processing with a true equatorial launch site. There are three distinct phases of operations: • Phase I takes place in the Payload Processing Facility, at Sea Launch Home Port in Long Beach, California. This phase starts with the delivery of the spacecraft to Home Port, and includes final processing of the spacecraft, stand-alone tests, ordnance installation fueling and encapsulation with the payload fairing. • Phase II also takes place at Home Port on the Assembly and Command Ship. This phase includes mating the encapsulated spacecraft to the Zenit-3SL launch vehicle, followed by integrated launch vehicle testing. • Phase III takes place on the Launch Platform. While still at Home Port, the integrated launch vehicle is erected and electrical checkout and radio frequency end-to-end tests are conducted. Following completion of these tests, the vehicle is de-erected and housed in an environmentally-controlled hanger in a horizontal position. The Launch Platform then transports the integrated launch vehicle to the launch site on the equator. Rehearsals are conducted during transit. At the launch site, the integrated launch vehicle is rolled out to the launch pad for erection and fueling. An automated system performs launch operations Figure 8-1 depicts the overall spacecraft processing flow. Figure 8-2 illustrates a typical 55-day processing schedule and indicates down-time on Sundays. Deviations from the 55-day schedule may be accommodated on a case-by-case basis. A typical 28-day start of combined operations flow through launch is illustrated in Figure 8-3.
REV D HPD-10041 8-1 Assembly and Launch Platform Payload Processing Facility Command Ship
Fueling Cell Encapsulation Cell
Zenit and Block DM-SL ILV erection, PLU electrical buildup and checkout checkout, and RF end-to- end test*
SC Installation on adapter* Mechanical and Block DM-SL Transit to launch site electrical checkout* Propellant Loading Installation on ISPS Rehearsals Zenit and Block Installation on strongback DM-SL Mating SC battery charging* assembly* Ordnance Payload unit mating* installation* Horizontal rotation Closeout operations*
PLF encapsulation Integrated test for Transfer to pad and complete launch vehicle* erection RF checkout* Countdown operations Transfer of launch Propellant Transfer of payload unit to ACS vehicle to LP hangar loading* Launch *Tasks that directly involve the SC SHG0681874-015 Figure 8-1. Spacecraft Processing Flow
8-2 HPD-10041 REV D 55 50 45 40 35 30 25 20 15 10 05 01 Description Shift plan Days Previous launch (at 1500 -55 0 or Later) – 55-day centers Launch at 1500 SC on-dock Home Port – (notional) Note: SC stand-alone operations Variable, -29 L-minus days (notional) SC controlled are notional -60 until applied to PLU on-dock Home Port – specific launch PLU stand-alone 8 hr/d, LOE -60 -29 date and time operations 5 d/wk to account for workweek SC required to support – -28 shifting and combined operations agreed-to Combined PLU/SC 8 hr/d, 6 -28 -22 holidays. operations 6 d/wk Integrated Block DM/Zenit – -21 available for PLU mate PLU rollout to ACS. Mate 10 hr/d, 3 -21 -19 to Block DM, integrated 6 d/wk electronic tests, transfer to LP1 -18 -15 ILV erect, KNSG, preps for 12 hr/d, 3 transit (no rerad test) 6 d/wk LP transit to launch site 24 hr/d, -14 -3 12 days ACS transit to launch 7 d/wk 9 days -3 0 Launch readiness 24 hr/d, 72 hr operations, launch 7 d/wk Post-launch securing 24 hr/d, 1 -54 7 d/wk -53 -47 ACS transit to Home Port 24 hr/d, 7 days 7 d/wk ACS returned from -47 previous launch – ILV stand-alone 10 hr/d, 24 -46 -22 operations 6 d/wk Integrated Block DM/Zenit -21 Available for PLU Mate SHG0681874-016.2 Figure 8-2. Nominal 55-Day Processing Schedule
REV D HPD-10041 8-3
Figure 8-3. Nominal Campaign Timeline
8-4 HPD-10041 REV D 8.1 Phase I ⎯ Spacecraft Processing in the Payload Processing Facility
Spacecraft Payload processing begins with delivery of the spacecraft to Sea Launch Delivery Home Port in Long Beach, California. Spacecraft in their respective shipping containers are delivered by truck to Home Port. Delivery of spacecraft via air freight to regional airports and subsequent transportation to Home Port can be facilitated. Spacecraft are received into the common airlock cell in the Payload Processing Facility (Figure 8-4).
265807J6-206 Figure 8-4. Payload Processing Facility at Home Port
Spacecraft Before the spacecraft arrives, the spacecraft adapter is mounted on an Adapter adapter support stand assembly and moved into one of two high-bay spacecraft processing/fueling cells in anticipation of spacecraft fit check operations.
REV D HPD-10041 8-5 Spacecraft Upon arrival, the spacecraft is unloaded into the common air lock and Processing moved into the spacecraft processing/fueling cell. A spacecraft-to- adapter fit check is performed (Figure 8-5). The spacecraft is then transferred to a customer-provided processing and/or fueling stand. The customer performs electrical and mechanical checkout of the spacecraft, stand-alone tests, ordnance installation and propellant loading.
265807J3-097 Figure 8-5. Fit Check on Adapter Support Stand
8-6 HPD-10041 REV D Facilities Astrotech Space Operations, a subcontractor to Sea Launch, provides facility support services while the spacecraft is in the Payload Processing Facility. Section 10 describes this facility and its services. • Access lifts are available to support personnel working on the spacecraft. • An adjacent control room is provided for spacecraft ground support equipment and personnel. • Garment change rooms, office facilities, remote control facilities and warehouse or storage facilities are also available.
Payload Fairing Prior to spacecraft processing, the payload fairing, interface skirt and Delivery payload structure arrive at the Payload Processing Facility, are unloaded in the common air lock cell and are then transferred to the encapsulation cell for standalone processing.
Strongback A strongback assembly is used to rotate the spacecraft from vertical to Assembly horizontal for encapsulation (Figures 8-6 and 8-7), as follows: • Boeing personnel move the interface skirt (IS) and payload structure (PS) into the encapsulation cell and place it on the strongback assembly. • All systems are fully tested before mating and all mechanical and electrical interfaces are verified by established procedures. • When spacecraft stand-alone processing and fueling are complete, the spacecraft is mated to the spacecraft adapter and separation ordnance is installed. The spacecraft is then transported into the encapsulation cell and mated to the interface skirt and payload structure on the strongback assembly (Figure 8-8). • Following mate to the IS/PS, and prior to horizontal rotation, communication checks of the hardline paths are performed and final spacecraft closures are completed.
REV D HPD-10041 8-7 3.5 m Top of payload structure
Payload structure Interface skirt 1.84 m Skirt base 0.65 m Free vertical height*
0 m Floor level
5. (2 54 18 m m .4 7.72 in in) ) (304
* “Crawlspace” height between SBA and PPF floor. 265807J6-288 Figure 8-6. Strongback Assembly Vertical Mating Position
Flexure Interface/spacer Payload mount ring assembly assembly
Drive system and control assembly
Multicomponent load cell
Base assembly 265807J6-289 Figure 8-7. Strongback Assembly Horizontal Raised Position
8-8 HPD-10041 REV D 265807J6-290 Figure 8-8. Strongback Assembly Rotating to Horizontal with Pathfinder Structure
Encapsulation Boeing personnel install the payload fairing on the encapsulated payload transportation mechanical equipment (EPTME). The EPTME (Figure 8- 9) supports the payload fairing in a horizontal position. Alignments are meticulously inspected. The EPTME slides the payload fairing over the spacecraft and mates with the interface skirt.
REV D HPD-10041 8-9 265807J6-337 Figure 8-9. Encapsulated Payload Transportation Mechanical Equipment with Encapsulated Spacecraft
Payload Unit The encapsulated spacecraft, adapter, interface skirt and payload structure, and payload fairing - are together referred to as a payload unit. RF communication checks are conducted on the spacecraft and conditioned airflow can be initiated into the payload fairing cavity.
Scaffolding Scaffolding is available to provide access while the spacecraft is in the horizontal position. Figure 8-10 illustrates this access. Figure 8-11 shows an example in use. Scaffolding is also available on board the Assembly and Command Ship to provide access for maintenance and to support mating of the payload unit to the Block DM-SL.
8-10 HPD-10041 REV D 688-95510-002 scaffold assembly used in encapsulation 34 in cell and on board ACS
Strongback 155 in assembly (simplified)
EPTME (simplified)
88 in
688-95510-001 scaffold assembly used in encapsulation cell only 265807J6-292 Figure 8-10. Scaffolding around Encapsulated Spacecraft
265807J6-293 Figure 8-11. Access to Encapsulated Spacecraft
REV D HPD-10041 8-11 Encapsulated The integrated payload unit is loaded onto the EPTME, for transportation Payload Transport to the Assembly and Command Ship. Vehicle Spacecraft thermal, contamination, vibration, acoustic and electromagnetic radiation environments are rigorously maintained throughout transportation to the Assembly and Command Ship.
Transport from Figure 8-12 shows the payload unit, which includes the encapsulated Payload spacecraft, as it exits the Payload Processing Facility. Processing Facility Figure 8-13 shows the payload unit in transit to the Assembly and to Assembly and Command Ship. Command Ship Figure 8-14 shows the payload unit moving onto the stern ramp of the Assembly and Command Ship.
SHG0681874-043 Figure 8-12. Payload Unit Exiting the Payload Processing Facility
8-12 HPD-10041 REV D SHG0681874-044 Figure 8-13. Payload Unit in Transit to the Assembly and Command Ship
265807J6-296 Figure 8-14. Payload Unit on Stern Ramp of the Assembly and Command Ship
REV D HPD-10041 8-13 Security Security is a high priority throughout spacecraft processing. In the Payload Processing Facility, access is restricted to authorized personnel only. Video cameras monitor controlled access areas at Home Port, on the Assembly and Command Ship and the Launch Platform.
8.2 Phase II ⎯ Mating to the Launch Vehicle and Integrated Testing on the Assembly and Command Ship
Zenit Processing Zenit-3SL launch vehicle processing is performed on the Assembly and Before Spacecraft Command Ship while the spacecraft is still in the Payload Processing Mating Facility. At the pier, complete systems for loading fuels, compressed gases and cryogens are provided. The Zenit first and second stages are fully tested, the Block DM-SL is tested and fueled with kerosene and hypergols, and the Block DM-SL upper stage is mated with the Zenit-3SL rocket. After mating and integrated testing, the launch vehicle is ready to receive the payload unit, which includes the encapsulated spacecraft (Figure 8-15). The fully assembled Integrated Launch Vehicle is shown in Figure 8-16.
265807J6-298 Figure 8-15. Payload Unit Mating to the Zenit-3SL
8-14 HPD-10041 REV D SHG0681874-045 Figure 8-16. Integrated Launch Vehicle in the Assembly and Command Ship
Spacecraft Mating After transfer to the Assembly and Command Ship, the payload unit is to the Zenit-3SL mechanically and electrically mated with the fully assembled Zenit-3SL vehicle. Integration testing is then performed and the complete integrated launch vehicle is fully checked out.
Transfer to the The Assembly and Command Ship is repositioned around the pier, Launch Platform directly in front of the Launch Platform, for transfer of the integrated launch vehicle from the Assembly and Command Ship to the Launch Platform. The launch vehicle is rolled out to the stern ramp of the Assembly and Command Ship assembly area by a ship rail system. Cranes are used to lift the vehicle up from the ramp onto the Launch Platform for placement in the hangar. Clean conditioned air is provided throughout this operation. Figure 8-17 shows the Zenit-3SL vehicle ready for lifting from the Assembly and Command Ship. Figure 8-18 shows the Zenit-3SL vehicle being transferred to the Launch Platform hangar.
REV D HPD-10041 8-15 SHG0681874-046 Figure 8-17. Integrated Launch Vehicle Ready for Lifting
SHG0681874-047 Figure 8-18. Lifting of Integrated Launch Vehicle
8-16 HPD-10041 REV D 8.3 Phase III ⎯Transit and Launch Operations on the Launch Platform
Transit Transit of the Launch Platform (Figure 8-19) from Home Port to the launch site on the equator takes approximately twelve days. The Assembly and Command Ship (Figure 8-20) departs three days after the Launch Platform, and the two vessels rendezvous.
SHG0681874-048 Figure 8-19. Launch Platform Sailing for the Equator
REV D HPD-10041 8-17 SHG0681874-049 Figure 8-20. Assembly and Command Ship Sailing for the Equator
Conditions During transit to the launch site, the Zenit-3SL vehicle is supported on the transporter/erector in the Launch Platform hangar. The ILV is not rolled out of the hangar until the Launch Platform is ballasted down at the launch site. Figure 8-21 shows the transporter/erector rolled out of the Launch Platform hangar. Conditioned air is continuously provided to the payload fairing throughout transit to the launch site. Battery charging is available through the interface skirt connection to spacecraft ground support equipment.
8-18 HPD-10041 REV D SHG0681874-050 Figure 8-21. Zenit-3SL Transporter/Erector
Communication Communication links (both direct and satellite) are provided between the Links spacecraft in the Launch Platform hangar and the customer’s electrical ground support equipment in the spacecraft support room on board the Assembly and Command Ship.
Access The spacecraft is functionally accessible by radio frequency and hard- line connections and is physically accessible through payload fairing doors if required. Access platforms are available in the hanger to facilitate access to most locations on the outside surface of the payload fairing. Specialized platforms to access the spacecraft through the fairing can be made available as non-standard items.
REV D HPD-10041 8-19 Launch Three rehearsals are performed for each launch campaign. The first Rehearsals rehearsal takes place prior to vessel departure. The next two rehearsals, including the mission dress rehearsal, are conducted before the vessels arrive at the launch site. The primary purpose of these launch rehearsals is to train launch management and teams on countdown documentation and protocols. Participation also includes spacecraft tracking and communication systems personnel. Participation of customer personnel on-board the Assembly and Command Ship, and personnel in the spacecraft control center is required. The mission dress rehearsal simulates pre-launch and post-launch operations up through, and including, spacecraft separation and completion of the Block DM-SL contamination and collision avoidance maneuvers. During these rehearsals, launch vehicle operations are simulated and the launch vehicle remains safely in the hangar. Stand-alone vehicle checkouts are completed by each segment to verify operability of the hardware.
Launch Platform At the launch site, the Launch Platform is ballasted to its semi- Stability submersible launch draft position of 22.5 m (73.8 ft). Platform positioning and orientation are maintained using global positioning system inputs to a dynamic positioning system of thrusters on board the vessel. Stability is maintained by an active trim-and-heel system that compensates for mass movements atop the platform.
Weather Like other launch sites, Sea Launch monitors cloud conditions, lightning, Conditions thunderstorms and winds. However, Sea Launch is unique in that wave heights and currents are also monitored. All of these elements are taken under consideration in determining launch availability. For historical weather data, reference Appendix C.
Year-round The combination of typically benign weather, inherent platform stability Launch Capability and active leveling systems makes it possible to launch the rocket year round in most expected weather conditions (Figure 8-22).
8-20 HPD-10041 REV D SHG0681874-051 Figure 8-22. Launch Operations at the Launch Site
Vessel to Vessel Sea Launch has the ability to transfer personnel between the ACS and LP Transfer by either helicopter or link bridge operations depending upon the Operations sequence within the mission timeline. Once the Launch Platform has ballasted after arrival at the launch site, the Assembly and Command Ship and Launch Platform are able to perform link bridge operations (Figure 8-23) and helicopter transfer operations. Following ILV erection at the launch site, all personnel transfer operations are conducted via helicopter.
REV D HPD-10041 8-21 265807J6-203 Figure 8-23. Assembly and Command Ship and Launch Platform Connected by Link Bridge
Launch Readiness After the vessels have arrived at the launch site, a launch readiness Review review is conducted and approval is given to the launch team to commence final launch operations.
Launch Vehicle The hangar doors are opened and the automatic sequence that moves the Erection Zenit-3SL to the launch pad is initiated. As the integrated launch vehicle moves to the pad, electrical, pneumatic, hydraulic and propellant lines are automatically connected. Before erection, air is switched from the hangar conditioned-air supply to the pad conditioned-air supply system. The integrated launch vehicle is erected automatically and rotated to a vertical position.
8-22 HPD-10041 REV D Evacuation of Personnel are evacuated from the Launch Platform via helicopter in a Personnel timely and orderly manner in accordance with the launch countdown documentation. All personnel, except for critical launch vehicle technicians, are transferred to the Assembly and Command Ship, positioned 6.5 km away. After the launch vehicle technicians complete final pre-launch checkouts, the helicopter transfers them to the Assembly and Command Ship, prior to rocket fueling. From this point on, the Launch Platform is operated under remote control from the Assembly and Command Ship. Radio frequency links provide communications between the ground support equipment rooms and the spacecraft.
Final Critical Launch operations progress with readiness polls at key milestones. A Operations poll of launch management, the spacecraft manufacturer and the customer is conducted to give the “go” for launch vehicle propellant loading. The Zenit-3SL is loaded with propellants and compressed gases starting at L-2:20. Prior to L-30 minutes, launch holds can be accommodated. Following this point, off-nominal conditions would result in a launch abort. The launch commit poll is conducted at L-36 minutes, prior to disconnecting the launch vehicle umbilical at L-17 minutes. The transporter/erector transitions to the hangar starting at L-12 minutes.
Holds and Aborts The Sea Launch system has the ability to hold the countdown up to L-30 minutes to address off-nominal conditions. Pre-determined maximum hold durations are available prior to key countdown milestones, including: L-5 hours start of auto-processing, L-2:20 ILV tanking, and the L-36 minute launch commit poll. Launch aborts called prior to LOX GSE chill down starting at L-4:20 minutes would result in a one- to two-day stand down. Following the start of ILV propellant loading, four days are required to reinitialize preparations for another launch attempt. Re-initialization entails draining oxygen and kerosene from the Zenit- 3SL, charging the spacecraft batteries if required, and returning the integrated launch vehicle to the hangar. Continual conditioned airflow to the spacecraft is maintained through any abort or recycle operations.
REV D HPD-10041 8-23 Zenit-3SL in The Zenit-3SL is held in place on the launch table by hold-down clamps Vertical at the base of the first stage. These clamps are not released until Configuration computers confirm that the Stage 1 operating thrust has been reached. In order to minimize exhaust effects on the Launch Platform and acoustic effects on the spacecraft, a water deluge system beneath the launch table is used. (Figure 8-24)
SHG0681874-052 Figure 8-24. Liftoff from the Launch Platform
8-24 HPD-10041 REV D
Launch Control The launch control center is located on the shelter deck of the Assembly Center and Command Ship (Figure 8-25). During all phases of launch vehicle processing and flight, the launch control center management has the ultimate responsibility and authority for all decisions and commands affecting the integrated launch vehicle. Internal communications on both the Assembly and Command Ship and Launch Platform include closed- circuit television, telephone, intercom, public address, data communications and computing systems, a line-of-sight system, a countdown clock system and a mission management display system. Three console positions are available for customer spacecraft personnel in the launch control center.
SHG0681874-053 Figure 8-25. Launch Control Center on the Assembly and Command Ship
REV D HPD-10041 8-25 Mission Director If any unplanned event occurs during a mission that affects nominal Control launch operations, the Mission Director will coordinate all required actions in accordance with instructions contained in the master procedures. These actions may include: • Overseeing delay-of-launch operations • Overseeing safety measures • Identification of the personnel required for assessment and resolution of the event • Coordination of the personnel team to assess the event and provide recommendations • Authorization of launch recycle or recovery activities • Preparation of required notifications and reports
8.4 Flight Operations
Range Tracking Range tracking and telemetry return is provided by Sea Launch using and Telemetry assets on board the Assembly and Command Ship through NASA’s Tracking and Data Relay Satellite System. Ground monitoring installations include the launch control center on board the Assembly and Command Ship and the Russian ground tracking station in Korolev, outside Moscow. Line-of-sight telemetry coverage is continued through payload fairing jettison. The Tracking and Data Relay Satellite System telemetry is routed through the central communication node in Brewster, Washington.
Range Sea Launch provides all necessary range coordination functions Coordination independent of the eastern or western range sites in the United States. These include: • FAA (air traffic) • National Geospatial-Intelligence Agency (notice to mariners) • NASA (orbital collision avoidance) There is little risk associated with range availability because there are typically no other launches competing for Sea Launch assets at the same time.
8-26 HPD-10041 REV D Weather The Assembly and Command Ship includes: • Self-contained weather station with C-band Doppler radar • Upper atmospheric balloon release capability • Surface wind instruments • Wave radar • Ambient condition sensors • Satellite imagery A full-time meteorologist is stationed on the Assembly and Command Ship while at sea. Historical weather data at the Launch Site is listed in Appendix C.
Other • A helicopter is provided to support mission operations. Both the Assembly and Command Ship and the Launch Platform are equipped with landing pads. A backup helicopter may also be utilized for contingencies and is available onboard the Assembly and Command Ship. • A comprehensive photo-optical recording system is used to document the launch. • External communications are handled through a satellite communication system linking both the Assembly and Command Ship and the Launch Platform to each other and to a Sea Launch ground station in Brewster, Washington.
Zenit-3SL Flight Zenit first- and second-stage flight operations are completely automatic. Operations A mission event timeline is included in Table 3-1, Section 3. All Stage 1 and most Stage 2 events occur within the view of the Assembly and Command Ship. The spent stages fall into the Pacific Ocean, far short of the coast of South America and the major coastal shipping lanes. Any deviation of flight trajectory from preprogrammed limits causes onboard systems to automatically terminate propulsion and end the mission. This approach to flight safety has been used successfully for decades and obviates the need for the traditional range safety officer. At second-stage separation from the Block DM-SL, four solid-propellant rocket motors at the base of Stage 2 fire to back the stage away from the Block DM-SL. The Block DM-SL lower and middle adapters are jettisoned during this period; the lower adapter is jettisoned with Stage 2, and the middle adapter is jettisoned just before Block DM-SL ignition.
REV D HPD-10041 8-27 Block-DM Flight Before launch, the Block DM-SL onboard systems are turned on and Operations initialized, its oxidizer tank level is adjusted, and power is transferred from the Launch Platform umbilical to the Block DM-SL internal power supply. During the Stage 1 and 2 flight, the Block DM-SL remains inactive, except for preparations for autonomous flight. Following Stage 2 separation, the Block DM-SL is inserted into the target orbit with a single main engine burn. For two-burn missions, the Block DM-SL performs a settling burn using the attitude control/ullage propulsion system before the second burn of the main engine. Burn program options include, but are not limited to, one- or two-impulse insertion geosynchronous transfer orbits and multiple burns (up to a maximum of five) to medium Earth orbit or planetary escape. Following spacecraft insertion into the target orbit, the Block DM-SL separates from the spacecraft and performs a contamination and collision avoidance maneuver. Disposal options include transfer of the Block DM-SL to a higher or lower disposal orbit that mitigates orbital debris generation risk.
Postlaunch Sea Launch provides near-real-time confirmation of key ascent events Reports (e.g., payload fairing jettison, engine burns and spacecraft separation). An injection report is delivered within fifty minutes of spacecraft separation. The injection report documents the state vector of the Block DM-SL just before spacecraft separation and is provided to customer personnel in the launch control center. A postflight report is delivered sixty days after launch. The postflight report documents performance of Sea Launch systems and interfaces as related to documented requirements. Data includes operations and launch environments (ground and flight), launch time, separation state vector, injection orbit and accuracy, separation attitude and rates, contamination and collision avoidance maneuver performance and flight event sequences.
8-28 HPD-10041 REV D 9.0 HOME PORT FACILITIES AND SEA LAUNCH VESSELS
Overview The Sea Launch Home Port complex provides the facilities, equipment, supplies, personnel and procedures necessary to receive, transport, process, test and integrate spacecraft and associated support equipment with Sea Launch. Figure 9-1 shows the general layout of the complex in Long Beach, California. Home Port also serves as the base of operations for both of the Sea Launch vessels. Figure 9-2 shows the Assembly and Command Ship and Launch Platform at sea. The personnel providing day-to-day support and service during prelaunch processing and launch operations for Sea Launch and its customers are located at Home Port. In addition to the overview provided here, the Sea Launch Facilities Handbook for customers provides a complete description of all buildings, support equipment and services available to Sea Launch customers while at Home Port or on the vessels. Characteristics of Home Port facilities and Sea Launch vessels are covered in Sections 9.1 through 9.3, including: • Home Port facilities • The Assembly and Command Ship (Sea Launch Commander) • The Launch Platform (Odyssey)
REV D HPD-10041 9-1 Integrated LV Transfer to LP Pier 16
Assembly and command ship Launch platform Payload Payload Unit Mate processing to Rocket facility Gate 4 Ordnance Gate 5 Gate 3 Gate 7 storage N Gate 13 Unassigned Gate 8 Gate 10 Facility Warehouse 10 6 2 fence and 1 Facility 7 8 9 customer SC Arrival offices 4 Gate 6 fence at Home Gate 2 11 12 13 3 Port 5 CIS office space
Access Gate 12 Warehouse Gate 1 Gate 11 road buildings Home Port Gate 9 management 265807J6-304/R1 Figure 9-1. Sea Launch Home Port Complex
9-2 HPD-10041 REV D 265807J6-305 Figure 9-2. Sea Launch Vessels Home Port Home Port is a sixteen-acre site located in Long Beach, California. Location Situated south of Los Angeles, the complex is within the boundaries of the Port of Long Beach at the eastern end of the former Navy mole that partially forms the Long Beach Harbor. Figure 9-3 depicts the location of Home Port. Access to the site is by the San Diego freeway (I-405) and then either I-110 or I-710 to Long Beach. Long Beach Airport (21 km or 13 mi), Los Angeles International Airport (40 km or 25 mi), and John Wayne Airport (38 km or 24 mi) are all in close proximity.
REV D HPD-10041 9-3 SHG0681874-054 Figure 9-3. Home Port Location and Vicinity
9-4 HPD-10041 REV D 9.1 Home Port Facilities
Payload The Payload Processing Facility (Figure 9-4), also known as Building 1, Processing is located on the east side of the Home Port complex. All spacecraft Facility processing, propellant transfer operations, pressurization, ordnance preparation and payload fairing encapsulation operations are accomplished in Building 1. Through the use of duplicate and independent processing cells, the Payload Processing Facility can concurrently support the discrete processing campaigns of two customers. The Payload Processing Facility is separated from the rest of the complex by an interior fence with controlled access during hazardous spacecraft operations. The Payload Processing Facility has an overall area of 2872 m2 (31,000 ft2). Major features include: • Spacecraft processing and fueling cells • Fuel storage rooms • Oxidizer storage rooms • Control rooms • Garment change rooms • An air lock • An encapsulation cell • A lobby and break room
REV D HPD-10041 9-5 Break room Encapsulation cell 15 m x 38 m (50 ft x 125 ft) Lobby
Change room 4 Air lock 15 m x 23 m (50 ft x 75 ft)
Processing/ Processing/ Fuel 1 fueling cell 1 fueling cell 2 15 m x 20 m 15 m x 20 m Oxidizer1 (50 ft x 65 ft) (50 ft x 65 ft) Fuel 2
Fire riser Oxidizer 2
Control Control Generator room room 1 room 2
Change room 1 Change room 2 265807J6-307 Figure 9-4. Payload Processing Facility—Interior and Exterior Views
9-6 HPD-10041 REV D Processing and The processing and fueling cells are designed to facilitate on-site Fueling Cells electrical testing, spacecraft fueling and final assembly before encapsulation of the spacecraft into the payload fairing. Spacecraft providers are assigned one of the two high-bay processing and fueling cells (Figure 9-4). Both cells have large-capacity overhead cranes, are certified cleanrooms, and have fueling islands that offer fuel spill and contamination control.
Fuel Storage A fuel storage room and an oxidizer storage room is dedicated to each of Room and the spacecraft processing and fueling cells. Both rooms fully support fuel Oxidizer or oxidizer storage and all fueling operations. Complete personnel Storage Room protection is provided during hazardous operations and these rooms offer containment systems in the event of accidental leakage.
Control Rooms Independent control rooms, from which spacecraft operations are conducted, are located adjacent to each of the processing and fueling cells. Penetrations between the control room and the processing and fueling cells allow direct communication between the spacecraft and its controlling support equipment. A window allows observation of operations in each processing and fueling cell.
Garment Change The garment change rooms are located adjacent to the processing and Rooms fueling cells. These rooms provide the supplies and environment appropriate for personnel to don cleanroom garments or other protective suits before entering a processing and fueling cell.
Air Lock The common air lock provides an environmentally controlled cleanroom that allows access to either the encapsulation cell or one of the two spacecraft processing and fueling cells. Customer equipment is delivered to this cell and distributed throughout the rest of the facility. An overhead crane, forklifts and other support equipment are available to assist the customer.
REV D HPD-10041 9-7 Encapsulation Cell The encapsulation cell is an environmentally controlled cleanroom used for the preparation of the payload fairing, the interface skirt and the payload structure. This cell contains all the equipment necessary to mate the spacecraft to the adapter and encapsulate that assembly within the payload fairing.
Lobby The lobby of the Payload Processing Facility is available for viewing the activities within the encapsulation cell. A break room is connected to the lobby.
Ordnance Facility An ordnance storage building (Building 2) is located near the Payload Processing Facility at Home Port. This building is used to store the packaged explosive items for use by the launch vehicle and its payload.
Customer Building 10 (Figure 9-1) is available to the customer for storing supplies, Warehouse/ shipping containers and other miscellaneous equipment. Storage areas Storage can be secured with access controlled by the customer. With prior Facilities coordination, additional customer storage may be arranged.
The Pier The pier (Figure 9-1) provides facilities for mooring, servicing and supplying the Sea Launch vessels. The pier has all the provisions necessary for providing communications, water and sewer services to the vessels while in port. It also has equipment for loading fuel, personnel and equipment on board the vessels.
Building 4 Building 4 provides office and conference space for customer personnel. Also within Building 4, a pair of remote control rooms is provided for customers who wish to remotely control or monitor their operations in the Payload Processing Facility.
Building 3 Building 3 (Figure 9-5) provides facilities for the resident Home Port administrative and professional staff. This building contains offices, a training area, conference rooms and a break area.
9-8 HPD-10041 REV D 265807J6-308 Figure 9-5. Building 3, Sea Launch Office Facility
Transportation When customer equipment destined for Home Port arrives by aircraft, Services Sea Launch provides logistical assistance with customs, transportation and delivery from local airports. Forklifts, trucks, dollies and air-bearing pallets are available to the customer to assist with equipment movement within the Home Port complex. Sea Launch is responsible for transportation of the encapsulated spacecraft between Home Port facilities.
REV D HPD-10041 9-9 9.2 Assembly and Command Ship
Functions and The Assembly and Command Ship (Figure 9-6) performs four functions Characteristics for Sea Launch operations: • Serves as the facility for assembly, processing and checkout of the launch vehicle • Houses the launch control center, which monitors and controls all operations at the launch site • Acts as the base for tracking the initial ascent of the launch vehicle • Provides accommodations for the marine and launch crews during transit to and from the launch site and for all personnel during the fi- nal stage of launch operations The Assembly and Command Ship is designed and constructed specifically to suit the unique requirements of Sea Launch. The ship’s overall dimensions are nearly 200 m (660 ft) in length and 32 m (110 ft) in beam. Its overall displacement is approximately 27,612 tonnes (30,400 tons). Major features of the Assembly and Command Ship include: • A rocket assembly compartment • The launch control center • A launch control center viewing room • Helicopter hangar and pad • Spacecraft contractor and customer work areas • Spacecraft contractor and customer accommodations
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3rd bridge deck 4th bridge deck 2nd bridge deck 1st bridge deck
Shelter deck Helicopter hangar
265807J6-309 Figure 9-6. Assembly and Command Ship
Rocket Assembly The rocket assembly compartment (Figure 9-7), located on the main deck Compartment of the Assembly and Command Ship, hosts the final assembly and processing of the launch vehicle. This activity is conducted while the vessels are at Home Port and typically in parallel with spacecraft processing. The bow of the main deck is dedicated to processing and fueling the Block DM-SL. After the completion of spacecraft processing and encapsulation in the Payload Processing Facility, the encapsulated payload is transferred into the rocket assembly compartment, where it is integrated with the 2-stage Zenit and Block DM-SL.
REV D HPD-10041 9-11 Rocket assembly Encapsulated Block DM-SL Stern ramp compartment spacecraft processing
Integrated launch vehicle Block DM-SL fueling SHG0681874-055 Figure 9-7. Launch Vehicle Processing—Main Deck
Spacecraft Accommodations are available for spacecraft manufacturers to check Checkout spacecraft health and verify interfaces before and after integration with Accommodations the launch vehicle. After the payload is integrated with the launch vehicle and all checkouts are complete, the fully integrated launch vehicle is transferred to the Launch Platform (Figure 9-8). Environmental conditioning and monitoring of the encapsulated spacecraft is ongoing, from encapsulation through liftoff.
265807J6-311 Figure 9-8. Launch Vehicle Transfer from Assembly and Command Ship to Launch Platform at Home Port Launch Control The launch control center is the primary center from which day-to-day Center and launch operations are conducted. The launch control center is located on the Assembly and Command Ship shelter deck, as shown in Figure 9- 9. Layout of the console and display features of the room are shown in Figure 9-10.
9-12 HPD-10041 REV D Launch control center
265807J6-312 Figure 9-9. Launch Control Center—Shelter Deck
265807J6-313 Figure 9-10. Launch Control Center Mission The launch control center contains operator consoles and seating, and Management and provides connections with all shipboard communication systems. Display System The mission management and display system receives and processes information from the shipboard weather equipment, marine systems, closed-circuit television, Energia shipboard launch control system, Energia shipboard flight control system, the global positioning system and other sources. The system displays information on large projection screens and television monitors at each operator console and on television monitors at other locations on the Assembly and Command Ship. The computer system provides standard applications software for the operations and interconnection with other shipboard and on-shore personnel (through the external communications system) for file transfer, e-mail and Internet capabilities.
REV D HPD-10041 9-13 Capabilities Launch control center capabilities available to the customer include: • Three console stations (Figure 9-11) • Headset and telephones to connect each console station to the mis- sion control communications network • Countdown clocks, mission control information and closed-circuit television camera views on the front wall Data management network connection is provided by a local area network that allows Internet access through a local firewall server.
265807J6-205 Figure 9-11. View of Launch Control Center Consoles and Front Wall (Boeing/Customer Side of Room)
Viewing Room Adjoining the launch control center is a viewing room for observing launch activities. This room contains seating for as many as four customers or prominent guests, allows observation of the launch control center and provides access to the communications networks and telephone system.
9-14 HPD-10041 REV D Customer Work The spacecraft contractor and customer work areas are located on the Areas first bridge deck directly above the launch control center (Figure 9-12). The area has ground support equipment rooms, office areas, conference rooms, individual offices and storage rooms. The ground support equipment rooms provide space for the equipment necessary to control and monitor the spacecraft during mission operations. The ship’s overhead cranes load the ground support equipment onto the deck area just behind each room. The ground support equipment rooms contain: • Patch panels for hard-line links to the launch control center • Countdown and count-up clocks • Public address, intercom and telephone services • Overhead monitor displays connected to closed-circuit television networks • 120/208-V, 60-Hz and 220/380-V, 50-Hz power • Uninterruptible power supplies for 60-Hz and 50-Hz power
SeaCustomer Launch storagestorage Office Office Office Office Office P P P P P, PC
Sea Launch weather SeaCustomer Launch office office area area offices CustomerSea Launch P, PC ground F, I, P, VCR support equipment VCR,CustomerSea Launch VTC, I, P room conference CCTV, P, F, I, M room Sea Launch office area CCTV, P, F, I, M VCR,Customer VTC, I, P Customer conference PC, P, F, I, VCR ground room support equipment room
Customer office area F, I, P,VCR Sea Launch EGSE room
Customer OfficeP OfficeP OfficeP P storage SHG0681874-056 Figure 9-12. Customer Work Area on First Bridge Deck
REV D HPD-10041 9-15 Customer Office The two office areas for customers consist of an open bay and three Areas individual offices. A conference room and a storage room dedicated to customer use are also provided adjacent to the office area. Additional conference space is provided directly behind the launch control center. Each conference room has seating, overhead projectors and teleconferencing equipment.
Communication Telephones, fax machines, teleconferencing and data links are provided Satellite Services from the Assembly and Command Ship and Launch Platform to customer home locations through commercial communication satellite services (Figures 9-13 and 9-14). Radio frequency links between the Assembly and Command Ship and the Launch Platform allow communications with the spacecraft in the hangar and on the launch pad, and between spacecraft ground support equipment rooms on the Assembly and Command Ship and the Launch Platform.
SHG0681874-057 Figure 9-13. Assembly and Command Ship Satcom Radome
9-16 HPD-10041 REV D SHG0681874-058 Figure 9-14. Launch Platform Satcom Radome Cabin Customer and spacecraft contractor accommodations are located on the Accommodations first bridge deck and the second bridge deck (Figure 9-15). The Assembly and Command Ship has accommodations for a total of eleven customer and spacecraft contractor personnel (Figure 9-16). Crew members assigned to the Assembly and Command Ship are allocated either individual or shared cabins. Sea Launch assigns as many individual cabins as possible. Some cabins are equipped with a separate fold-down bunk to serve as a temporary berth for Launch Platform crew members who transfer to the Assembly and Command Ship during final launch preparations.
REV D HPD-10041 9-17 Second bridge deck Legend: Office space Crew quarters GSE rooms Conference rooms
First bridge deck
Crew quarters
265807J6-316 Figure 9-15. Spacecraft Customer Accommodations – First and Second Bridge Decks
SHG0681874-059 Figure 9-16. Assembly and Command Ship Customer Cabin Accommodations Dining, The Assembly and Command Ship provides various off-hour Recreation and entertainment opportunities during the transit to and from the launch site. Medical The Sea Launch catering staff serves regular meals in the main dining hall (Figure 9-17). Mealtimes are rearranged as necessary to support continuous operations at the launch site during final preparations. Other features of the Assembly and Command Ship include a lounge, coffee lounge, theater, recreation room, exercise room and sauna (Figure 9-18). The ship also houses a hospital, and qualified medical personnel are members of the marine crew staff.
9-18 HPD-10041 REV D 265807J6-318 Figure 9-17. Main Dining Hall
REV D HPD-10041 9-19 • Dining Hall
• Coffee bar
• Lounge
• Theater
• Forward lounge
• Smoking lounge
• Fitness room
• Hospital SHG0681874-060.1 Figure 9-18. Assembly and Command Ship Facilities
9-20 HPD-10041 REV D 9.3 Launch Platform
Characteristics The Launch Platform has all the necessary systems for erecting the integrated launch vehicle into launch position, fueling the launch vehicle and conducting launch operations (Figure 9-19). The Launch Platform transports the integrated launch vehicle, in a horizontal position, to the launch site in the enclosed hangar on the launch deck. It also provides accommodations for the marine and launch crews staffing the Launch Platform during transit to and from the launch site. After the launch vehicle has been erected and all launch system checks are complete, all crew members are evacuated from the Launch Platform to the Assembly and Command Ship using a link bridge between the vessels or by helicopter. Launch Platform station keeping and launch operations are conducted remotely from the Assembly and Command Ship through redundant radio frequency links between the vessels.
REV D HPD-10041 9-21 Launch vehicle in launch position
Crew accommodations Launch deck
GSE rooms Rocket hangar
Launch pad
Launch vehicle in transit position
Water line—Launch operations
Water line— transit SHG0681874-061 Figure 9-19. Launch Platform During Final Countdown
9-22 HPD-10041 REV D Specifications The Launch Platform is an extremely stable platform from which to conduct launch operations. The Launch Platform rides catamaran style on a pair of large pontoons. It is self-propelled by a four-screw propulsion system (two in each hull). Each screw is powered by direct- current double armature-type 3,000 hp motors. Once at the launch location, the pontoons are submerged to a depth of 21.5 m (70.5 ft) to achieve a very stable launch position, level to within approximately one degree. Ballast tanks are located in the pontoons and in the lower part of the columns. Six ballast pumps, three in each pontoon, serve these tanks. The Launch Platform has an overall length of approximately 133 m (436 ft) at the pontoons, and the launch deck is 78 by 66.8 m (256 by 219 ft). Its overall transit displacement is approximately 24,400 tonnes (26,900 tons).
Accommodations The spacecraft contractor work area (Figure 9-20) is located on the starboard side of the Launch Platform on the 38500 deck. This area has rooms for customer ground support equipment, offices and storage. Crew accommodations are located on decks 38500, 42500 and 45500, on either side of the launch vehicle hangar. The Launch Platform has accommodations for three spacecraft contractor personnel. Crew members assigned to the Launch Platform are allocated either individual or shared cabins.
Mag tape Data 214 213a 213b Reserved storage process
211 210 212c Time system Launch Marine equipment measurement storage Meeting room complex LAN, PBX, VTC Storage SCC ground sup- and spares Office 209b port equipment CCTV, LAN, 218c LAN, PBX, PBX VTC Central Storage 221c telecomm. and spares 216a systems Control panels room 218b Ground support equipment 221d 204 Storage 209a LAN, PBX, and spares VTC SCC office CCTV, LAN, 218a 216b PBX
SHG0681874-062.1 Figure 9-20. Spacecraft Contractor Work Areas on Launch Platform, 38500 Deck Level
REV D HPD-10041 9-23 Communications Telephones, fax machines, teleconferencing and data links are provided and Access on the Launch Platform. Radio frequency links allow communication between the Assembly and Command Ship and the spacecraft, in the hangar and on the launch pad. Radio frequency links are also provided between spacecraft contractor ground support equipment rooms on the Assembly and Command Ship and the Launch Platform. In addition to radio frequency and hard-line connections to the spacecraft, physical access through doors in the payload fairing is available during the transit to the launch site as a non-standard option. Physical access to the spacecraft is not available once the launch vehicle is erected on the pad.
Dining, The Launch Platform has a recreation and video room for off-hour Recreation and relaxation during transit to and from the launch site. The Sea Launch Medical catering staff serves regular meals in the cafeteria. In addition, there is a small self-serve dining area with food and beverage service. The Launch Platform also houses a hospital with an attending doctor as part of the marine crew staff.
9-24 HPD-10041 REV D 10.0 OPERATIONS SUPPORT SYSTEMS
10.1 Mission Support Systems
Overview Sea Launch provides a number of support systems that are available for the customer’s use in support of the launch, including: • Communications • Weather systems • Helicopter support The systems are described in further detail in this section.
Communications Internal communication systems are distributed between the Assembly and Command Ship and the Launch Platform. This system includes closed-circuit televisions, telephones, intercom, public address, video teleconferencing and vessel-to-vessel communications (also known as the line-of-sight system). This system links with the external communication system providing a worldwide network that interconnects the various segments of the Sea Launch system. The external communication system includes SATCOM with three ground stations. The ground stations are located in Brewster, Washington, and on the Assembly and Command Ship and the Launch Platform, and provide the primary distribution gateways to other communication nodes. Customers can connect to the Sea Launch communication network through the convenient Brewster site. The SATCOM system ties in with the Assembly and Command Ship and Launch Platform Private Branch Exchange (PBX) systems to provide telephone connectivity (see Figures 9-13 and 9-14). Additionally, critical telephone capability can be ensured by INMARSAT.
Weather Systems The Assembly and Command Ship has a self-contained weather station with the central office located on the first bridge deck (Figure 10-1). Components of the weather system include a motion-stabilized C-band Doppler radar, upper-atmospheric balloon release station, surface wind instruments, wave radar and ambient condition sensors. The onboard meteorologist has access to satellite imagery and information from an onsite buoy (Figure 10-2). Forecasts are provided routinely throughout the processing flow to monitor any potential operational weather constraints.
REV D HPD-10041 10-1 SHG0681874-063 Figure 10-1. Weather Office
10-2 HPD-10041 REV D 265807J6-324 Figure 10-2. On-Site Meteorological Buoy
Helicopter Sea Launch operates two helicopters (Figure 10-3), the primary Sea Operations Launch One Bell 230, and a backup helicopter, to support mission operations such as prelaunch and postlaunch platform inspections, crew transit, downrange clearing of the launch corridor, perimeter security patrol and additional functions. Each aircraft can carry as many as seven passengers. The Assembly and Command Ship and the Launch Platform are equipped with helicopter landing pads rated at 12,000 lb. The Assembly and Command Ship is also equipped with a helicopter hangar and a 4,000-gal fuel storage tank. The helicopter has an operational range of 300 nautical miles and can be used for medical emergencies, if required.
REV D HPD-10041 10-3 265807J6-325 Figure 10-3. Sea Launch Bell 230 Helicopter
Photographic The Sea Launch photo-optical measurement system has several Services components on the Assembly and Command Ship and the Launch Platform. On the Launch Platform, a high-speed film system and several sound-initiated still cameras are installed to document liftoff of the launch vehicle. On the Assembly and Command Ship, an optical tracking system is provided (Figure 10-4). This tracking mount is equipped with a combination of high-speed film and video cameras outfitted with telescopic and zoom lenses to document the ascent phase through fairing separation. Digital cameras also are available on both vessels and at Home Port to support flight hardware closeouts, configuration documentation and other photographic requirements.
10-4 HPD-10041 REV D SHG0681874-064 Figure 10-4. Optical Tracker
External Agency Sea Launch handles all coordination with external agencies required for Coordination launch notifications and to assure compliance with the launch license. These agencies issue the necessary advisories: • FAA (air traffic) • National Imaging and Mapping Agency (notice to mariners) • NASA (orbital collision avoidance) Sea Launch issues additional notices to mariners and fisheries that may have vessels in the area and any local air traffic coordinators who may be affected.
REV D HPD-10041 10-5 10.2 Telemetry Assets Overview Sea Launch uses line-of-sight telemetry systems for the initial flight phase and a down range tracking station at Nuku Hiva (Figure 10-7) for interim coverage, as well as the NASA Tracking and Data Relay Satellite System (TDRSS) for later phases. The line-of-sight system, which in- cludes the Proton antenna and the S-band system, is located on the As- sembly and Command Ship (Figures 10-5 and 10-6). Other telemetry as- sets include Russian ground tracking stations, as required, and the Energia Moscow Control Center. The following sections apply to launch vehicle and payload unit telemetry reception and routing.
265807J6- 327 Figure 10-5. Proton Tracking Antenna
10-6 HPD-10041 REV D 265807J6-328 Figure 10-6. S-Band System Radome
Flight Vehicle Line Zenit and Block DM-SL data is received by the Proton antenna complex of Sight Coverage on board the Assembly and Command Ship from shortly before liftoff until approximately seven or eight minutes into the flight. Simultaneously, Payload Unit data is being received in the Boeing telemetry room, from the S-band system (Figure 10-6). The S-band system is capable of relaying very large amounts of data. Its duration covers liftoff to payload fairing separation, which occurs approximately three minutes into the flight. Following fairing separation, the Block DM-SL and payload unit telemetry system begin to deliver telemetry data through the TDRSS.
REV D HPD-10041 10-7 Nuku Hiva Track- Zenit data is received by a tracking station (Figure 10-7) on Nuku Hiva ing Station island (of the Marquesas) from approximately seven or eight minutes into the flight until the Block DM-SL separates from the Zenit stage 2. The full rate telemetry data is recorded for later playback, and is sent to Yuzhnoye SDO facilities as well as the ACS. Simultaneously, a low rate subset of the telemetry data is transmitted via Inmarsat through the Sea Launch network to the ACS. On board the ACS, the data is routed as with the line-of-sight system.
SHG0681874-065 Figure 10-7. Nuku Hiva Ground Station
10-8 HPD-10041 REV D Tracking and Data When the payload fairing separates, the payload unit transmitter shifts Relay Satellite from sending high-rate payload accommodation data by line-of-sight to System sending combined data from payload unit/Block DM-SL by TDRSS. The combined data is routed from the NASA White Sands Ground Station to the Sea Launch Brewster Ground Station, where it is separated into Block DM-SL and payload unit data and then sent to the Assembly and Command Ship. The data is received on board the ship through the satellite communications terminal and routed to separate rooms for upper-stage data and payload unit data. Simultaneously, Brewster routes Block DM-SL data to the Energia Moscow Control Center. TDRSS coverage continues until after playback of the recorded Block DM-SL data. Data is also recorded at White Sands and Brewster for later playback.
Russian Tracking For missions where there may not be adequate TDRSS coverage, Facilities and Russian tracking facilities may be brought on line. For low-inclination Moscow Control trajectories, the Block DM-SL will not be within line-of-sight of the Center Russian tracking facilities until it obtains sufficient altitude. When the Block DM-SL is within line-of-sight of a Russian tracking facility, telemetry is transmitted to the tracking facility and passed on to the Moscow Control Center and to the launch control center on the Assembly and Command Ship by way of communication satellites. Telemetry recorded on board the Block DM-SL is also relayed to the ground station. A system called Kvant is also used during the countdown to load the flight data on board the Block DM-SL. When in view of ground stations, Kvant allows for an independent orbit determination by comparing uplink and downlink radio frequency carrier phase shifts. This orbit determination is separate from the onboard guidance telemetry. During all phases of launch processing, the launch control center on the Assembly and Command Ship has ultimate responsibility and authority for all decisions and commands affecting the launch vehicle, Block DM- SL, and the spacecraft.
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11.0 SAFETY
Overview This section describes the integration and licensing process for customer spacecraft safety, and details the requirements for Sea Launch safety approval. These include: • Safety integration approach • Safety assessment phases • Launch licensing • Safety of operations All launches conducted by Sea Launch require safety evaluation and approval by Sea Launch Safety and licensing by the U.S. Department of Transportation, Office of the Associate Administrator for Commercial Space Transportation. To obtain safety approval, the customer must identify, develop and verify, through the required Sea Launch safety documentation, that its equipment and operations comply with established Sea Launch safety requirements. It is essential that the customer provide required data in a timely manner to assess all potential hazards introduced by spacecraft processing, launch activities and effectiveness of control measures used by both the customer and Sea Launch.
Authority Sea Launch is responsible for the safety of operations and compliance with federal, state and local laws, codes and regulations. Sea Launch Safety has prime responsibility and authority for management, implementation and maintenance of the Sea Launch system safety program, as documented in the Sea Launch System Safety Plan, Boeing document D688-10014-1.
11.1 Safety Integration Approach
Overview Sea Launch uses a three-phase approach to define required data for the purpose of conducting reviews for launch customers’ spacecraft processing safety. Follow-on missions using spacecraft and operations similar to those Sea Launch has previously performed incorporate an abbreviated version of this approach, in which changes to the previously approved mission are identified and evaluated for compliance. This process and the required safety data are identified in Chapter 3 of the Sea Launch Safety Regulations Manual, Boeing document D688-10024-1.
REV D HPD-10041 11-1
Customer The customer must demonstrate to Sea Launch that the spacecraft Responsibilities payload, support equipment and operations comply with established Sea Launch safety regulations. For Sea Launch approval, the customer submits required safety documentation, in accordance with the schedule set forth in the Statement of Work, identifying hazards, developing risk assessment and verifying the hazard controls in accordance with Sea Launch risk acceptance criteria.
Sea Launch Sea Launch incorporates customer safety data into a spacecraft interface Responsibilities hazard analysis to assess impacts to the Sea Launch system. A safety review schedule and data submittal plan is established during initial customer interface discussion and documented in the customer/Sea Launch mission integration schedule and customer/Sea Launch safety plan.
Safety and Sea Launch conducts safety review phases concurrent with mission Readiness Reviews preliminary design review and critical design review phases, where applicable. A ground operations readiness review is conducted before beginning spacecraft and payload unit operations at Home Port. This completes the safety review and approval process. A mission readiness review is conducted before the Sea Launch vessels leave Home Port for the launch site. A launch readiness review is conducted before launch countdown operations begin. These reviews ensure that preparations for launch have been completed and all conditions of the launch license have been met. Any open safety issues following the mission readiness review and launch readiness review must be fully dispositioned completed before further operations begin.
11.2 Safety Assessment Phases
The three safety assessment phases required by Sea Launch are: • Definition phase • Development phase • Verification phase This section defines responsibilities and completion criteria for the customer and Sea Launch.
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Phase I ⎯ The responsibilities of the customer during the definition phase of the Definition Sea Launch Safety Assessment are to: • Review the Sea Launch safety requirements and identify applicability and tailorings necessary • Provide a system safety plan in compliance with Sea Launch safety requirements with final definition of safety data deliverables and re- view/approval schedules • Prepare and submit documentation addressing all hazards associated with the spacecraft and spacecraft processing and the requirements to control the identified hazards in the preliminary spacecraft design The responsibilities of Sea Launch during this phase are to: • Review for approval the definition phase of the customer prelaunch safety package and coordinate questions and issues with the customer for resolution • Compare the customer equipment designs and hazard control re- quirements with Sea Launch safety requirements • Accept the customer safety plan This definition phase will be completed when the safety plan is accepted and methods to identify all hazards, hazard causes, and the means for eliminating, reducing or controlling the hazards have been defined.
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Phase II ⎯ The customer responsibilities during the development phase of the Sea Development Launch Safety Assessment are to: • Submit prelaunch safety package documentation describing final haz- ard definition, detail hazard control requirements and safety-critical spacecraft system design information • Provide results of a risk-based hazard analysis of spacecraft design, integration, operations procedures and interface with Sea Launch systems and equipment per the Sea Launch Safety Regulations Man- ual • Define safety-critical GSE with qualification and acceptance re- quirements • Define qualification and training requirements for each job function in hazardous operation procedure • List tailorings, waivers/deviations for unmitigated risks and Sea Launch closure of safety noncompliances • List hazardous materials with quantities and type of containers aboard the spacecraft, hazardous materials used during ground op- erations and provide copies of Material Safety Data Sheets (MSDSs) • Submit spacecraft propellant tank, pressure tank and pressurized sys- tem design data and system verification that is in accordance with recognized industry design guidelines and meet Sea Launch system payload flight and ground environments per the Sea Launch Safety Regulations Manual • Define design data, safety factors, recognized industry standards and test methods employed in constructing and verifying spacecraft me- chanical, electrical and propellant processing GSE • Define planned hazardous systems manufacturing, qualification and acceptance verification methodology (e.g., hazard analyses, test re- ports and inspection records) that satisfy the hazard control require- ments and system design • Define safety-critical operational methodology and systems to be used for spacecraft processing and for hazard mitigation. Include pre- liminary identification of procedures to be used at Home Port and on board the ACS and LP to control hazardous spacecraft processing operations • Submit preliminary hazardous operations procedures for Sea Launch review and approval • Confirm that the maximum credible events during spacecraft hazard- ous operations are in compliance with the Sea Launch Risk Manage- ment Plan
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• Submit a preliminary Process Safety Management Plan in compli- ance with the requirements of California Code of Regulations Proc- ess Safety Management Plan requirements for propellant storage and process – California Occupational Safety and Health Administration (OSHA) 5189 • Submit explosive license/permit for the matchmate test site as well as explosive policies/procedures for the matchmate test site and any site specific safety procedures, if a matchmate is required The responsibilities of Sea Launch safety during this phase are to review and coordinate questions and issues with the customer for the approval of the following: • Development phase of customer prelaunch safety package • Preliminary hazardous procedures • Customer process safety management plan • Customer equipment designs and hazard control requirements
The development phase is deemed complete when design and operational hazard risk levels, spacecraft safety-critical system design, preliminary hazardous procedures and required hazard controls have been identified and approved by Sea Launch. This phase is to be concluded at the mission critical design review.
Phase III ⎯ Customer responsibilities for the verification phase of the Sea Launch Verification Safety Assessment are to: • Submit verification and acceptance documentation indicating com- pliance to established Sea Launch safety requirements as defined in Phase II • Submit final hazardous operational procedures in accordance and compliance with California OSHA regulations • Provide specific operations supporting information relative to hazard- ous materials, personnel qualifications and training, and critical support equipment certification. Individual customer personnel training for use and maintenance of equipment and appropriate medical evaluations are to be identified in the qualifications and training certification records • Submit spacecraft operations final Process Safety Management (PSM) Plan • Submit spacecraft Risk Management Plan (RMP) certificate
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The responsibilities of Sea Launch during this phase are to: • Review and approve the verification phase of the customer prelaunch safety package and to coordinate questions and issues with the cus- tomer for resolution. Sea Launch will review and compare verifica- tion data to safety requirements and verifications identified in Phases I and II. • Review and approve operations support information relative to haz- ardous materials, personnel qualifications and training, and critical support equipment certification. Personnel qualifications and training includes all of the qualifications and training certification records for each individual participating in customer hazardous operations as well as applicable medical evaluations. • Review and approve the final spacecraft hazardous procedures for compliance with California OSHA regulations • Review and approve final spacecraft PSM Plan for compliance with California OSHA regulations • Accept spacecraft RMP certificate • Report spacecraft propellant, pressurant, batteries and heat pipe qualification and acceptance data to the U.S. Department of Trans- portation During this verification phase, the spacecraft interface hazard analysis will be completed and closed. After implementation of all necessary changes and approval of operating procedures and supporting data, Sea Launch Safety will approve the planned customer operations. This phase is anticipated to conclude at the ground operations readiness review and must be completed before the start of spacecraft processing at Sea Launch facilities. Following completion of Phase III and the readiness review, customer- processing operations at Home Port may begin.
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11.3 Launch Licensing and Flight Safety
Authority and Sea Launch operates under a license issued by the U.S. Department of Licensing Transportation, Office of the Associate Administrator for Commercial Space Transportation (FAA/AST).
Licensing Information supplied by the customer is incorporated into a mission- Requirements specific application package submitted to the FAA/AST by Sea Launch in satisfaction of licensing requirements. This information includes: • Data on the breakup characteristics of the spacecraft, to verify that casualty probabilities remain within program FAA guidelines, should a launch failure occur • Definition of spacecraft ownership and operator, final orbit and li- censing for radio frequency transmissions • Listing of hazardous materials on board the spacecraft to support review of the potential environmental impacts in the event of a launch failure
FAA/AST All launch licenses issued by the FAA/AST require review of the potential Compliance environmental impacts of the Sea Launch system and launch operations. Spacecraft hazardous materials are considered to have potential impacts on the environment should a launch failure occur. The FAA/AST conducted an environmental assessment to cover the entire scope of Sea Launch operations and a finding of no significant impact was issued. Spacecraft environmental characteristics must not compromise the terms of this finding. Sea Launch complies with FAA/AST requirements to minimize orbital debris by demonstrating a contamination and collision avoidance maneuver of the orbiting upper stage (Block DM-SL). The contamination and collision avoidance maneuver moves the stage away from the spacecraft to a safe distance to a disposal orbit. At the completion of the contamination and collision avoidance maneuver, all Block DM-SL main engine propellants and gases are vented, all valves are latched open, and all power is depleted.
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Launch Operation Before and during launch operations, Sea Launch coordinates with Coordination appropriate U.S. Government agencies to issue warnings in the interest of public safety. Coordination includes identification of four points along the trajectory: • Immediate launch area • Stage 1 impact area • Fairing impact area • Stage 2 impact area Sea Launch coordinates launches with the following agencies: • FAA/Central Altitude Reservation Function • United States Space Command (through NASA) • National Imagery and Mapping Agency FAA/AST representatives accompany the launch team on board the Assembly and Command Ship at the launch site to fulfill their responsibilities for oversight of the launch and flight. Sea Launch assists onboard FAA/AST representatives in compliance monitoring of the terms and conditions of the license.
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11.4 Safety of Operations
Operational Safety Operational safety requirements governing customer operations in Sea Requirements Launch facilities are defined in Sea Launch safety reference documents. These requirements ensure safety of personnel and equipment in preparing and launching the spacecraft. The customer responsibilities for activities that occur on Sea Launch facilities and vessels are to: • Coordinate all activities with Sea Launch operations and safety per- sonnel • Identify all hazardous spacecraft processing operations anticipated on Sea Launch facilities and vessels • Develop customer and spacecraft-unique procedures • Personnel training (conducted by both Sea Launch and the customer) • Provide and maintain all unique personnel protection equipment re- quired to support spacecraft operations independent of or in addition to normal Sea Launch facility and vessel operations • Ensure that requested Sea Launch-provided personal protective equipment and services are identified in the applicable interface re- quirements document and spacecraft campaign document • Provide a safety contact within the processing team who will work with the Sea Launch safety organization at Home Port and on board the vessels Sea Launch will monitor safety during customer hazardous operations at Home Port.
Safety Training Customer personnel shall be skilled in spacecraft processing operations conducted on Sea Launch facilities and vessels. Customer personnel working with Sea Launch equipment in Sea Launch facilities are required to complete orientations in equipment familiarity, emergency procedures and facility-unique constraints. Customer personnel needed for operations on board Sea Launch vessels will be provided specialized safety training for operations and emergency procedures unique to those vessles and maritime environments.
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12.0 MISSION ASSURANCE
Overview Qualification, acceptance, quality, safety and reliability are all subject to careful and systematic scrutiny and oversight of the Sea Launch Mission Assurance organization. The Sea Launch Mission Assurance organiza- tion is the independent overseer of all Sea Launch processes and activi- ties that effect mission success. Mission assurance is a continuous effort that starts with the acceptance of qualified hardware and continues through the mission post-flight review and the implementation of correc- tive actions. Through these activities, Mission Assurance ensures that the Sea Launch system, and specifically launch vehicle reliability, is main- tained through a controlled set of established processes and procedures.
Mission Assurance Organization of Sea Launch Mission Assurance represents the culmina- Organization tion of years of effort to blend the best practices of Boeing and other U.S. companies, European companies, and companies from Russia and Ukraine. Figure 12-1 indicates the system of checks and balances in place for the review process. Each partner and contractor conducts its initial reviews according to its own established and proven processes and Sea Launch then performs an oversight supervisory evaluation of those processes. This disciplined approach of multiple independent layers of checks and verifications results in superior quality controls, which leads to higher probability of mission success.
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Engineering Vice President Chief Systems Engineer
Systems Mission Mission Advanced Engineering Assurance Integration Systems Director Director Director Director Systems Engineering Mission Assurance Mission Integration Advance Systems
Audits Independent reviews Insurer communication
1. Certification 1. Hardware acceptance 1. Problem resolution management 2. Reliability and maintainability 2. Production review 2. Bi-annual briefing 3. First flight items 3. Engineering review 3. SRR briefing 4. Flight anomaly 4. Risk management 4. MRR briefing 5. Postflight 5. LRR briefing 6. Mission design 6. FFI letter 7. Mission readiness 7. Action items closure 8. System readiness 8. Post-MRR letter 9. Ground operation readiness 10. Combined operation readiness 11. PLU transfer readiness 12. ILV transfer readiness 13. Launch readiness 14. Postflight 15. Mission flow
Denotes independent Denotes direct oversight and review line responsibility responsibility SHG0681874-066 Figure 12-1. Sea Launch Mission Assurance
Customer An essential part of Sea Launch mission assurance is the involvement of Participation the customer in the mission assurance process. The purpose is to under- stand all customer concerns and to assure the customer that everything reasonably possible is being done to assure mission success. Sea Launch welcomes and encourages the customer’s involvement. Sea Launch also provides formal written reports to the customer and its insurer.
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Qualification The Sea Launch mission assurance methodology starts with the accep- Status tance of qualified flight-proven hardware from Sea Launch partners and suppliers. All hardware used in a mission passes qualification through analysis, similarity to other flight vehicles, and qualification through test- ing. During the development of the hardware and through its subsequent testing, Sea Launch assures that the integrated launch vehicle compo- nents are qualified to levels during flight with significant margins of safety. The successful flights of the Sea Launch system and its record of on-time, accurate spacecraft insertion is evidence of Sea Launch’s thor- ough system qualification process.
Hardware Each set of campaign hardware is accepted in compliance with the Sea Acceptance Launch acceptance plan that sets forth the verification process for accep- tance of flight hardware and assurance of conformity to all Sea Launch specified requirements. The Sea Launch hardware acceptance team conducts a thorough, inde- pendent review of flight hardware non-conformance reports and key ac- ceptance test results, and physically inspects the hardware in the factory. Upon completion of hardware acceptance, Sea Launch certifies the hard- ware for its designated mission or other purpose. The Sea Launch Chief Systems Engineer typically assigns hardware to a specific mission seven months prior to launch.
Launch Vehicle Launch vehicle certification is an essential element of mission assurance. Certification The responsible contractors certify each launch vehicle segment prior to Sea Launch acceptance. This certification process, shown in Figure 12-2, uses a bottom-up approach wherein subsystem testing and certification are followed by system-level, stage-level and finally, the fully integrated launch vehicle testing and certification. Configuration management is also an essential element in the acceptance process. Sea Launch partner contractors Energia and Yuzhnoye use their own traditional, established system of records called “technical pass- ports” to track each subsystem, system and element under their authority. The technical passport is the pedigree of the processes and procedures performed at the partners’ sites. The technical passport travels with its respective component during each stage of production in the factory.
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Certification method Documentation Production
• Biser-3 • Biser-3 • Batteries Components • Accelerometers • Accelerometers • Qualification tests • Separation joint Passports • Gyroscopes • Gyroscopes • Acceptance tests •Processors • Engines • Engines • Independent quality review • SC adapter • Tanks • Tanks
Sea Launch/Boeing hardware acceptance review Passports Zenit and Block Payload • Acceptance tests stage 1 stage 2 DM-SL unit • Independent quality review
Integrated launch vehicle • Interface verification tests Certification of • Integrated functional mission Zenit and Block Payload checkout readiness stage 1 stage 2 DM-SL unit • Sea Launch review and acceptance
SHG0681874-067 Figure 12-2. Sea Launch Hardware Certification and Acceptance Process
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Sea Launch It is the policy of Sea Launch to systematically integrate quality into eve- Quality Assurance rything we do, individually and collectively, through all mission phases, Program from identification of market needs to final satisfaction of customer and regulatory requirements. The Sea Launch Quality Management System is process-based. It is a results- and improvement-oriented, customer-focused approach to build- ing confidence in our systems while identifying and validating improve- ments and creating and maintaining customer satisfaction. Objectives of the Sea Launch Quality Management System are: • Customer expectations are clearly understood and are achieved or exceeded • Program processes deliver safe, on time and flawlessly executed launches for every customer, every time • Sea Launch team members have defined, documented and re- leased procedures for all aspects of their work and scrupulously follow these procedures in the performance of the work • Assurance that the proper level of quality oversight is applied to critical processes and operations • Rigorous development and maintenance of documented history of all required activities • Establishment of appropriate metrics to track achievement of these objectives, identify weaknesses, and measure progress to improve our processes and systems Sea Launch management provides the leadership, resources, process management, training and environment necessary to ensure that each Sea Launch employee has ownership, personal responsibility and account- ability for the quality of their work. This includes the capability to con- trol and continuously improve the quality of their processes and services and a full understanding of the metrics by which their performance is evaluated. It is the responsibility of every member of the Sea Launch team to ensure the quality and integrity of our processes and to seek innovative ways to improve our methods in order to deliver safe, on-time, cost-effective and flawlessly executed launches to every customer, every time.
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Sea Launch In addition to the Sea Launch Quality Assurance Program for Sea Quality Oversight Launch hardware and operations, Sea Launch maintains oversight and insight into the quality programs of its partner companies. • RSC Energia is ISO 9002 Registered, utilizes an internal quality group and on-site government quality inspectors, and upholds CIS Government (GOST) standards. • BCSC is ISO 9001 Registered, utilizes an internal quality group and Boeing independent quality inspectors, and upholds Boeing Quality processes. • SDO Yuzhnoye provides engineering oversight, while the Yuzhmash quality group oversees production. Yuzhmash utilizes on-site government quality inspectors, and upholds CIS Govern- ment (GOST) standards. Sea Launch Mission Assurance supported by Sea Launch Quality Assur- ance ensures these high quality standards are upheld through hardware acceptance reviews at the factory and thorough quarterly production re- views.
Reliability The predicted and demonstrated reliability of the Zenit-3SL is compara- ble with that of other heavy-lift launch vehicles. The high reliability of the Sea Launch system is the result of years of evolved problem solving, enhancements and best practices for both Zenit stages and the Block DM. The Block DM-SL upper stage and Zenit-2S each have minor modifica- tions to adapt them for use within the Sea Launch system. Please contact your Sea Launch representative for the current reliability estimates.
Reliability Zenit-3SL reliability is enhanced by the use of mature subsystems and Subsystems and components, the extensive use of redundancy for fault tolerance in Components mission-critical areas, ample design parameter margins (e.g., strength, stability and propulsion), manufacturing quality monitoring, and exten- sive testing of flight hardware.
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Anomaly The Sea Launch anomaly resolution system, based on the existing Boe- Resolution System ing failure review, analysis and corrective action system (as used on sev- eral integrated air vehicles and avionics platforms), has been established along with a system database. The anomaly resolution system is in place to acquire and assess off-nominal history, criticality and effectiveness of changes, and to be used as a database from which recommendations can be validated. The anomaly resolution system is also used to gather data on test and flight failures, design changes, test results and repair deci- sions. All in-flight anomalies are dispositioned prior to launch. Details of such anomalies are typically shared with the current and subsequent launch customer.
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12-8 HPD-10041 REV D APPENDIX A. CUSTOMER INPUTS In order to successfully integrate a spacecraft with the Sea Launch system, a spacecraft-to-Sea Launch Interface Control Document (ICD) will be prepared by Sea Launch. In addition to the ICD deliverables, spacecraft models such as a CAD model, spacecraft thermal math model, spacecraft external surfaces geometric math model, and a dynamic model will be required. A full list of deliverables required from the customer will be described in the mission-specific Statement of Work. A two-cycle mission integration activity will include a preliminary design review and critical design review based on two releases of the ICD. A one-cycle mission integration activity will include a critical design review based on only a single release of the ICD. The customer will be required to provide the following types of information at the time of contract signing or shortly thereafter to be included in the ICD. To assist the customer in providing these inputs, Sea Launch will provide a draft ICD at the beginning of the integration activity.
Flight Design Data Item Contents Targeting • Nominal spacecraft mass specification • Apogee radius (injection orbit) • Perigee radius • Inclination Launch Window • Launch times or parameters that define launch times Spacecraft mass • Spacecraft mass properties • Three axis center of gravity • Three-axis moments and products of inertia about center of gravity • Mass properties uncertainty ranges Spacecraft • Spin rate Separation Data • Nominal angular momentum attitude • Angular momentum attitude error • Time average sun angle • Spacecraft pointing requirements
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Electrical/RF Interface Data Spacecraft • Connector part numbers (both halves) umbilical definition • Umbilical pinout and signal nomenclature • Wiring shield arrangements • Voltage, Current, Resistance and capacitance allocations Spacecraft • Baseband signal characteristics command and • Discrete command requirements telemetry requirements • Any special instrumentation requirements Spacecraft • Electrical power requirements electrical ground • Dimensions and mass support equipment definition • Desired locations • Remote command and control communication system requirements • Communication links between ACS, LP, and customer facilities • Test configurations and test support requirements Wiring/cabling • Electrical wiring diagrams and schematics for electrical equipment requirements and that interface with or through the launch vehicle or facility wiring signal characteristics • Types of connectors, cables, wires, shields, bonding, and grounding Spacecraft • Conducted and radiated emissions and immunity data on the electromagnetic spacecraft system, the spacecraft individual components, or both compatibility data • EMC design approach and requirements • EMC test data and damage thresholds with frequency and power levels Spacecraft radio • Characteristics of transmitters, receivers and antennas (including frequency system antenna location, boresight and beamwidth) characteristics • Spacecraft radiation and reception requirements • Radio frequency protection techniques • Radio frequency circuit losses for transmit and receive Radio frequency • On/off times of the radio frequency systems, including any time systems on/off during autonomous testing, checkout and actual countdown timeline
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Mechanical Interface Documentation and Requirements Fluid/propellant • Inventory of fluid/propellant types, mass and densities systems • Three-axis center of gravity for fluids in each tank with fluid types characteristics specified • Tank capacities and geometry • Parameters of maximum loadable mass • Ullage/fill fraction Coordinate systems • Definition of spacecraft orientation (coordinate systems and clocking) and clocking relative to the ILV based on analysis of ascent heating, access, preliminary CLA results and restrictions on spacecraft orientation during horizontal processing PLF upper access • Quantity and location of upper PLF access doors doors defined PLF lower access • Quantity and location of lower PLF access doors doors defined Mating surface • Definition of mating surface and push-off characteristics Customer Logo • Customer logo for application on the PLF, delivered in electronic media (Adobe Illustrator Vector File Format; .eps or .ai file format)
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Thermal, Venting, and Contamination Data Spacecraft thermal • Facility requirements requirements • Dedicated provisions (e.g., cooling air, purges) • Prelaunch thermal soak requirements • Maximum airflow velocities • Thermal interface requirements and constraints • Maximum aero-heating/free molecular heating/solar exposure constraints Spacecraft venting • Displaced volume requirements • Internal cavity volume • Airflow characteristics between spacecraft internal cavities • Maximum allowable values for: pressure differential between cavities, residual pressure in fairing at jettison, and rate of pressure decay in fairing Spacecraft • Facility requirements contamination requirements • Spacecraft surface molecular and particulate contamination requirements at launch and at spacecraft/upper stage separation • Planned access to spacecraft after encapsulation • Locations of spacecraft contamination-sensitive surfaces • Bagging and covering procedures during ground processing Launch Site/Operations Data SC Safety Data • Safety Integration Control Agreement (formerly referred to as the Memorandum of Agreement) Nonstandard • Power, communications, air conditioning and any deviations or Spacecraft facility additions to standard services or equipment provided by Sea Launch and processing and described in the Sea Launch Facilities Handbook and Facility requirements Cable Drawings • Deviations or additions to standard Sea Launch processing operations
A-4 HPD-10041 REV D APPENDIX B. SEA LAUNCH STANDARD OFFERINGS AND NONSTANDARD SERVICE OPTIONS
This appendix details the Sea Launch standard offerings as well as nonstandard service options, and provides a top-level description of all services offered by Sea Launch. A detailed description of such services is set forth in the Sea Launch Statement of Work which, when incorporated into an executed Launch Services Agreement (LSA), sets forth the actual work scope to be provided to the customer. Standard offerings are those items that comprise the Sea Launch all-inclusive, commercial launch service and are typically considered to be in-scope to an executed LSA. Conversely, nonstandard service options are considered to be outside the scope of the standard LSA, but may be provided upon request to address the customer’s unique mission requirements. Nonstandard services are subject to additional consideration. • Standard-offering hardware and software • Standard-offering launch services • Standard offering facilities and support services • Nonstandard service options
Standard-Offering: Hardware and Software Services
The subsections that follow identify the standard hardware and software items and associated services that are typically used to provide launch services in support of a customer mission.
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Payload Standard accommodations include: Accommodations • One payload fairing, corresponding to requirements stated in Section 7, including: • Up to two access doors (one per half shell) • Two customer logos (2.95 meters tall by 2.69 meters wide). One logo is positioned at –Zplu (starboard side) and one is positioned at +Yplu (forward side) • Acoustic blankets • Spacecraft adapter with spacecraft compatible mechanical interfaces and one separation system • Current standard-offering spacecraft adapters are: SCA1194, SCA1194VS, SCA1666, SCA 1666MVS, SCA702+, SCA702+GEM and SCA937S • The clocking capability of the PLA allows any spacecraft to be clocked relative to the PLU at any angle within 2.5° increments from the PLU +Z axis • Payload Structure • Interface Skirt • In-fairing re-radiation system, which accommodates spacecraft command and telemetry frequencies in the 3 to 18 GHz range. The frequency used is mission dependent and the amplifiers have bandpass filters to pass the desired frequency. Operational capability outside this range may be accommodated as a non-standard option.
Launch Vehicle Sea Launch vehicle configuration includes: • One Zenit two-stage booster with mission-specific flight software • One Block DM-SL upper stage with mission-specific flight software
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Electrical Sea Launch electrical interfaces include: Interfaces • Radio frequency interfaces • Radio frequency (command and telemetry) links between spacecraft and spacecraft electrical ground support equipment in the Sea Launch Payload Processing Facility and when the integrated launch vehicle is erect at the launch site by means of the Spacecraft Radio Frequency Support System (SCRFSS). • Ship to ship 802.11G wireless bridge link: This IEEE standard link is available for customer use for establishing Ethernet connectivity between customer equipment located on the Launch Platform and the Assembly and Command Ship. Throughput varies with distance and weather conditions but is generally acceptable for command and telemetry traffic up to approximately 1.0 Mbps. At the time of publishing, this option is still under development to be implemented on all missions. Please contact your Sea Launch representative for status. Hard-line interfaces include: • Two in-flight disconnect connectors for telemetry, command and power interfaces to the spacecraft • Hard-line (command and telemetry) umbilical links • Up to two serial data telemetry streams • Up to four analog telemetry channels • Up to eight discrete telemetry inputs • Up to four redundant (primary and backup) commands • Two in-flight disconnect break wires (thirty-seven or sixty-one pin electrical connectors)
Environmental Sea Launch will maintain spacecraft environments within the limits Controls specified in the spacecraft/Sea Launch interface control document while the spacecraft is in the Payload Processing Facility and encapsulated in the payload unit, throughout mission integration up to launch.
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Interface Test • For each first-of-type spacecraft, Sea Launch will perform one matchmate test at the spacecraft factory that includes the spacecraft adapter and a flight-equivalent separation system. The matchmate test consists of mechanical and electrical interface testing between the spacecraft and the spacecraft adapter, shock transmission tap-tests and spacecraft separation test. • First-of-type indicates a spacecraft bus that has not been launched previously on a Zenit-3SL. Exceptions may be made for spacecraft bus types that have been previously integrated through matchmate, but have not yet launched (see Section 2, Table 2-1).
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Standard-Offering Launch Services The following subsections identify the basic launch services that are provided for a Sea Launch mission.
Mission • Sea Launch management—includes program-level services such as Management contract administration, scheduling and program management. • Mission management—includes coordination and oversight of mission-specific activities such as integration, analysis and operations. This also includes reasonable support by Sea Launch to customer insurance placement requirements.
Meetings and Meeting and milestone reviews include (all meetings held at Sea Launch Reviews Home Port unless otherwise noted): • Customer kickoff meeting • Customer site survey • Customer hardware acceptance review • System readiness review (customer attendance is optional) • Two ground operations working groups • Technical interchange meetings as required • Ground operations readiness review (ready to receive spacecraft) • Combined operations readiness review (ready to mate spacecraft to payload accommodations) • Payload unit transfer readiness review (ready to integrate payload unit to launch vehicle) • ILV transfer readiness review (ready to transfer the ILV from the Assembly and Command Ship to the Launch Platform) • Mission readiness review (ready to leave Home Port; customer attendance is optional) • Launch readiness review (ready to launch) held on board the Sea Launch vessels at the launch site • Postlaunch customer review (meeting is on request, a postlaunch report is a standard deliverable)
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Documentation Generic The following documents will be provided to the customer according to the agreed-to mission integration schedule, export licensing permitting: • Sea Launch Facilities Handbook for customers • Electromagnetic Compatibility (EMC) Control Plan • Contamination Control Plan for Sea Launch Operations • Sea Launch Safety Regulations Manual • Sea Launch System Safety Plan • Standard Security Guidelines for Customers at Sea Launch Mission specific The following documents will be provided to the customer according to the mission integration schedule (developed by Sea Launch with customer concurrence after mission identification), export licensing permitting: • Mission integration schedule • Spacecraft/Sea Launch system interface control document • Spacecraft Campaign Document (SCD) • Spacecraft Test Requirements Document (STRD) • Launch schedule • ILV Home Port operations master procedure (MP-05) • Launch description documents (LDD-01, LDD-02) • Launch commit criteria (LCC-01, LCC-02) • Countdown Master Procedure (MP-02) • Plus Count Master Procedure (MP-03) • Matchmate plan (if required)
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Mission Analyses Integration The following analyses will be conducted and presented at design review meetings. (Unless noted otherwise, each analysis will involve two cycles, preliminary and final, for first-time integration and one cycle for repeat or similar mission integration. If more than one specific analysis is required per cycle, this is subject to a contractual change.) • Mission design • Spacecraft acoustic analysis • Payload fairing separation analysis* • Spacecraft separation analysis • Critical clearance analysis (dynamic) • Coupled loads analysis • Spacecraft shock analysis • Spacecraft vibration analysis • SC – PLA clearance analysis (static, for encapsulation) • Flight thermal and pressure analysis • Venting analysis • Contamination assessment • Electromagnetic compatibility assessment • Avionics and electrical interface assessment* • Spacecraft on-pad link analysis* • Operational analysis of mission specific changes • Post-flight analysis* (not presented at design review meetings as it is conducted after the launch) *One cycle for both first-time and repeat integration Verification Sea Launch will perform verification of spacecraft requirements as defined in the spacecraft/Sea Launch system interface control document. Separation state vector A spacecraft separation state vector is provided to the customer within fifty minutes of separation. Test Support Sea Launch will provide personnel to support spacecraft sine/vibe testing.
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Transportation Sea Launch provides: • Ground transportation for the spacecraft and support equipment from Long Beach, Los Angeles or Ontario airports to Home Port, and return of support equipment to the airport after the launch • Coordination with the US Department of Homeland Security including its US Customs and Border Protection Agency for the spacecraft and its associated personnel and support equipment
Home Port Primary prelaunch services conducted at Home Port include: Activities • Fit check—a spacecraft-to-spacecraft adapter fit check • Support to accommodate spacecraft manufacturer’s tasks for checkout and fueling • Spacecraft mate to the SCA and IS/PS • Encapsulation • Payload accommodations checkout • Payload accommodations/launch vehicle integration • Integrated launch vehicle integrated test • Transfer of the integrated launch vehicle to the Launch Platform • Payload accommodations and spacecraft radio frequency end-to-end test
Transit Activities At-sea operations provided by Sea Launch include: • The Assembly and Command Ship and Launch Platform will make one round-trip transit from Home Port to the launch site, including transportation of ground support equipment and customer launch support personnel • Two launch countdown rehearsals (plus one at Home Port prior to departure) • Monitoring of spacecraft environments and support of spacecraft operations
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Standard-Offering Facility Support Services
The following subsections identify the facilities and support services available to each Sea Launch customer.
Home Port (HP) • Home Port facilities are described in detail in the Sea Launch Facilities Handbook. Facilities at Home Port are available to the customer for approximately sixty days during the launch campaign. All customer equipment shall be removed from Sea Launch facilities within three workdays after the vessels return to Home Port following launch. • Customer operations may be scheduled on a twenty-four hour basis with access to the customer office building and the Payload Processing Facility.
HP Facilities • Processing and fueling cell • Control room • Remote control room • Offices • Conference room • Storage
HP • Intercom Communications • Spacecraft command and telemetry cabling • Closed-circuit television monitors and cameras • Telephone • Launch video • Countdown net • Handheld radios • Lightning pagers
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HP Security • The facilities have twenty-four hour perimeter security. • Security devices are installed on all internal and external doors leading to the payload processing areas. • The facilities have twenty-four hour electronic monitoring for ingress and egress, and smoke and fire detection.
HP Support • Separate technical and facility electrical power Services • Use of Home Port mechanical support equipment (e.g., manlifts and forklifts) • Access to video and still photography services, subject to Sea Launch Security restrictions • Approved copy of Sea Launch video highlights of the launch campaign • Protective garments for cleanrooms and hazardous operations, including fueling • Negotiated quantities of gaseous nitrogen, liquid nitrogen, gaseous helium, isopropyl alcohol, and other general-purpose cleaning agents and solvents • Analysis of up to ten samples of gases, propellants or cleaning materials • Monitoring system for detection of hazardous vapors • Receipt and staging of spacecraft propellants (thirty-day limit for propellants on site) • Receipt and storage of ordnance components • Use of calibrated weights • Photocopier and facsimile machine
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Communications Sea Launch provides the following communications support for mission Support coverage • Press outreach, beginning with the transit to the launch site, continuing throughout launch operations.
• Dedicated mission webpage on the Sea Launch website, featuring background information about the customer and its satellite, a photo album illustrating operations at Home Port and at sea, and mission specific flight illustrations.
• Launch broadcast/webcast, including live coverage from countdown to spacecraft acquisition, pre-packaged features about the mission and its team, interviews with key customer participants; and customer-provided video about the satellite, its footprint, applications and market potential, as well as marketing/advertising pieces. This program is fed into the Sea Launch-hosted customer event for launch viewing.
• Sea Launch provides a satellite feed of the broadcast program to the point of downlink at the customer's headquarters, and satellite coordinates to the customer and the media for downlinking capabilities. The broadcast program is also hosted live on the Sea Launch website, for global access.
Launch Event Sea Launch provides full coordination and event planning of a combined customer/employee event. The launch event is held at the Sea Launch Home Port Facility in conjunction with the launch broadcast.
Assembly and Assembly and Command Ship capabilities are described in detail in the Command Ship Sea Launch Facilities Handbook for customers. (ACS)
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ACS Facilities • Customer ground support equipment room • Customer conference room • Customer offices as shown in the Sea Launch Facilities Handbook • Berthing to accommodate up to eleven spacecraft personnel; fourteen personnel during Launch Platform evacuation at time of launch. The accommodations include the following: • VIP Cabin (2) • First Class Cabin (1) • Regular Single Occupancy Cabin (3) • Double Occupancy Cabins (3) • Three positions in the launch control center (each console can accommodate two people) • Two seats in the prominent launch control center guest viewing area • Storage
ACS Shared Areas • Escorted access to launch vehicle assembly compartment (Room 101) • Lounge • Observation areas • Dining facility • Recreational facilities
ACS Internal The following communication links are available within the Assembly Communications and Command Ship and between the Assembly and Command Ship and the Launch Platform when they are in proximity: • Intercom (including countdown net) • Closed-circuit television monitors and cameras (including launch video) • Telephone
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ACS External Sea Launch has 256-kbps of bandwidth that may be configured to Communications customer specification. The following external links are available to the customer during the deployment phase, by way of a satellite link to
Brewster, Washington: • One countdown audio net link • 128-kbps local area network (configurable to 64-kbps RS422 serial digital data link, V.35, Ethernet and RS232) • 64-kbps RS422 serial digital data link • 64-kbps telemetry link between the Assembly and Command Ship and the Launch Platform • Five dedicated telephone lines with access to long-distance circuits • Internet connectivity for the spacecraft provider • Internet connectivity for the spacecraft owner
ACS Security • Card readers control access to various secure areas of the ship • Twenty-four hour electronic monitoring for smoke and fire detection
ACS Support • Meal service Services • Hospital (staffed by a medical doctor and a nurse) • Photocopier and facsimile machine
Launch Platform Launch Platform capabilities are described in detail in the Sea Launch (LP) Facilities Handbook.
LP Facilities • Customer ground support equipment room • Customer office • Berthing to accommodate up to three spacecraft personnel • Storage
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LP Shared Areas • Recreation and video room • Dining facility • Conference room
LP Internal The following communication links are available within the Launch Communications Platform and between the Assembly and Command Ship and the Launch Platform when they are in proximity:
• Intercom (including countdown net) • Closed-circuit television monitors and cameras (including launch video) • Telephone
LP External The following external links are available to the customer during the Communications deployment phase, by way of the satellite link to Brewster, Washington: • One countdown audio net link • One live launch video link • 128-kbps local area network (configurable to 64-kbps RS422 serial digital data link) • One dedicated telephone line with access to long-distance circuits • Internet connectivity
LP Security • Card readers control access to various secure areas of the vessel • Twenty-four hour electronic monitoring for smoke and fire detection
LP Support • Monitoring system for detection of hazardous vapors Services • Meal service • Hospital (staffed by a medical doctor) • Photocopier and facsimile machine
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Nonstandard Options
Nonstandard service options are considered to be outside the scope of the Sea Launch Launch Service Agreement (LSA). However, to address the customer’s unique mission requirements, the following nonstandard services may be provided by Sea Launch upon customer request, subject to additional consideration. Customer requests for nonstandard services not described in this document will be considered by Sea Launch on a case-by-case basis.
Mission Analysis Customer requests to repeat any standard offering mission integration analysis listed, or to perform any additional analysis or design work due to changes made, or for any other reason, may be provided on customer special request. Additional reporting requirements may also be provided.
Licensing Any missions that are considered to be “hybrid” or “dual use,” or spacecraft containing military components, are non-standard for Sea Launch. These missions require additional licensing, cost, and/or government approvals/authorizations.
Spacecraft Purge Sea Launch can provide localized purging to the spacecraft on the launch pad up to L-17 minutes (at the time of transporter-erector disconnect).
Directed Air While the ILV is in the Launch Platform hangar, Sea Launch can provide directed air for cooling the spacecraft batteries with a given flow rate and temperature setpoint. Modifications to the payload accommodations are required for this nonstandard option, therefore the request for this service must be provided no later than L-12 months.
Fairing If the satellite volume falls outside of the Sea Launch static envelope Modifications (Figure 7-1), Sea Launch may accommodate acoustic foam modifications or a “blister” on the fairing. Protrusions must be such that payload fairing encapsulation can still be accommodated.
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Additional Fairing The standard Sea Launch payload fairing provides up to two access doors Doors for spacecraft access in the Payload Processing Facility—one door per fairing half shell. As an option, the payload fairing can accommodate additional access doors beyond the standard quantity of two, as well as doors outside the areas specified in Section 7. Size, quantity and location of doors are subject to analysis.
Spacecraft Access Sea Launch can support environmentally-controlled access to the encapsulated spacecraft while in the Launch Platform hangar at the launch site and on the launch pad prior to ILV erection. For safety reasons, this option is available only while the Launch Platform is ballasted.
Communications Communications requirements outside those described as standard may be accommodated. Examples of previous nonstandard items are as follows: • One live launch video link, export compliance permitting • Hard-line command receiver tones • Additional ground support equipment cabling • Additional external communications interfaces
Fairing Logo Any customer requests for deviations from the standard offering described under payload accommodations will be evaluated on a case-by- case basis.
Nonstandard The performance for many scenarios is provided in Section 3 of this Trajectories guide and is for informational and planning purposes only. There are some described orbits that would require modifications to the Sea Launch system to accommodate specific requirements. Missions to these orbits are considered nonstandard. Some examples include: • Orbits that require a long-duration mission, such as geosynchronous orbits • Complex mission designs that would require multiple flight design scenarios for a single satellite • Trajectories that require moving the location of the launch site.
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Transportation In addition to the standard offering, Sea Launch can provide requested Handling nonstandard handling equipment. Equipment
Video Sea Launch has examined various possibilities for providing launch video from onboard the ILV. Sea Launch can make these options available as a nonstandard service upon customer request.
Interface Tests Matchmate and separation shock testing at the spacecraft factory for repeat of spacecraft type can be accommodated. Additional interface testing can also be accommodated.
Electrical • Spacecraft serial PCM telemetry codes other than NRZL with Interface accompanying clock (RS422) • Radio frequency link (SCRFSS) capability to the Assembly and Command Ship launch vehicle assembly room and the Launch Platform hangar. • Re-radiation capability in the 2 to 3 and 18 to 40 GHz range • Umbilical hard-line command and telemetry links other than baseband PCM
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Support Services Arrangements for the following services will be made by Sea Launch based upon customer request. Charges for requested services will be billed to the customer. • Long-distance telephone calls (from telephones, facsimile machines and teleconferencing facilities) • Access to Internet service provider • Bandwidth that exceeds standard offering • Office supplies • Office equipment (computers) • Secretarial services—bilingual for non-English-speaking customers • Translation and interpretation services • Telecommunications charges associated with forwarding data from the communications node in Brewster, Washington, to the destination • Standard equipment calibration services • Support of Home Port personnel during off-shifts (e.g., second or third shift on weekdays and any weekend shifts) • Analysis of particulate and nonvolatile residue in the cleanrooms • Customer-requested evacuation of personnel during marine operations • Specialized handling equipment
Materials • Gaseous nitrogen • Oxygen • Helium • Liquid nitrogen • Isopropyl alcohol • Deionized water • Solvents and gases in addition to negotiated quantities • Provision or importation of spacecraft propellant
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Facilities • Use of additional office space may be negotiated as available • Use of any facilities beyond the time frame allocated to the customer may be negotiated as available • Storage of the spacecraft beyond the time frame allocated to the customer may be negotiated as available
Launch Event Sea Launch can assist with coordinating an event at a location designated by the customer, or providing additional VIP services at the Home Port launch event as a non-standard service.
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APPENDIX C. WEATHER HISTORY AT THE LAUNCH SITE
This appendix provides historical data on weather conditions at the launch site normally used for launching. Additional data may be available at the website provided by the National Data Buoy Center at: http://www.ndbc.noaa.gov/station_page.php?station=51028.
Station 51028
Station 51028 is a weather buoy currently located at 0 degrees 0'1" N latitude, 153 degrees 54'46" W longitude, near the launch site. It is a 3- meter discus buoy with an ARES payload. It is owned and maintained by the National Data Buoy Center (NDBC). Weather data is collected by the NDBC and is available for public access. The data collected includes wind direction, wind speed, wind gust, wave height, dominant wave period, atmospheric pressure, pressure tendency, air temperature and water temperature.
Historical Weather Data
The following sections provide historical weather data gathered by the weather buoy at the launch site.
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Air Temperature Air temperatures at the launch site over the period from October 1997 to December 2006 are charted in Figure C-1.
Monthly Air Temperature Statistics from Oct 1997 to Dec 2006 32
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28
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24 Air Temperature (Celsius)
22
20 Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec Time (month) SHG0681874-068 Figure C-1. Historical Air Temperatures
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Wind Gusts Peak wind gusts at the launch site over the period from October 1997 to December 2006 are charted in Figure C-2.
Monthly Wind Gusts Statistics from Oct 1997 to Dec 2006 18
16
14
12
10
8 Wind Gusts(m/s) 6
4
2
0 Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec Time (month) SHG0681874-069 Figure C-2. Historical Wind Gusts
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Atmospheric Atmospheric pressure at sea level at the launch site over the period from Pressure October 1997 to December 2006 is charted in Figure C-3.
Monthly Pressure Statistics from Oct 1997 to Dec 2006 1018
1016
1014
1012
1010
1008 Pressure (mbar) 1006
1004
1002
1000 Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec Time (month) SHG0681874-070 Figure C-3. Historical Atmospheric Pressure at Sea Level
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Water Water temperatures at the launch site over the period from October 1997 Temperature to December 2006 are charted in Figure C-4.
Monthly Water Temperature Statistics from Oct 1997 to Dec 2006 34
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24 Water Temperature (Celsius)
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20 Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec Time (month) SHG0681874-071 Figure C-4. Historical Water Temperatures
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Average Wind Average Wind Speed, over 8 min averages as measured at the buoy, at Speed the launch site over the period from October 1997 to December 2006 is charted in Figure C-5.
Monthly Wind Speed Statistics from Oct 1997 to Dec 2006 16
14
12
10
8
6 Wind Speed (m/s)
4
2
0 Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec Time (month) SHG0681874-072 Figure C-5. Historical Average Wind Speed
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