section 1 introduction - geo ring

Land Launch User’s Guide ____________________________________________________________________________________________________________________________________________________________________________________________________________________________________ Original Release Date: 28 July 2004 Initial Revision

The Land Launch User’s Guide has been cleared for public release by the

Department of Defense, Directorate for Freedom of Information and Security Review, as stated in letter 04-S-1323, dated 28 July 2004.

Prepared by Sea Launch for distribution by:

Boeing Launch Services One World Trade Center, Suite 950

Long Beach, CA 90831, USA

on behalf of the Sea Launch Company, L.L.C.

Copyright pending by

Sea Launch Company, L.L.C.

Land Launch User’s Guide

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REVISIONS

LTR DESCRIPTION DATE APPROVAL

NC Initial release July 2004 James Ellinthorpe Program Manager


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TABLE OF CONTENTS

Page

1. INTRODUCTION ............................................................................................................ 1-1 Purpose................................................................................................................................ 1-1

1.1 Overview of the Land Launch System................................................................... 1-1 What the System Includes.......................................................................................... 1-3 Advantages to the Customer ...................................................................................... 1-3 Timeline ..................................................................................................................... 1-4 Baikonur Cosmodrome .............................................................................................. 1-5

1.2 Land Launch Organization..................................................................................... 1-7 Overview.................................................................................................................... 1-7 Sea Launch Company, LLC....................................................................................... 1-8 The Boeing Company ................................................................................................ 1-8 Space International Services, Ltd .............................................................................. 1-8 SDO Yuzhnoye .......................................................................................................... 1-9 PO Yuzhmash .......................................................................................................... 1-10 Design Bureau of Transport Machinery (KBTM) ................................................... 1-11 Center for Ground Space Infrastructure Operations (TsENKI)............................... 1-12 RSC Energia............................................................................................................. 1-13 NPO Lavochkin ....................................................................................................... 1-14 Russian Space Agency............................................................................................. 1-14

2. VEHICLE DESCRIPTION ............................................................................................. 2-1 Overview............................................................................................................................. 2-1 Design ................................................................................................................................. 2-2 Zenit Flight History ............................................................................................................ 2-2 Block DM Flight History.................................................................................................... 2-2 Flight Success Ratios .......................................................................................................... 2-3

2.1 Land Launch Zenit .................................................................................................. 2-4 Design Heritage ......................................................................................................... 2-4 Changes Made for Sea Launch .................................................................................. 2-5 Avionics ..................................................................................................................... 2-5 Overall Specifications and Configurations ................................................................ 2-5 2.1.1 Zenit Stage 1................................................................................................ 2-8 Overall Configuration................................................................................... 2-8 RD-171M Engine ......................................................................................... 2-9


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2.1.2 Zenit Stage 2.............................................................................................. 2-10 Overall Configuration................................................................................. 2-10 RD-120 Main Engine ................................................................................. 2-11 RD-8 Vernier Engine.................................................................................. 2-11

2.2 Block DM-SLB Upper Stage ................................................................................. 2-12 Overall Configuration .............................................................................................. 2-12 11D58M Main Engine ............................................................................................. 2-12 Attitude Control/Ullage Engines.............................................................................. 2-12 Avionics ................................................................................................................... 2-12 Changes Made for Sea Launch ................................................................................ 2-12 Block DM-SLB Versus the Block DM-SL.............................................................. 2-13

2.3 Zenit-3SLB Ascent Unit ........................................................................................ 2-15 Components and Integration .................................................................................... 2-15 Payload Fairing ........................................................................................................ 2-15 Fairing Access Characteristics................................................................................. 2-15 Conditioned Air Supply to the Fairing..................................................................... 2-16 Fairing Thermal Protection ...................................................................................... 2-16 Payload Structure Support ....................................................................................... 2-16

2.4 The Zenit-2SLB Payload Unit .............................................................................. 2-17 Components and Integration .................................................................................... 2-17 Payload Fairing ........................................................................................................ 2-17 Fairing Access Characteristics ................................................................................. 2-17 Conditioned Air Supply to the Fairing..................................................................... 2-18 Fairing Thermal Protection ...................................................................................... 2-18 Intersection Bay ....................................................................................................... 2-18 Spacecraft Adapters ................................................................................................. 2-18 Unique Interfaces and Multi-Spacecraft Launches.................................................. 2-18

3. PERFORMANCE............................................................................................................. 3-1 Overview............................................................................................................................. 3-1 Performance Ground Rules................................................................................................. 3-2 Launch Window Availability.............................................................................................. 3-3

3.1 Launch Site and Accessible Orbits......................................................................... 3-4 Site Location .............................................................................................................. 3-4 Accessible Orbits ....................................................................................................... 3-5


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3.2 Ascent Trajectory – Generic Zenit-3SLB GTO Mission...................................... 3-6 Mission Profile........................................................................................................... 3-6 Stage 1 Flight ............................................................................................................. 3-7 Stage 2 Flight ............................................................................................................. 3-8 Block DM-SLB Powered Flight ................................................................................ 3-9 Flight Timeline......................................................................................................... 3-10 Ground Track ........................................................................................................... 3-11

3.3 Ascent Trajectory – Generic Zenit-2SLB Mission to 51.6° LEO ...................... 3-12 Stage 1 Flight ........................................................................................................... 3-12 Stage 2 Flight ........................................................................................................... 3-12 Flight Timeline......................................................................................................... 3-13 Flight Profile ............................................................................................................ 3-14 Ground Track ........................................................................................................... 3-15

3.4 Payload Capability – Three Stage Zenit-3SLB ................................................... 3-16 Geosynchronous Transfer Orbit............................................................................... 3-16 MEO, HEO, Circular, and Elliptical Orbits ............................................................. 3-17 High-Energy and Earth-Escape Trajectories............................................................ 3-19

3.5 Payload Capability – Two Stage Zenit-2SLB...................................................... 3-20 Circular LEO Orbits................................................................................................. 3-20 Elliptical Orbits ........................................................................................................ 3-21

3.6 Coast Phase Attitude Maneuvers ......................................................................... 3-22 Zenit-3SLB............................................................................................................... 3-22 Zenit-2SLB............................................................................................................... 3-22

3.7 Injection Accuracy ................................................................................................. 3-23

3.8 Spacecraft Separation and Post-Separation Events ........................................... 3-24 3.8.1 Zenit-3SLB ................................................................................................ 3-24 Separation Event......................................................................................... 3-24 Separation Capabilities............................................................................... 3-24 CCAM ........................................................................................................ 3-24 State Vector Delivery ................................................................................. 3-24 3.8.2 Zenit-2SLB ................................................................................................ 3-25 Separation Event......................................................................................... 3-25 Separation Capabilities............................................................................... 3-25 CCAM for Second Stage............................................................................ 3-26 State Vector Delivery ................................................................................. 3-26


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4. SPACECRAFT ENVIRONMENTS .............................................................................. 4-1 Overview............................................................................................................................. 4-1 Ground and Flight Environments ....................................................................................... 4-1 Reference Coordinate System............................................................................................. 4-1 Environmental Monitoring ................................................................................................. 4-2

4.1 Structural Loads ...................................................................................................... 4-3 Overview.................................................................................................................... 4-3 Quasi-Static Load Factors, Ground Handling, and Transportation ........................... 4-3 Quasi-Static Load Factors, Flight .............................................................................. 4-4 Sinusoidal Equivalent Vibration During Flight ......................................................... 4-5

4.2 Random Vibration ................................................................................................... 4-6 Ground Random Vibration for Components Near Spacecraft Interface.................... 4-6 Flight Random Vibration Environment ..................................................................... 4-6

4.3 Acoustics ................................................................................................................... 4-8 Fairing Space Average Sound Pressure Levels.......................................................... 4-8

4.4 Shock ....................................................................................................................... 4-10 Overview.................................................................................................................. 4-10 Zenit-3SLB............................................................................................................... 4-10 Zenit-2SLB............................................................................................................... 4-13

4.5 Electromagnetic Environment.............................................................................. 4-14 Overview.................................................................................................................. 4-14 Coordination............................................................................................................. 4-14 Ambient Cosmodrome Electromagnetic Environment ............................................ 4-14 Launch Vehicle Radio Equipment ........................................................................... 4-17 Radio Frequency Environment at the SC Separation Plane..................................... 4-18

4.6 Spacecraft Thermal and Humidity Environments ............................................. 4-20 Introduction.............................................................................................................. 4-20 4.6.1 Ground Thermal and Humidity Environments .................................... 4-20

General Overview, Ground Thermal and Humidity Environments ........... 4-20 Facility Clean Air Systems......................................................................... 4-20 Transportation Clean Air Systems ............................................................. 4-20 Launch Pad Clean Air System.................................................................... 4-21 Impingement Velocity of Airflow Upon SC Surface................................. 4-21

4.6.2 Flight Thermal Environments................................................................. 4-24 General Overview, Flight Thermal Environment....................................... 4-24


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4.7 Pressure Venting .................................................................................................... 4-26 Overview.................................................................................................................. 4-26 Pressure Decay Rate ................................................................................................ 4-26 Pressure Differential at Fairing Jettison................................................................... 4-26 4.8 Contamination........................................................................................................ 4-30 Contamination Control During Ground Processing................................................. 4-30 Contamination Control During Flight...................................................................... 4-31 Fairing Design Features to Minimize Contamination.............................................. 4-31 Plume Impingement ................................................................................................. 4-31

5. SPACECRAFT INTERFACES....................................................................................... 5-1 5.1 Mechanical Interfaces ............................................................................................. 5-1 5.1.1 Mass Properties and Modal Frequencies ................................................. 5-1 Spacecraft Mass and Longitudinal Center of Gravity Location................... 5-1 Spacecraft Center of Gravity Radial Offset ................................................. 5-2 Modal Frequencies ....................................................................................... 5-2 5.1.2 Payload Fairing Mechanical Interfaces.................................................... 5-3 Payload Fairings ........................................................................................... 5-3 Useable Volume ........................................................................................... 5-7 Useable Volume Inside Payload Structure................................................... 5-9 Access Doors................................................................................................ 5-9 RF Windows................................................................................................. 5-9 Customer Insignia......................................................................................... 5-9 5.1.3 Spacecraft Adapters ................................................................................. 5-11 Saab Spacecraft Adapters ........................................................................... 5-11 Zenit-2 Adapter for Use with Zenit-2SLB ................................................. 5-13 Multi-Satellite Dispensers for Use with Zenit-2SLB ................................. 5-13 5.2 Electrical Interfaces............................................................................................... 5-14 Overview ................................................................................................................. 5-14 5.2.1 Hard Line Links (Spacecraft Umbilical)................................................ 5-14 Umbilical Circuits ...................................................................................... 5-14 Umbilical Use During Processing and Launch .......................................... 5-14 Umbilical Connectors................................................................................. 5-15 5.2.2 Radio Frequency Links............................................................................ 5-15 5.2.3 In-Flight Commands, Measurements and Telemetry ........................... 5-15 General ....................................................................................................... 5-15 Separation Verification............................................................................... 5-15 Satellite Environments Measurements ....................................................... 5-16 Commands.................................................................................................. 5-16 5.2.4 Electrical Power for EGSE...................................................................... 5-17 Ground Power............................................................................................. 5-17


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Uninterruptible Back-up Power.................................................................. 5-17 5.2.5 Bonding and Grounding .......................................................................... 5-18 Bonding ...................................................................................................... 5-18 Grounding................................................................................................... 5-18

6. LAND LAUNCH FACILITIES....................................................................................... 6-1 Overview............................................................................................................................. 6-1

6.1 Transportation of Personnel and Cargo to and From Baikonur ........................ 6-2 Krainy Airport ........................................................................................................... 6-2 Yubileiny Airport....................................................................................................... 6-2 Transportation at the Cosmodrome............................................................................ 6-3 6.2 Site 31 Payload Processing Facility ........................................................................ 6-4 Overview.................................................................................................................... 6-4 Buildings 40/40D, PPF .............................................................................................. 6-6 Building 40D Office Areas ........................................................................................ 6-6 Building 44, HPF ....................................................................................................... 6-6 6.3 Site 254 Payload Processing Facility .................................................................... 6-10 Overview.................................................................................................................. 6-10 Site 254 PPF Layout ................................................................................................ 6-10 Site 254 PPF Features .............................................................................................. 6-11 6.4 Zenit Technical Complex Site 42 .......................................................................... 6-12 Overview.................................................................................................................. 6-12 Integration Area Layout/Features ............................................................................ 6-13 Spacecraft Equipment Room ................................................................................... 6-13 Customer Office Areas ............................................................................................ 6-14 6.5 Zenit Launch Complex (LC) – Site 45 ................................................................. 6-14 Overview.................................................................................................................. 6-14 Launch Complex Automated Systems..................................................................... 6-14 Customer EGSE Room (Bunker)............................................................................. 6-16 Command Center ..................................................................................................... 6-17 6.6 Cosmodrome Amenities......................................................................................... 6-18 Visa and Access Authorization................................................................................ 6-18 Customs Clearances ................................................................................................. 6-18 Transportation .......................................................................................................... 6-18 Consumables ............................................................................................................ 6-18 Security .................................................................................................................... 6-18 Schedules ................................................................................................................. 6-18 External Communications........................................................................................ 6-19 Medical Care............................................................................................................ 6-19 Accommodations and Dining .................................................................................. 6-19


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7. SPACECRAFT DESIGN & VERIFICATION REQUIREMENTS ............................ 7-1 Overview............................................................................................................................. 7-1 7.1 Additional Spacecraft Design Consideration ........................................................ 7-1 7.1.1 Constraints on Spacecraft Transmitting and Receiving......................... 7-1 7.1.2 Horizontal Handling................................................................................... 7-2 7.1.3 Safety Design Considerations .................................................................... 7-3 Pressurized Systems ..................................................................................... 7-3 Ordnance Systems ........................................................................................ 7-3 7.1.4 Ground Support Equipment (GSE) Considerations............................... 7-3 7.2 Spacecraft Structural Capability ........................................................................... 7-4 Flexibility................................................................................................................... 7-4

7.2.1 Spacecraft Structural Capability.............................................................. 7-4 Factors of Safety........................................................................................................ 7-4

Test Verified Model Required for Final CLA ........................................................... 7-4 Test Requirements ..................................................................................................... 7-5 Modal Survey Test..................................................................................................... 7-5 Static Loads Test........................................................................................................ 7-5 Sine Vibration Testing............................................................................................... 7-6 Acoustic Testing ........................................................................................................ 7-7 Shock Qualification ................................................................................................... 7-7 7.2.2 Matchmate Test .......................................................................................... 7-8

8. MISSION INTEGRATION AND OPERATIONS ........................................................ 8-1 Overview............................................................................................................................. 8-1

8.1 Mission Management............................................................................................... 8-1 Mission Manager ....................................................................................................... 8-1 Mission Team Roles and Responsibilities ................................................................. 8-2 8.2 Mission Documentation and Schedule ................................................................... 8-2 Overview.................................................................................................................... 8-2 Integration Documentation ........................................................................................ 8-2 Spacecraft/Land Launch System Interface Control Document ................................. 8-2 ICD Verification Matrix ............................................................................................ 8-2 Mission Integration Schedule .................................................................................... 8-3 8.3 Mission Analyses ...................................................................................................... 8-4 Mission Analyses ....................................................................................................... 8-5 8.4 Operations Planning ................................................................................................ 8-6 Launch Campaign Planning....................................................................................... 8-6


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8.5 Launch Campaign.................................................................................................... 8-7 Overview.................................................................................................................... 8-7 Spacecraft Arrival and Transport............................................................................... 8-7 Spacecraft Processing ................................................................................................ 8-7 Spacecraft Fueling ..................................................................................................... 8-7 Spacecraft Mating with Launch Vehicle Elements in PPF Integration Bay.............. 8-8 Launch Vehicle Autonomous Processing .................................................................. 8-8 Mating with Zenit Stages ........................................................................................... 8-9 Integrated Testing ...................................................................................................... 8-9 Transfer Readiness Review and Transfer to Launch Pad .......................................... 8-9 Launch Pad Operations .............................................................................................. 8-9 Launch Readiness Review ....................................................................................... 8-10 Propellant Loading................................................................................................... 8-10 Second “Go” Poll and Launch ................................................................................. 8-10 Launch Control Center............................................................................................. 8-10 8.6 Post-Flight Activities.............................................................................................. 8-11 8.7 Safety....................................................................................................................... 8-11 8.8 Quality Assurance.................................................................................................. 8-11 General..................................................................................................................... 8-12 Hardware Review..................................................................................................... 8-12

APPENDIX A USER QUESTIONNAIRE ............................................................................. A-1 Spacecraft Physical Characteristics .................................................................................. A-2 Spacecraft Orbit Parameters ............................................................................................. A-3 Guidance Parameters ........................................................................................................ A-3 Electrical Interface ............................................................................................................ A-4 Thermal Environment ....................................................................................................... A-8 Dynamic Environment ...................................................................................................... A-9 Ground Processing Requirements ................................................................................... A-10 Contamination Control Requirements ............................................................................. A-14

APPENDIX B SEA LAUNCH STANDARD OFFERINGS AND OPTIONS......................B-1

Standard-Offering Hardware ..............................................................................................B-1 Launch Vehicle ..........................................................................................................B-1 Payload Fairing and Spacecraft Adapter....................................................................B-1 Electrical interfaces....................................................................................................B-2


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Standard-Offering Launch Vehicle Performance ...............................................................B-2 Orbit and Mass ...........................................................................................................B-2 Orbit Accuracy...........................................................................................................B-2

Standard-Offering Launch Services....................................................................................B-2 Mission Management.................................................................................................B-2 Meetings and Reviews ...............................................................................................B-3 Documentation ...........................................................................................................B-3 Mission Integration ....................................................................................................B-4 Interface Test..............................................................................................................B-4

Standard-Offering Facilities And Support Services ...........................................................B-5 Payload Processing Facilities.....................................................................................B-5 PPF Communication ..................................................................................................B-6 PPF Security...............................................................................................................B-6 PPF Support Services.................................................................................................B-6 Launch Vehicle Integration Facility, Area 42............................................................B-7 Launch Complex (LC) Facilities, Area 45.................................................................B-7 Launch Complex Communications............................................................................B-7 Launch Complex Area 45, Security...........................................................................B-7 Environmental Controls .............................................................................................B-8 Range Services...........................................................................................................B-8 Logistics Support .......................................................................................................B-8

Optional Services................................................................................................................B-8 Mission Analysis........................................................................................................B-8 Interface Tests ............................................................................................................B-9 Support Services ........................................................................................................B-9 Facilities .....................................................................................................................B-9 Materials.....................................................................................................................B-9


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LIST OF FIGURES

Page 1-1 Sea Launch Zenit-3SL ...................................................................................... 1-2 1-2 Mission Integration Timelines .......................................................................... 1-4 1-3 The Baikonur Cosmodrome.............................................................................. 1-6 1-4 Land Launch Organizational Structure............................................................. 1-7 1-5 The Zenit Vehicle(Yuzhnoye) ........................................................................ 1-10 1-6 Zenit Complex at Baikonur (KBTM) ............................................................. 1-11 1-7 The Block DM (Energia) ................................................................................ 1-13 2-1 The Zenit-3SLB ................................................................................................ 2-1 2-2 The Zenit-2SLB ................................................................................................ 2-1 2-3 Cosmonaut Access Tower at the Zenit Launch Complex................................. 2-4 2-4 Zenit Stage 1 and Stage 2 Configuration .......................................................... 2-7 2-5 Land Launch Zenit Stage 1 ............................................................................... 2-8 2-6 RD-171M Engine.............................................................................................. 2-9 2-7 Zenit Second Stage ......................................................................................... 2-10 2-8 Block DM-SLB............................................................................................... 2-14 2-9 Zenit-3SLB Ascent Unit ................................................................................. 2-15 2-10 Zenit-2SLB Payload Unit ............................................................................... 2-10 2-11 Zenit 2 Spacecraft Adapter ............................................................................. 2-19 3-1 Flight Corridors for Land Launch from Baikonur ............................................ 3-4 3-2 Three-burn Block DM Mission Profile to GTO ............................................... 3-6 3-3 Approved Land Launch Ground Track and Drop Zones for GTO ................... 3-7 3-4 Injection Ground Track for Generic Zenit-3SLB GTO Mission .................... 3-11 3-5 Typical Ascent Profile to ISS Orbit ................................................................ 3-14 3-6 Flight Ground Track for Zenit-2SLB Mission to ISS Orbit ........................... 3-15 3-7 Zenit-3SLB Payload Capability to GTO......................................................... 3-16 3-8 Zenit-3SLB Performance to Circular Orbits................................................... 3-17 3-9 Zenit-3SLB Performance to Elliptical Orbits ................................................. 3-18 3-10 Zenit-3SLB High-Energy and Earth Escape Payload Capability ................... 3-19 3-11 Zenit-2SLB Payload Capability for Circular Low Earth Orbits ..................... 3-20 3-12 Zenit-2SLB Performance to Elliptical Orbits ................................................. 3-21 4-1 Reference Coordinate System to Define Spacecraft Environments ................. 4-2 4-2 Typical Quasi-Static (Max Expected) Design Loads in Flight......................... 4-4 4-3 Random Vibration Environment During Flight................................................ 4-7 4-4 Max Expected Acoustic Pressure Envelope inside Zenit-2SLB Fairing .......... 4-9 4-5 Max Expected Acoustic Pressure Envelope inside Zenit-3SLB Fairing .......... 4-9 4-6a Zenit-3SLB Spacecraft SRS with Standard SAAB 937-mm Adapter ............ 4-11 4-6b Zenit-3SLB Spacecraft SRS with Standard SAAB 1194-mm Adapter .......... 4-11 4-6c Zenit-3SLB Spacecraft SRS with Standard SAAB 1666-mm Adapter .......... 4-12 4-6d Zenit-2SLB Spacecraft SRS with Standard SAAB 2624-mm Adapter .......... 4-13 4-7a Ambient Electromagnetic Environment within PPF Site 254 ........................ 4-15 4-7b Ambient Electromagnetic Environment within PPF Area 31......................... 4-15 4-8 Ambient Electromagnetic Environment within ILV Assembly Bldg............. 4-16 4-9 Ambient Electromagnetic Environment at Zenit Launch Complex ............... 4-16


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4-10 Max Field Intensity Levels at SC Separation Plane, Zenit-2SLB .................. 4-19 4-11 Max Field Intensity Levels at SC Separation Plane, Zenit-3SLB .................. 4-19 4-12 Zenit-2SLB Ascent Unit Air-Conditioning and Venting Scheme .................. 4-23 4-13 Zenit-3SLB Ascent Unit Air-Conditioning and Venting Scheme .................. 4-23 4-14 Zenit-3SLB Free Molecular Heating Environment ........................................ 4-25 4-15 Zenit-2SLB Ascent Venting Scheme.............................................................. 4-27 4-16 Zenit-3SLB Ascent Venting Scheme.............................................................. 4-27 4-17 Typical Zenit-2SLB Fairing Internal Pressure Profile During Ascent ........... 4-28 4-18 Typical Zenit-3SLB Fairing Internal Pressure Profile During Ascent ........... 4-28 4-19 Typical Zenit-2SLB Fairing Internal Pressure Profile Decay Profile............. 4-29 4-20 Typical Zenit-3SLB Fairing Internal Pressure Profile Decay Profile............. 4-29 5-1 Fairing for Zenit-3SLB ..................................................................................... 5-3 5-2 Fairing for Zenit-2SLB ..................................................................................... 5-4 5-3 General Lay-out of the 4.1-meter 17S72 Fairing (Zenit-3SLB) ....................... 5-5 5-4 General Lay-out of the 3.9-meter Fairing (Zenit-2SLB) .................................. 5-6 5-5 Spacecraft Static Envelope within Zenit-2SLB Fairing ................................... 5-7 5-6 Spacecraft Static Envelope within Zenit-3SLB Fairing.................................... 5-8 5-7 Locations for Access Doors, Zenit-3SLB Payload Fairing ............................ 5-10 5-8 Locations for Access Doors, Zenit-2SLB Payload Fairing ............................ 5-11 6-1 Location of Land Launch Facilities at Baikonur .............................................. 6-1 6-2 Krainy Airport .................................................................................................. 6-2 6-3 Spacecraft Off-Load at Yubileiny Airport ........................................................ 6-3 6-4 Ascent Unit Transportation with Thermostating Car........................................ 6-4 6-5 Area 31 Partial Facility Lay-out ....................................................................... 6-5 6-6 Lay-Out of Buildings 40 and 40D .................................................................... 6-7 6-7 SC Processing and Joint Operations Area in Buildings 40 and 40D ................ 6-8 6-8 Hazardous Processing Facility, Building 44, at Site 31.................................... 6-9 6-9 Lay-out of SC PPF at Site 254 with Proposed Adjacent Building ................. 6-10 6-10 Encapsulation Operations in Site 254 Room 102 ........................................... 6-11 6-11 North Rail at the Zenit Technical Complex, Site 42....................................... 6-12 6-12 Clean Room at Area 42................................................................................... 6-13 6-13 Lay-out of the Zenit Launch Complex, Area 45............................................. 6-15 6-14 Location of Room 114 (Customer EGSE Room) ........................................... 6-16 6-15 Customer Location Options in the Launch Command Center........................ 6-17 6-16 Hotel 1 at Site 2Zh Near the Site 254 PPF...................................................... 6-19 7-1 Maximum Intentional Spacecraft Electric Field Impingement on Launch

Vehicle .............................................................................................................. 7-2 7-2 Electrical and Mechanical Matchmate Test ...................................................... 7-8


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LIST OF TABLES

Page 2-1 Sea Launch Stages, Cumulative Flight History, All Related Configurations... 2-3 2-2 Reliability and Flight History, Sea Launch Configuration Only ...................... 2-3 2-3 Land Launch Zenit Specifications .................................................................... 2-6 2-4 Block DM-SLB Specifications ....................................................................... 2-13 3-1 Launch Operational Features ............................................................................ 3-3 3-2 Zenit Launch Azimuths and Inclinations from Baikonur ................................. 3-4 3-3 Accessible Orbits on Land Launch ................................................................... 3-5 3-4 Flight Timeline – GTO by Zenit-3SLB .......................................................... 3-10 3-5 Flight Timeline – Zenit-2SLB ISS Mission.................................................... 3-13 3-6 Zenit-3SLB Payload Capability to GTO......................................................... 3-16 3-7 Zenit-3SLB Performance to Circular Orbits................................................... 3-17 3-8 Zenit-3SLB Performance to Elliptical Orbits ................................................. 3-18 3-9 Zenit-3SLB High-Energy and Earth Escape Payload Capability ................... 3-19 3-10 Zenit-2SLB Payload Capability for Circular Low Earth Orbits ..................... 3-20 3-11 Zenit-2SLB Performance to Elliptical Orbits ................................................. 3-21 3-12 Zenit-2SLB and Zenit-3SLB Orbital Insertion Accuracy............................... 3-23 3-13 Zenit-3SLB Direct GEO Insertion Accuracy.................................................. 3-23 3-14 Spacecraft Motion after Separation – Single Payload, Zenit-2SLB ............... 3-25 3-15 Spacecraft Motion after Separation – Multiple Payloads, Zenit-2SLB .......... 3-25 4-1 Maximum Quasi-Static Accelerations During Ground Operations .................. 4-3 4-2 Sinusoidal Vibrations at Spacecraft Interface................................................... 4-5 4-3 Random Vibration during Ground Transport, not in Spacecraft Container...... 4-6 4-4 Random Vibration Environment During Flight................................................ 4-6 4-5 Maximum Expected Acoustic Pressure Envelope Inside Fairings ................... 4-8 4-6 Zenit-3SLB Spacecraft SRS with Standard SAAB Adapters ......................... 4-10 4-7 Zenit-2SLB Spacecraft Shock Response Spectra (SRS) ................................ 4-13 4-8 Characteristics of the Sirius Transmitters (Zenit-2SLB and Zenit-3SLB) ..... 4-17 4-9 Characteristics of BITC-B Telemetry Equipment (Zenit-3SLB Only)........... 4-17 4-10 Characteristics of Glonass Receiver (Zenit-2SLB and Zenit-3SLB).............. 4-18 4-11 Max Field Intensity Levels Generated by Launch Vehicle at SC Separation

Plane, Without Fairing Attenuation ................................................................ 4-18 4-12 Spacecraft Ground Thermal and Humidity Environment............................... 4-22 4-13 Flight Thermal Environments ......................................................................... 4-24 4-14 Fairing Internal Surface Cleanliness Levels at Encapsulation........................ 4-30 5-1 Expected Spacecraft Mass and CG Limits – Zenit-3SLB ................................ 5-1 5-2 Expected Spacecraft Mass and CG Limits – Zenit-2SLB ................................ 5-2 5-3 Recommended Spacecraft Fundamental Frequencies ...................................... 5-2 5-4 Standard SAAB Ericsson Space Spacecraft Adapters .................................... 5-12 5-5 Spacecraft Umbilical Links............................................................................. 5-14 5-6 Umbilical Hook-up Locations and Availability.............................................. 5-14 5-7 Characteristics of Commands from Launch Vehicle to Spacecraft ................ 5-16 5-8 Electrical Power Supplies for Customer EGSE.............................................. 5-17 5-9 Uninterruptible Power Supply for Customer EGSE ....................................... 5-17


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7-1 Factors of Safety and Test Options................................................................... 7-4 7-2 Sine Vibration Amplitudes and Sweep Rates ................................................... 7-6 7-3 Spacecraft Acoustic Margins and Test Durations............................................. 7-7 8-1 Typical Mission Integration Schedule .............................................................. 8-3 8-2 Typical Launch Campaign Schedule ................................................................ 8-6


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Abbreviations and Acronyms

A ampere(s) A0 azimuth A/C air conditioning ASTM American Society for Testing and Manufacture ATB Assembly & Test Building AU ascent unit B Baikonur BER bit error rate BLS Boeing Launch Services BPS bits per second C Celsius or Centigrade C3 velocity squared at infinity CA California CC command center 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 CM centimeter(s) CVCM collected volatile condensable material DB decibel(s) DC direct current DP dew point EGSE electrical ground support equipment EMC electromagnetic compatibility ESD electro-static discharge F Fahrenheit FM frequency modulation FMH free molecular heating FSA Federal Space Agency FT foot/feet G gravity GEO geosynchronous or geostationary Earth orbit


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GOWG ground operations working group GTO geosynchronous transfer orbit GSE ground support equipment H; HR hour HEO high Earth orbit HPF hazardous processing facility HZ hertz I inclination

moment of inertia ICAO International Civil Aviation Organization ICD interface control document I/F interface ILV integrated launch vehicle IRD interface requirements document ISS International Space Station K thousand(s) KBTM Design Bureau of Transport Machinery KG kilogram(s) KGF Kilogram(s) force KM kilometer(s) KN kilonewton(s) KVA kilo volt-ampere(s) KW kilowatt(s) LB(S) pound(s) LBF pound(s) force LEO low Earth orbit LC launch complex LLC limited liability company LOX liquid oxygen LRR launch readiness review LSA launch services agreement LV launch vehicle M meter(s) MS millisecond(s) ME main engine MEO medium Earth orbit MHZ megahertz


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MM millimeter(s) N/A not applicable PA pascal(s) PCM pulse control modulation PDR preliminary design review PLF payload fairing PLU payload unit PPF payload processing facility PPM parts per million PSI pounds per square inch PSM payload systems mass PSS payload support structure R&D research and development RF radio frequency RH relative humidity S second(s) S/C; SC spacecraft SCA spacecraft adapter SCAPE self-contained atmospheric protection ensemble SIS Space International Services, Ltd SOW statement of work SPST solid propellant separation thrusters SRS shock response spectra T Time

tonne(s) TBD to be determined TBR to be revised or reviewed TM telemetry TML total mass loss TRR transfer readiness review TsENKI Center for Ground Space Infrastructure Operations UPS uninterruptible power supply USA United States of America USSR Union of Soviet Socialist Republics V volt(s)

velocity VAC volt(s) of alternating current


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W watt(s) 0 degrees σ sigma (one standard deviation) µ micro (10–6)

Land Launch User’s Guide Section 1

Rev. Initial Release HPD-19000 1-1

1. INTRODUCTION Purpose The purpose of the Land Launch User’s Guide is to familiarize members of the cus-

tomer community with the Land Launch system and associated services. This docu-ment is the starting point for understanding the Land Launch spacecraft integration process, Land Launch interfaces and the overall capabilities of the system. Land Launch services are provided by the Sea Launch Company, LLC, acting in cooperation with Space International Services, Ltd (SIS), and are marketed by Boeing Launch Ser-vices (BLS). Further information may be obtained by contacting BLS directly (www.boeing.com/launch).

1.1 Overview of the Land Launch System

Overview As its name implies, Land Launch is Sea Launch on land: the proven hardware, proc-

esses and people of Sea Launch shifted to a land-based launch site at the Baikonur Cosmodrome. Land Launch and Sea Launch complement each other by addressing dif-ferent classes of payloads. Whereas Sea Launch is a heavy-lift launch system that de-livers more than 6,000 kilograms to a geosynchronous transfer orbit (GTO) requiring less than 1500 meters/second to geostationary orbit (GEO), Land Launch is a medium-lift launch system that delivers 3,600 kg to an equivalent GTO. The difference in GTO performance is due to the change in launch site.

There are two Land Launch configurations:

• The Zenit-3SLB (“B” for Baikonur), a three-stage integrated launch vehicle (ILV) closely derived from the Sea Launch Zenit-3SL (Figure 1-1), is suited for delivering payloads to medium and high, circular and elliptical Earth orbits, in-cluding GTO and GEO, as well as escape trajectories.

• The Zenit-2SLB, a two-stage ILV based on the first two stages of the Sea Launch Zenit-3SL, is designed for delivering payloads to inclined low Earth circular and elliptical orbits.

This edition of the Land Launch User’s Guide addresses a representative set of missions and launch services that can be cost-effectively accomplished with these two companion launch systems. Potential users are invited to complete the User Questionnaire (Appendix A) and return it to the address indicated therein.


1-2 HPD-19000 Rev. Initial Release

Figure 1-1. The Proven Sea Launch System Provides a Solid Foundation for Land Launch



What the sys-tem includes

The standard Land Launch system comprises:

• The three-stage Zenit-3SLB launch vehicle

• The two-stage Zenit-2SLB launch vehicle

• The Zenit launch complex at Baikonur cosmodrome, and approved downrange stage and fairing impact zones

• Support facilities at Baikonur cosmodrome for ground processing of the spacecraft and launch vehicle including fueling, check-out, ILV assem-bly and launch

• Transportation equipment including rolling stock and thermostating sys-tems for moving people and hardware between locations at Baikonur cosmodrome

• Tracking, meteorological and communications assets

A thorough summary of Land Launch interfaces, operations, services and facilities is provided herein as Appendix B: Land Launch Standard Offerings and Options.

Advantages to the Customer

Land Launch provides:

• The most mature flight hardware in its payload class, closely derived from the proven Sea Launch configuration. Maturity is greatest with the Land Launch upper stage, the venerable Block DM, which is the most experienced and reliable upper stage in any payload class with continu-ous service since 1974 on more than 220 missions with a demonstrated reliability exceeding 97%

• The versatility of the Block DM, which has the capability for multiple re-starts, long duration missions, roll and coast maneuvers, accurate orbital insertion and tightly controlled SC separation parameters of motion and attitude

• Dedicated launch, eliminating schedule and technical risk associated with co-passengers

• Existing, operational, proven ground facilities

• The teamwork and proven expertise of the Land Launch partners, which include the same core companies that work together on Sea Launch

• The responsiveness of the world’s only dedicated commercial launch services family: Sea Launch and Land Launch



Timeline It is expected that the first experience of a new spacecraft type can be integrated by eighteen months after contract signature. Repeat missions of the same spacecraft type can be integrated by twelve months after contract signature. Figure 1-2 provides a top-level summary of the relative mission phases, while more extensive details on the mis-sion flow process are presented in Section 8.

Land Launch has the capability for conducting successive launches on 30-day centers at Baikonur. The vehicles are assembled off the launch complex, and the nominal time between vehicle roll-out and launch is on the order of twenty-eight hours, provided that spacecraft check-out on the pad does not exceed eighteen hours.

Land Launch is dedicated to reducing the time required for integrating and launching spacecraft. Deviations from the standard flows may be accommodated on a case-by-case basis, particularly for two-stage missions.

18 Month Integration Timeline

Sign Contract

Mission Analysis

Operations Planning

Launch Ops S/C Delivery and processing

18 mo 12 mo 6 mo 0 mo

New Mission

12 Month Integration Timeline

Sign Contract

Mission Analysis

Operations Planning

Launch Ops S/C Delivery and processing

12 mo 6 mo 0 mo

Repeat Mission

Figure 1 – 2. Land Launch Mission Integration Timelines Take Advantage of In-Place Sea Launch Procedures and Processes



Baikonur Cosmodrome

The Zenit Launch Complex is located at the Baikonur Cosmodrome in Kazakhstan at 63°E, 46°N (Figure 1-3). Baikonur was a primary launch site for the Soviet Union and, in the post-Soviet era, it continues to be the principal launch site for both the Russian and Ukrainian space industries.

Many of the greatest events of the Space Age have occurred at Baikonur including launch of the world’s first satellite in 1957, the first mission (unmanned) to the moon in 1959 and the first manned launch in 1961. Thousands of other launches have taken place at Baikonur over the ensu-ing decades up to the present day. In recent years Baikonur has become an important commercial spaceport. Since 1995 there have been more than fifty launches of satellites made outside the Commonwealth of In-dependent States (CIS).

The cosmodrome is linked to major cities in the CIS by air, road and railway transport. The local area of the cosmodrome also has a developed road and railway network. The closest residential area is the city of Baikonur, located just south of the cosmodrome approximately 60 km from the Zenit Complex. Baikonur city lies on the north bank of the Syrdarya river. The Tyuratam settlement and the railway station of the same name (Kazakh railway) adjoin Baikonur. The Baikonur airport (Krainy) is linked to Moscow by regular and charter passenger flights and can accommodate most cargo and passenger airplanes. Spacecraft and cargo typically arrive via Yubileiny airport located within the cos-modrome itself.

The Zenit space rocket complex is the newest operational launch facility at Baikonur, and conducted its first launch in 1985. It is located far from large populated areas, ensuring safety of launches and allowing for easy allocation of impact zones. Furthermore, its position on the eastern side of the cosmodrome enables greater flexibility with respect to launch azi-muths than other local launch complexes can offer.

The specific Baikonur cosmodrome facilities utilized by Land Launch are described more extensively in Section 6.



Figure 1-3. The Baikonur Cosmodrome is Readily Accessible from Moscow



1.2 Land Launch Organization Overview Land Launch contracts will be managed by the existing Sea Launch or-

ganization in Long Beach, California. Such co-location and shared use of resources and personnel is key to enabling Land Launch to provide a “western” interface to its customers that is comparable to that experi-enced with Sea Launch. Launch services out of Baikonur are obtained via subcontract from Sea Launch to Space International Services, Ltd (SIS). SIS is a limited liability company based in Moscow consisting of key Land Launch members from Ukraine and Russia, all of which also participate in Sea Launch missions. The Land Launch organizational structure is presented in Figure 1- 4.

Figure 1-4. The Land Launch Program Brings Together an Experienced Team

Boeing

Space International Services, Ltd.

(SIS)

Yuzhnoye SDO Yuzhmash PO KBTM TsENKI RSC Energia NPO Lavochkin

Federal Space Agency(FSA)

- CIS Licensing

- CIS Governmental Interfaces

- Launch Services

- Mission Integration

- Baikonur Operations

- Sales & Marketing

- Mission management

- Quality and Technical Oversight

- Hardware Acceptance Review

- Customer Safety and Logistics Support

- Program Management

- Contracts and Legal Support

- Insurance Interface

- US Licensing

Sea Launch Company, LLC




Sea Launch Company, LLC, headquartered in Long Beach, California, is the world leader in commercial heavy-lift launch services with its highly successful and innovative ocean-based launch system. Sea Launch is a partnership comprised of Boeing, RSC Energia, SDO Yuzhnoye, the Kvaerner Group and PO Yuzhmash. Additional information related spe-cifically to Sea Launch can be found in the Sea Launch User’s Guide and on the corporate website at: www.sea-launch.com.

For the Land Launch service, Sea Launch responsibilities include:

• Management of the overall endeavor

• Contracts & legal management

• Insurance and financing interfaces

• US licensing and US governmental interfaces

The Boeing Company

The Boeing Company is enlisted by Sea Launch to provide the following capabilities on Land Launch missions:

• Mission management

• Payload integration support

• Hardware quality reviews

• Overall technical and quality oversight

• Satellite safety assessments and customer logistics support

• Marketing & sales

Space International Services, Ltd

Space International Services, Ltd (SIS) is a company comprised of SDO Yuzhnoye, PO Yuzhmash, RSC Energia, the Center for Ground Space Infrastructure Operations (TsENKI) and the Design Bureau of Transport Machinery (KBTM). Its office is in Moscow, Russia. SIS direct respon-sibilities include:

• Launch services and Baikonur operations

• Supplier management (all CIS and launch hardware suppliers)

• CIS licensing and regulation

• Third party insurance



SDO Yuzhnoye SDO Yuzhnoye is the leading Ukrainian aerospace organization with vast experience in the design and development of launch vehicle tech-nology. The company is established under the laws of Ukraine, with its principal place of business in Dnepropetrovsk, Ukraine. The SDO Yuzhnoye team along with that of PO Yuzhmash has conducted hun-dreds of successful launches from Baikonur (Fig. 1-5). Additional infor-mation related to SDO Yuzhnoye can be found on the website at: www.yuzhnoye.dp.ua.

SDO Yuzhnoye performs the following work in support of the Land Launch program:

• Design and configuration management of the Zenit stages for both the Zenit-2SLB and Zenit-3SLB, as well as design support during their manufacturing

• Design of the Zenit-2SLB fairing and integrated launch vehicle (ILV) as a whole

• Systems engineering and integration

• Payload integration and mission analysis

• Technical management and participation in ILV processing and launch operations



Figure 1-5. The Zenit Vehicle Reflects Five Decades of SDO Yuzhnoye and PO Yuzhmash Experience in Designing, Building and Operating Launch Vehicles

PO Yuzhmash PO Yuzhmash is another leading Ukrainian aerospace enterprise having

vast experience in the development and production of major launch vehi-cles. The company is incorporated under the laws of Ukraine and, like SDO Yuzhnoye, its principal place of business is in Dnepropetrovsk, Ukraine.

PO Yuzhmash performs the following work in support of the Land Launch program:

• Manufacturing of the first two stages, and the Zenit-2SLB fairing

• ILV integration

• Participation in ILV processing and launch operations



Design Bureau of Transport Machinery (KBTM)

The principal offices of KBTM are located in Moscow, Russia. In its fifty year history KBTM has designed numerous launch complexes in-cluding the Zenit complex at Baikonur (Figure 1-6), and also performs vehicle integration and launch operations. More than 900 orbital launches have been conducted from launch complexes built by KBTM. KBTM is a major subcontractor on the Sea Launch program responsible for ground support equipment maintenance including the trans-porter/erector. On Land Launch, KBTM will have overall responsibility for:

• ILV assembly area and launch complex

• ILV processing for launch

• Ground support equipment

• Launch operations

Figure 1-6. KBTM Developed the ILV Transporter and Erector Equipment for both the

Sea Launch and Land Launch Programs



Center for Ground Space Infrastruc-ture Operations (TsENKI)

The principal offices of TsENKI are located in Moscow, Russia in close association with the Russian Space Agency. TsENKI is responsible for operation of the ground aerospace infrastructure facilities at Baikonur cosmodrome and downrange sites. TsENKI will have fundamental re-sponsibility on Land Launch for:

• Russian launch licenses and third party insurance

• Recording, acquisition and processing of telemetry data

• Security and guard services

• Telecommunication services and communication systems

• Logistics support (propellant components and compressed gases)

• Securing cosmodrome support services during ILV processing and launch

• Processing operation and maintenance of impact zones

• Coordinating electromagnetic compatibility for ILV processing and launch operations



RSC Energia RSC Energia is the premier Russian space company. RSC Energia, de-veloper of launch vehicles and propulsion systems, spacecraft, space sta-tions, as well as manned and cargo modules, brings its legendary experi-ence in space exploration and launch system integration to Land Launch. Energia is a joint stock company established under the laws of the Rus-sian Federation, with its principal place of business in Korolev (near Moscow), Russia. Additional information related to RSC Energia can be obtained at: www.energia.ru. Energia has the responsibility on Land Launch for:

• Design and manufacture of the Block DM-SLB upper stage (Figure 1-7)

• Integration of the Ascent Unit comprising the Block DM-SLB, fairing, adapter and spacecraft

• Mission analysis support

• Launch operations support

• Customer support

Figure 1-7. Energia’s Block DM is the Most Successful and Most Proven Upper Stage Available Today



NPO Lavochkin NPO Lavochkin is located in Khimki, Russia and has a distinguished his-tory of achievement in the design, development and manufacture of air-craft, launch vehicle upper stages and spacecraft including many deep space missions to the moon, Venus and Mars. Lavochkin has also pro-vided more than 100 fairings for various launch vehicles and will be pro-viding Land Launch with a flight-proven fairing for the Zenit-3SLB.

Federal Space Agency

Land Launch enjoys the support of the Federal Space Agency which will be providing Land Launch with the use of its facilities at Baikonur, launch licensing and associated regulatory support including relations with other CIS governments on whose territory Land Launch activities will be conducted.



2. VEHICLE DESCRIPTION Overview Land Launch uses either two or three in-line, liquid oxygen (LOX) and kero-

sene stages. The three-stage Zenit-3SLB configuration (Figure 2-1) is used for medium-lift missions to medium and high, circular or elliptical orbits includ-ing GTO and GEO, as well as escape trajectories. The two-stage Zenit-2SLB configuration (Figure 2-2) is used for missions to low earth circular and ellip-tical orbits. Each configuration uses a different fairing. All elements of either configuration have extensive flight heritage.

The principal components of the Land Launch vehicles are:

• Zenit Stage 1

• Zenit Stage 2

• Block DM-SLB upper stage (Zenit-3SLB configuration)

• Fairing and Payload Support Structure

Figure 2-1. The Zenit-3SLB

Figure 2-2. The Zenit-2SLB

58.65 m (192.4 ft)

Zenit Stage 132.9 m (107.9 ft)


BlockDM-SLB

Fairing10.4 m (34.1 ft)

Ø3.7 m (12.1 ft)Ø4.1 m (13.5 ft) Ø3.9 m (12.8 ft)



Fairing13.7 m (44.8 ft)

Ø3.9 m (12.8 ft)

Intersection Bay0.35 m (1.1 ft)

57.4 m (188.3 ft)



Design The Zenit first and second stages used on Land Launch are interchange-able with the Sea Launch first and second stages. They are manufactured by PO Yuzhmash in Ukraine, with design oversight provided by SDO Yuzhnoye. The Block DM-SLB third stage, used only on the Zenit-3SLB, is closely adapted from the Block DM-SL used by the Sea Launch program (the differences are described in section 2.2) and is manufac-tured by RSC Energia in Russia. The fairing for the Zenit-3SLB is 4.1 meters in diameter and is manufac-tured by NPO Lavochkin in Russia. It was designed specifically for the Block DM and has an unblemished flight history dating to 1996. The Zenit-2SLB fairing is 3.9 meters in diameter and is manufactured by PO Yuzhmash. It was designed specifically for the two-stage Zenit configu-ration and has a flight history dating to 1985.

The payload support structure for the Zenit-3SLB is provided by RSC Energia. It consists of a spacecraft adapter (SCA) typically procured from Saab Ericsson Space (937, 1194 or 1666 interfaces) and a transfer compartment manufactured by Energia. The payload support structure for the Zenit-2SLB is provided by SDO Yuzhnoye and will typically consist of a Saab SCA mounted on a truss manufactured by PO Yuzhmash. Unique interfaces and multi-satellite dispensers can also be provided if required.

Zenit Flight History

The original Zenit-2 was first launched in 1985 from Baikonur Cos-modrome. As of March 2004, it has completed 30 successful missions in 35 launch attempts. The Zenit first-stage booster also served as the strap-on for the Energia launch vehicle (four per launch) and logged an addi-tional eight successes in two flights in this capacity. The modified and improved Zenit-2S, the version used on Sea Launch, has flown twelve times as of February 2004. Land Launch also uses the Zenit-2S.

Block DM Flight History

From its introduction in 1974 through March 2004 the Block DM has completed 222 successful missions in 228 attempts in various versions, including eleven successes in eleven attempts for the Block DM-SL ver-sion used on Sea Launch, making it far and away the most proven, reli-able and mature upper stage in the launch industry. Past missions have included GTO and direct insertion GEO for commercial and for govern-ment customers, high elliptical orbits, low and high circular orbits, dedi-cated launches and launches of multiple satellites, and escape trajectories (to Halley’s Comet, Venus and Mars).



Flight Success Ratios

Tables 2-1 and 2-2 list flight records for each of the three Zenit-3SL stages as of March 2004, as well as engineering reliability estimates. The closely-related Zenit-3SLB stages are expected to achieve identical reliability levels.

The engineering reliability estimates account for:

• Extensive testing performed when modifications are made to flight hardware or ground support equipment

• Expected reliability growth, using statistics of other boosters using similar processes and procedures that were also built and launched in the former USSR

• An exhaustive failure analysis team that investigates any flight anomalies and implements measures to ensure that the anomalies never recur

• The Sea Launch mission assurance and audit process currently in place and operating at the factory level in Ukraine and Russia

Table 2-1. Sea Launch Stages, Cumulative Flight History, All Related Configurations

Stage Year Introduced

Versions Flown

Cumulative Flight Record

Zenit Stage 1 1985 3 53 of 55 Zenit Stage 2 1985 2 41 of 44 Block DM 1974 9 222 of 228

Table 2-2. Reliability and Flight History, Sea Launch Configuration Only

Stage Year Introduced Flight Record Reliability

Estimate Zenit Stage 1 1999 12 of 12 Zenit Stage 2* 1999 11 of 11 } 98.0%

Block DM-SL 1999 11 of 11 98.5% * The one Sea Launch failure (mission 3) occurred during second stage opera-tion, but was not caused by the second stage and no design changes to the sec-ond stage resulted from the failure investigation. The failure cause was a fault in ground software that left open a second stage valve.



2.1. The Land Launch Zenit

Land Launch uses the Sea Launch configuration of the Zenit, retaining the improvements and modifications that were made for Sea Launch to the heritage Zenit-2. SDO Yuzhnoye designed the original two-stage Zenit-2 during the late 1970s and early 1980s in response to requirements from the Soviet Ministry of Defense for a launch system that would be able to quickly and efficiently reconstitute military satellite constellations. Conse-quently, the design emphasizes robustness, ease of operation and fast reac-tion times, which are achieved through extensive automation. It incorpo-rates state-of-the-art launch and processing technologies, developed by Land Launch partner KBTM, in contrast to systems developed during pre-vious decades. A second intended use for the original Zenit-2 was manned launches to space station MIR (figure 2-3). Though ultimately it was never used for this purpose due to the break-up of the Soviet Union, in order to be man-rated, the Zenit was designed with a significant degree of internal redundancy and other features to ensure high reliability.

Design Heritage

Figure 2-3. Cosmonaut Access Tower at the Zenit Launch Complex



Changes Made for Sea Launch

Significant configuration differences between the heritage Zenit-2 and the Sea Launch Zenit-2S, which are also retained on the Land Launch Zenit-2SLB, are:

• New navigation system

• Next generation flight computer

• Increased performance due to mass reductions and an increase in sec-ond stage main engine thrust from 87 tonnes force to 93 tonnes force

Avionics Just as on Sea Launch, the Land Launch Zenit contains its own complete complement of avionics for telemetry, guidance and navigation functions even when lifting an upper stage in a three-stage configuration. The on-board Sirius telemetry packages transmit telemetry data on separate RF links to existing ground stations located in Russia and, for sun-synchronous missions, to a remote ground station located on the Arabian Peninsula. For three-stage missions, these Zenit links are complemented by an independent set of data that is provided simultaneously by the Block DM-SLB telemetry system.

Overall Specifications and Configura-tions

Zenit specifications and performance parameters are shown in Table 2-3. Stage 1 and Stage 2 configurations are pictured in Figure 2-4. With pro-pellant mass fractions exceeding 90%, the designs of both stages rank among the most structurally efficient in the world. In the case of the first stage, this is due in large part to the highly efficient RD-171M engine and the lack of strap-on boosters.

The absence of strap-on boosters greatly simplifies pre-launch processing and is a major feature distinguishing Zenit from most other large launch systems. Without strap-ons, the stage structure is more efficient, ordnance count is reduced and overall reliability is enhanced by eliminating expo-sure to the failure of booster separation mechanisms or of the boosters themselves. Furthermore, the streamlined configuration lends itself to ro-bust control margins during all phases of flight which enable the Zenit to fly through a broad range of wind and weather conditions, further ensuring on-time and on-target launch performance.



Table 2-3. Land Launch Zenit Specifications

Stage 1 Stage 2 Zenit

Zenit-2SLB and -3SLB Zenit-2SLB Zenit-3SLB

Burn Time 140 - 150 s 300 - 1,100 s 360 - 370 s

Inert Mass 27,564 kg (60,768 lb) 8,367 kg (18,446 lb) 8,307 kg (18,314 lb)

Fueled Mass 354,350 kg (781,200 lb) 90,854 kg (200,297 lb) 90,794 kg (200,164 lb)

Fuel (kerosene) 90,219 kg (198,897 lb) 23,056 kg (50,829lb)

Oxidizer (LOX) 236,567 kg (521,536 lb) 59,431 kg (131,022 lb)

Length 32.9 m (108 ft) 10.4 m (34 ft)

Diameter 3.9 m (12.8 ft) 3.9 m (12.8 ft)

Engines One RD-171 (four thrust chambers)

One RD-120 Main Engine One RD-8 Vernier Engine (four thrust chambers)

Thrust (sea level) 740,000 kgf (1.63 million lbf) Not applicable

Thrust (vacuum) 806,400 kgf (1.78 million lbf)

Main Engine: 93,000 kgf (205,028 lbf) Vernier Engine: 8,100 kgf (17,857 lbf)

Specific impulse (sea level) 309.5 s Not applicable

Specific impulse (vacuum) 337.2 s Main Engine 350 s

Vernier Engine 342.8 s

Attitude Control Nozzle gimbal + 6.3 deg Vernier engine nozzle gimbal + 33 degrees



Avionics bay

Liquid oxygen tank

Kerosene tank

Main engine

SPST (4) Steering engine

Interstage frame

∅3.9 m (12.8 ft)

Liquid oxygen tank

Kerosene tank

Single turbopumpMain engine nozzles (4)

Solid-propellant separation thrusters (SPST) (4)

View A-А

Separation plane

Stage 210.4 m (34 ft)

Stage 1 32.9 m (108 ft)

АА

Figure 2-4. Land Launch Uses the Same Zenit Stages that are Used

on the Sea Launch Zenit-3SL



2.1.1 Zenit Stage 1 Overall Configuration

The Land Launch Zenit Stage 1 (Figure 2-5) features an aluminum primary structure with integrally machined stiffeners, and environmentally-friendly LOX/kerosene propellants. The upper LOX tank fits in a concave depression at the top of the kerosene tank, and the LOX feed line runs through the mid-dle of the lower tank. With a Zenit-2SLB gross lift-off mass of 450,000 – 460,000 kg, and a Zenit-3SLB gross lift-off mass of 462,000 – 466,000 kg, the 740,000 kgf produced by the first stage yields a very healthy ~1.6 take-off thrust-to-weight ratio for both vehicles. Separation is achieved with four solid retro-rockets located at the base of the stage.

The Land Launch Zenit first stage design is intentionally kept common to that of the Sea Launch stage. Both are manufactured on the same production line at PO Yuzhmash.

Figure 2-5. PO Yuzhmash Achieves Significant Economies of Scale by Manufacturing Both Land Launch and Sea Launch Zenit Stages on the Same Production Line



The RD-171M engine (Figure 2-6), which powers Zenit Stage 1, burns liquid oxygen (LOX) and kerosene. It provides an impressive 740,000 kgf (1.6 million lbf) of thrust at sea level and is one of the most powerful rocket engines in the world, featuring advanced rocket engine technolo-gies developed by leading Russian propulsion organizations. It was de-veloped specifically for the Zenit, in parallel with the closely related RD-170 that served as the strap-on booster for the Energia/Buran. An exhaus-tive test program consuming more than 200 test engines preceded first flight in the mid 1980’s. The four thrust chambers are fed by a single, vertically mounted turbopump, which in turn is powered by two gas gen-erators feeding hot oxidizer-rich gas to a single turbine. Flight control is achieved by gimbaling the independently suspended combustion cham-bers, while the ability to throttle down to ~ 74 % of nominal full-engine thrust provides great flexibility in trajectory design.

RD-171M Engine

265807J3-038 Figure 2-6. The RD-171M is the Most Powerful Liquid Rocket Engine

Presently in Operation



2.1.2 Zenit Stage 2 Overall Configu-ration

Like the first stage, the Zenit second stage (Figure 2-7) features an inte-grally stiffened aluminum construction and environmentally-friendly LOX/kerosene propellants. Propulsion is provided by an RD-120 main engine with steering provided by an RD-8 vernier engine fed from the same propellant tanks. The lower kerosene tank is toroidally shaped and surrounds the main engine, while the upper LOX tank is a domed cylin-der. The stage is topped by an instrument compartment containing the avionics. The Sea Launch and Land Launch Zenit second stages, like the first stages, are manufactured on a common Yuzhmash production line, thereby benefiting from common inventory and Boeing quality oversight processes. The second stage generates 101,000 kg (222,887 lbs) of thrust (RD-120 and RD-8 engines combined). As on the first stage, separation is achieved with four aft-mounted solid retrorockets.

265807J3-037R1 Figure 2-7. The Second Stage’s Toroidally-Shaped Fuel Tank Results in a Shorter, More

Efficient Vehicle Structure



RD-120 Main Engine

The 2nd stage main engine is a single-chamber, fixed nozzle liquid-propellant rocket engine that uses LOX and kerosene to generate 93,000 kgf (205,028 lbf) of thrust. The RD-120 is throttled down to ∼ 78% of nominal full-engine thrust at the end of flight. The RD-120 was devel-oped specifically for the Zenit launch system.

RD-8 Vernier Engine

The RD-8 vernier engine mounted in the aft end of Stage 2 provides three-axis attitude control. The RD-8 uses the same propellants and pro-pellant storage system as the RD-120, with one turbo-pump feeding four gimbaling thrusters spaced around the outside of the RD-120. The RD-8 produces 8,100 kgf (17,900 lbf) of thrust, and was specifically developed for Zenit. The ability to modulate its operation from 65 to 900 seconds following main engine cut-off provides flexibility in mission design for Zenit-2SLB launches to a wide range of circular LEO orbits.



2.2 The Block DM-SLB Upper Stage Overall Configuration

The Block DM-SLB (Figure 2-8) used on the Zenit-3SLB is closely de-rived from the Block DM-SL used on Sea Launch. It is a LOX/kerosene upper stage capable of igniting up to three times during a mission. Basic specifications are provided in Table 2-4.

The basic structure of the Block DM-SLB is provided by the upper adapter together with an internal truss. The middle and lower adapters that enclose the stage are jettisoned before first ignition of the Block DM-SLB. Kerosene is contained in a toroidal tank connected by a truss to the upper adapter which encircles the turbopump of the 11D58M main engine. The spherical LOX tank and the avionics/payload truss are located above the kerosene tank, and also connect to the upper adapter. Two attitude control/ullage engines, which provide stabilization during coast periods, are located on the bottom of the kerosene tank.

11D58M Main Engine

The Block DM-SLB upper stage is powered by the 11D58M engine, which operates on liquid oxygen and kerosene. Its carbon-carbon nozzle is gimbaled to provide pitch and yaw control during powered flight, with turbopump bleed gas used for roll control.

Attitude Control/Ullage Engines

Three–axis stabilization and attitude control during coast periods, includ-ing continuous rolls, are provided by two attitude control/ullage engines using hypergolic propellants that are located on the aft end of the main engine kerosene tank, on either side of the main engine nozzle.

Avionics

The Land Launch Block DM-SLB uses the same avionics as the Sea Launch Block DM-SL, with the exception of differences in the telemetry system more suitable for launches from Baikonur using associated fixed and mobile Russian ground receiving stations.

Changes Made for Sea Launch

Significant configuration differences between the heritage Block DM and the Sea Launch Block DM-SL, which are also retained on the Land Launch Block DM-SLB, are:

• New navigation system • Next generation flight computer • The autonomous control system provided by the R&D and Production Center for Automation and Instruments Manufacturing (NPTs AP) – the premier Russian avionics and space software company • An extended nozzle and various mass reductions for performance im-provement



Block DM-SLB Versus the Block DM-SL

The main differences between the Block DM-SLB and the Block DM-SL are:

• The Block DM-SLB forward structural interfaces are made to be compatible with the Russian-made fairing and payload structure that are used on Land Launch, while the Block DM-SL forward interfaces are compatible with the Boeing-made payload unit hardware that is used on Sea Launch

• The single, large (and heavy) toroidal avionics bay on the Block DM-SL is replaced on the Block DM-SLB with several discrete avion-ics containers for a net reduction in launch mass

• Some sensors and harnesses are removed that are a legacy of early qualification flights and are no longer needed

• A deployable antenna and telemetry system are replaced with a lighter system also used on Zenit that features fixed antennas with two independent radio links

• An uplink command system and its antenna are removed • One set of fuel tanks for the attitude control/ullage engines are re-

moved. Previously, these tanks were routinely under-filled by the equivalent of one set of tanks.

• The LOX tank is pressurized with helium instead of an oxy-gen/helium mixture

• The minimum useable propellant criterion for the final re-start is lowered from 4000-kg to 1500-kg, by adding two 10-kgf thrusters to ensure settling prior to ignition

• An external heat radiator is removed with this function being as-sumed by the upper adapter structure

Table 2-4. Block DM-SLB Specifications

Length 1 5.93 m (19.4 ft)

Diameter (primary) 3.7 m (12.1 ft)

Maximum Launch Mass2, 3 (fueled) 17,800 kg (39,240 lb)

Maximum3 Useable Propellant Reserve 14,580 kg (32,140 lb)

Thrust (vacuum) 8,103 kgf (17,864 lbf)

Note 1: The fairing overlays 1.03 m (3.4 ft) of the length of the Block DM-SLB, as shown in Figure 2-8

Note 2: Includes the lower and middle adapters, which are jettisoned prior to first burn of the Block DM-SLB

Note 3: Fuel is off-loaded for heavier payloads launching east (includes GTO missions), due to drop zone constraints



Avionics Container

Avionics/Payload Truss

Coolant Piping

Upper Adapter

LOX Tank

Middle Adapter

Main Engine

Lower Adapter

Kerosene Tank

Attitude Control/Ullage Engine

Avionics Container

Avionics/Payload Truss

Coolant Piping

Upper Adapter

LOX Tank

Middle Adapter

Main Engine

Lower Adapter

Kerosene Tank

Attitude Control/Ullage Engine

Figure 2-8. Block DM-SLB (dimensions in millimeters)



2.3 Zenit-3SLB Ascent Unit Components and Inte-gration

The Zenit-3SLB Ascent Unit (figure 2-9) consists of the spacecraft, Block DM-SLB, fairing and payload support structure (PSS). These elements are integrated in a Class 100,000 clean environment during ground processing. The PSS is comprised of an industry-standard spacecraft adapter typically procured from Saab Ericsson Space and a transfer compartment provided by RSC Energia.

Figure 2-9 Zenit-3SLB Ascent Unit

(dimensions in millimeters)

Payload Fairing

The payload fairing (PLF) provides environmental protection for the space-craft from encapsulation in the payload processing facility through launch and ascent. The PLF for the Zenit-3SLB is based on the 17S72 fairing manufactured by NPO Lavochkin. It was designed specifically for the Block DM and has an unblemished flight record on Block DM missions dating to 1996. The fairing is a bi-conic, aluminum construction that is 10.4 meters (34.1 feet) in length by 4.1 meters (13.5 feet) in its primary diameter. Spacecraft interfaces provided by the PLF are described in further detail in Section 5.

Fairing Access Characteristics

Once inside the PLF, physical access to the spacecraft is gained through fairing doors. Two doors are standard, one in each fairing half, up to 420 mm x 420 mm (16.5 inches x 16.5 inches) in size. Further information about access doors including allowable locations is provided in Section 5. Because there is no access tower at the Zenit launch pad, the customer/user can directly access their Land Launch payload(s) as late as 28 hours before launch, inside a clean enclosure at the Launch Vehicle Integration Facility. This capability improves opportunities for final adjustments, battery installation and other spacecraft-unique pre-launch operations.

Stage 2/Stage 3Interface Plane

Stage 2/Stage 3Interface PlaneStage 2/Stage 3Interface Plane

PLF/BDM Interface Plane

Spacecraft Interface PlanePayload Support Structure (PSS)

PSS/BDMInterface

SpacecraftPayloadFairing (PLF)

Block DM-SLB (BDM)



Conditioned Air Supply to the Fairing

Clean, conditioned air is provided to the payload fairing volume from en-capsulation until launch including during transport between facilities. Flow rates, cleanliness, temperatures, humidity levels and other details of the clean air supply to the payload volume are provided in Section 4.

Fairing Thermal Protection

The internal and external thermal insulation of the PLF nose cone protects the PLF structure against overheating and preserves acceptable thermal conditions for the spacecraft during ascent. Spacecraft environments are described in Section 4.

Payload fairing jettison is constrained to ensure that the free molecular heating does not exceed the allowable limit defined in Section 4 and that the fairing elements land in pre-approved drop zones.

Payload Support Structure

The payload support structure for the Zenit-3SLB is provided by RSC Energia. It consists of a transfer compartment manufactured by Energia and an industry-standard spacecraft adapter (SCA) typically procured from Saab Ericsson Space Company (PAS937, PAS1194 or PAS1666) that interfaces with the spacecraft. Further details are provided in Section 5. Unique spacecraft base interfaces can normally be accommodated within the stan-dard integration time span.



2.4 The Zenit-2SLB Payload Unit Components and Integration

The Zenit-2SLB payload unit (PLU), shown in Figure 2-10, consists of the spacecraft, fairing, intersection bay, interface truss and spacecraft adapter. These elements are integrated in a Class 100,000 clean environ-ment during ground processing.

Figure 2-10 Zenit-2SLB Payload Unit

Payload Fairing

The PLF for the Zenit-2SLB is based on the Zenit-2 fairing manufactured by PO Yuzhmash. It was designed specifically for the two-stage Zenit and has an extensive and unblemished flight record dating to 1985. The fairing is a mono-conic, aluminum construction that is 13.65 meters (44.8 feet) in length by 3.9 meters (12.8 feet) in its primary diameter, and provides a 3.48 meter (11.4 feet) useable diameter. Spacecraft interfaces provided by the PLF are described in further detail in Section 5.

Alternative and modified fairings are also available. Interested customers are encouraged to contact Boeing Launch Services for further informa-tion.

Fairing Access Characteristics

Access doors up to 500 mm x 500 mm (19.7 inches x 19.7 inches) can be provided in the Zenit-2SLB fairing for this purpose. Additional informa-tion can be found in Chapter 5. Because there is no access tower at the Zenit launch pad, the customer/user can directly access their Land Launch payload(s) as late as 28 hours before launch, inside a clean enclosure at the Launch Vehicle Integration Facility. This capability improves oppor-tunities for final adjustments, battery installation and other spacecraft-unique pre-launch operations.

Fairing

IntersectionBay

Interface TrussSpacecraft Adapter

13.6 m ( 44.8 ft)

14.0 m ( 45.9 ft)

Ø3.

9m

(12.

8 ft)



Conditioned Air Supply to the Fairing

Clean, conditioned air is provided to the payload fairing volume from en-capsulation until launch including during transport between facilities. Flow rates, cleanliness, temperatures, humidity levels and other details of the clean air supply to the payload volume are provided in Section 4.

Fairing Thermal Protection

External thermal insulation protects the payload structure from overheat-ing and the internal thermal insulation limits the interior payload fairing surface temperature. Payload fairing jettison is constrained to ensure that the free molecular heating does not exceed the allowable limit defined in Section 4 and that the fairing elements land in pre-approved drop zones.

Intersection Bay The intersection bay serves to preserve the mating interfaces on the for-ward end of Zenit Stage 2 for the Block DM, thus maximizing inventory flexibility by allowing each stage 2 to be used on any Sea Launch or Land Launch configuration as needed. On Zenit-2SLB the intersection bay also provides a solid base for the payload support structure (truss and adapter) and enables full encapsulation of the spacecraft while in the payload proc-essing facility, creating an enclosed payload volume for easy cleanliness and environmental control with a conditioned air supply.

Spacecraft Adapters

For dedicated launches of a single spacecraft on the Zenit-2SLB, Land Launch can provide the customer any of the available standard adapters manufactured by Saab Ericsson Space, or an adapter provided by SDO Yuzhnoye and PO Yuzhmash using their experience in developing, test-ing and manufacturing adapters and separation systems for past Zenit-2 missions (Fig. 2-11) and for other launchers produced by Yuzhnoye and Yuzhmash (Cyclone, Dnepr). Further information on interfaces is pro-vided in Section 5.

Unique Interfaces and Multi-Spacecraft Launches

For spacecraft that have unique interface and separation requirements, Land Launch can examine other heritage spacecraft adapter designs, in-cluding those that incorporate bolt-type attachment and separation mecha-nisms. Yuzhnoye and Yuzhmash also have extensive experience design-ing and launching multi-spacecraft mechanisms on several launch systems. Unique spacecraft base interfaces, or multi-spacecraft dispens-ers, can normally be accommodated within the standard integration time span.



Figure 2-11. Zenit 2 Spacecraft Adapter Developed by Yuzhnoye and Yuzhmash

O2062

2-я ступень РН LV Stage 2

КАSC



3. PERFORMANCE Overview The Land Launch vehicles, the Zenit-2SLB and the Zenit-3SLB, can

deliver spacecraft to a broad set of orbits. These include low, medium and high Earth orbits (LEO, MEO and HEO), geosynchronous transfer orbits (GTO), highly elliptical orbits, direct geostationary insertion (GEO) and Earth escape trajectories. Data presented in this section is intended to enable prospective users to make preliminary performance assessments. Please contact Boeing Launch Services for a performance quote specific to your mission requirements. Characteristics of performance are covered in Sections 3.1 through 3.8, including: • Launch Window Availability • Launch Site and Accessible Orbits • Generic Ascent Trajectories • Mass Performance • Coast Phase Maneuvers • Injection Accuracy • Spacecraft Separation Conditions



Performance Ground Rules

Performance data in this section is based on the following set of ground rules:

• Payload capability, defined in terms of Payload Systems Mass (PSM), consists of the combined mass of the separated spacecraft and the spacecraft adapter including wire harnesses.

• For preliminary planning of missions manifesting a single payload, spacecraft adapter (and harness) masses of 140 kg and 200 kg are assumed for the Zenit-3SLB and Zenit-2SLB respectively. The masses of dispensers for multiple payloads, typical for Zenit-2SLB missions to LEO, are application unique.

• The maximum PSM for Zenit-3SLB is 5,000 kg due to structural limitations. For Zenit-2SLB the structural limit is not a factor since it exceeds the vehicle’s maximum performance.

• To achieve orbit within the desired accuracy, and perform Contamination and Collision Avoidance Maneuver (CCAM), sufficient propellant reserves are assured for each individual stage to account for all launch vehicle dispersions and possible ambient conditions at any time of day on any day of the year with at least 99.65% probability.

• The spacecraft is injected into orbit via trajectories that are consistent with existing, approved launch corridors and drop zones.

• At the time of fairing jettison, the free molecular heating (FMH) is less than 1,135 W/m2, accounting for all launch vehicle dispersions and possible ambient conditions at any time of day on any day of the year.

• Orbital altitudes are specified with respect to an Earth radius of 6,378 km.

• The Zenit-3SLB uses its standard payload fairing that is 4.1 m in diameter and 10.4 m long.

• The Zenit-2SLB uses its standard payload fairing that is 3.9 m in diameter and 13.65 m long.

• Mission-unique customer requirements that may affect performance (e.g. specific argument of perigee, restricted mission duration, ground station visibility, extended launch windows) are not factored.



Launch Window Availability

The launch vehicle and associated ground systems can support a launch window any day of the year at any time of the day. Furthermore, inherent features of the Land Launch system enable it to provide the maximum flexibility to accommodate shifting satellite readiness dates with little or no perturbation to the launch schedules of other customers (Table 3-1). • Minimal Turn-Around Time - the Zenit launch complex was designed for maximum throughput and minimum refurbishment between launches. The complex can support launches as little as 10 days apart. Factory output limits the theoretical launch rate to twelve per year, of which seven may be Zenit-3SLB. • Robust Flight Hardware – Both the Zenit and Block DM launch systems were designed to withstand environmental conditions at Baikonur. • Heritage Hardware – The Land Launch configurations are composed of heritage systems with as many as 220 flights to their credit. This maturity, combined with robust commit criteria, give Sea Launch and Land Launch the highest launch-on-time probability for heavy and medium lift launch services, respectively. All but two of twelve Sea Launch launches to date have taken place in the first second of the first launch window on the first attempt.

Table 3-1. Launch Operational Features Dates Available year around

Times Available at any hour

Ambient Temperature -29 °C to +45 °C (-20 °F to +113 °F)

Average Ground Winds (at 10m above ground surface)

Zenit-2SLB: 20 m/s (45 miles/hour) Zenit-3SLB: 18 m/s (40 miles/hour)

Pad Turn-around Time Between Launches 10 Days

Nominal Turn-around Time After Launch Scrub

1 Day (if scrub precedes LV fueling) < 3 Days (scrub after LV is fueled)

Maximum Annual Launch Rate (Factory Limited)

Twelve (of which no more than seven Zenit-3SLB)

Launch-on-Time Probability Zenit-2SLB: 98% Zenit-3SLB: 97%



3.1 Launch Site and Accessible Orbits

The coordinates for the Zenit Launch Complex are: latitude = 46 o North, longitude = 63 o East. The currently approved launch azimuths available from this complex, as constrained by overflight and drop zone considerations, are shown below in Table 3-2 and Figure 3-1.

Table 3-2. Zenit Launch Azimuths and Inclinations from Baikonur Azimuth Inclination of Initial Orbit

64.2º 51.4º 35.0º 63.9º

Site Location

194.2º 98.8º

For special cases, arrangements can be made to open a corridor and allocate drop-zones for the launch azimuths of Ao = 82.1o (i = 46.2o) and A0=178.8º (i=88.1º). Approval of new launch corridors for Land Launch is eased by its use of environmentally-friendly fuels.

A=64.20, i=51.40

A=82.10, i=46.20

A=178.80, i=88.10A=194.20, i=98.80

A=35.00, i=63.00

A=64.20, i=51.40

A=82.10, i=46.20

A=178.80, i=88.10A=194.20, i=98.80

A=35.00, i=63.00

Approved AzimuthPotential Azimuth

Baikonur

Figure 3-1. Flight Corridors for Land Launch from Baikonur



Table 3-3 shows the orbit inclinations (i) that can be reached by Land Launch from its three approved launch corridors. LEO orbit inclinations several degrees different from the three approved launch corridors can be obtained by cross-range yawing maneuvers (“doglegs”) of the second stage commencing after fairing jettison. Such maneuvers are generally associated with missions provided by the Zenit-2SLB, where LEO is the final destination. For Zenit-3SLB missions involving higher orbits in which the desired inclination differs from the three approved corridors, it is typically most efficient for plane changes to be carried out primarily by the Block DM-SLB third stage. In these cases, the first two stages usually perform a direct ascent into a parking orbit inclination coinciding with one of the approved corridors.

Table 3-3. Accessible Orbits on Land Launch

Orbit Type Accessible Inclinations Vehicle Usual Plane Change

Method

LEO 46.2 < i < 71 o 84.0 < i < 105 o Zenit-2SLB Second Stage Yaw

MEO, HEO, GTO, Elliptical, escape trajectories

0.0 < i < 110 o Zenit-3SLB

Third Stage Perigee, Apogee or Post-Perigee Burn (mission-specific)

Accessible Orbits

Performance losses due to plane changes are highly sensitive to a variety of mission parameters. Consequently, prospective Land Launch customers are encouraged to contact Boeing Launch Services for a performance estimate that is specific to their needs.



3.2 Ascent Trajectory – Generic Zenit-3SLB GTO Mission Mission Profile For GTO missions the Zenit-3SLB flies a classic three-burn Block DM

mission profile (Figure 3-2) using the approved corridor and drop zones at Ao=64.2°, i=51.4° (Figure 3-3).

Block DM-SLB second burnInjection into transfer orbit

Zenit Stage I/IITransfer orbitHP = 200 kmHA = 35,950 kmInc = 48.6 deg

Parking orbitHP = 180 kmHA = 417 kmInc = 51.4 deg

Block DM-SLB third burnInjection into target transfer orbit

Target transfer orbitHP = 4,100 kmHA = 35,786 kmInc = 23.2 deg

Block DM-SLB first burnInjection into parking orbit

Figure 3-2. Land Launch Uses the Proven Three-Burn Block DM Mission Profile From Baikonur for GTO Launches (orbit parameters correspond to PSM=3600 kg)



First StageL=∼884 km

FairingL=1924 km

Second StageL=6850 km

Launch point

Figure 3-3. Approved Land Launch Ground Track and Drop Zones for GTO Missions

Stage 1 Flight The Zenit first stage provides the thrust for the first 149 seconds of flight.

The roll maneuver begins at 10 seconds after launch. During the final seconds of its burn the engine is throttled to limit the maximum axial acceleration. The approved drop zone for the separated first stage is at distance of approximately 884 km from the launch point, within the Republic of Kazakhstan as shown in Figure 3-3. Throughout this phase of the mission, telemetry is received by ground stations within Baikonur cosmodrome.



Stage 2 Flight The Zenit Stage 2 vernier engine ignites just prior to first stage separation.

Upon first/second stage separation, the first stage solid retrorockets fire and second stage main engine ignition occurs. The second stage main and vernier engines continue to operate in tandem for the next five minutes of flight. After second stage main engine cut-off, the vernier engine continues to function for 75 seconds to provide attitude control up through second/third stage separation. Payload fairing jettison occurs at approximately 320 seconds into flight (175 seconds into second stage operation) with the drop zone located in Siberia approximately 1924 km downrange of the launch site. At this point the free molecular heating rate has dropped to below 30 W/m2, well below the industry norm of 1,135 W/m2. Cross-range yaw maneuvers by the second stage, if required, take place after fairing separation. Telemetry coverage during second stage flight is typically provided by ground stations at Baikonur cosmodrome, and at Krasnoyarsk in Russia. The second stage drop zone is located within the neutral waters of the Pacific Ocean at a downrange distance of 6850 km.



Block DM-SLB Powered Flight

At approximately 65 to 90 seconds after second stage main engine shutdown and an altitude of 180 to 400 km, the second stage vernier engine shuts down. This event is quickly followed by second/third stage separation and the subsequent jettison of the middle adapter surrounding the Block DM-SLB. The Block DM-SLB can perform one to three burns. For most multiple-burn missions, including the generic three-burn GTO mission described here, the initial burn establishes a stable parking orbit, begins approximately ten seconds after separation of the second stage and lasts approximately 200 seconds, with telemetry coverage provided from Krasnoyarsk. The Block DM-SLB then begins a coast in the parking orbit lasting about 64 minutes. Attitude control during Block DM-SLB coast phases is provided by its two attitude control/ullage engines. The second Block DM-SLB burn occurs at the first ascending node of the parking orbit, over the Atlantic Ocean, to transfer to an intermediate elliptical orbit with a synchronous or super-synchronous apogee as dictated by customer requirements and the capabilities of the satellite platform. Ignition starts at approximately 75 minutes after launch and typically continues for approximately 6 minutes, with telemetry coverage provided by a mobile receiving station. After a 5-hour coast the Block DM-SLB and payload reach GTO apogee, where a third burn is performed to optimize the delivery orbit by raising perigee and reducing inclination. Telemetry coverage during the third burn is simplified by the altitude at which it occurs, and is typically provided by multiple sites located at Moscow, Baikonur, Krasnoyarsk and elsewhere. The target injection orbit for a payload mass of 3600 kg features a perigee of 4100 km, an apogee of 35786 km and inclination of 23.2°, resulting in a velocity shortage of 1500 meters/second required to achieve GEO. Payloads lighter than 3600 kg are delivered to orbits requiring progressively less than 1500 meters/second delta-velocity to GEO to the point that payloads weighing 1,600 kg and less are inserted directly into GEO, a mission that the Block DM family has already performed more than one hundred times. Spacecraft separation conditions and post-separation events including collision avoidance maneuvers are described in Section 3.8.



Flight Timeline Table 3-4 provides a typical sequence of events for a representative three-

burn Zenit-3SLB mission to GTO for a 3600-kg payload. Event timing is only slightly dependent on payload mass. Apart from spacecraft separation, variation (dispersion) of any planned event timing for a nominal mission is typically within 15 seconds from the reference sequence.

Table 3-4. Flight Timeline— GTO Mission by the Zenit-3SLB with Three

Burns of the Block DM-SLB

Time [seconds] Event 0 Ignition

∼3.9 Liftoff 12 Begin pitch over 14 Roll to launch azimuth 59 Maximum dynamic pressure 115 Maximum axial acceleration

115 to 132 Stage 1 engine throttle down to 74% 144 Stage 2 vernier engine ignition 147 Stage 1 engine shutdown 149 Stage 1 separation 154 Stage 2 main engine ignition 319 Payload fairing jettison 432 Stage 2 main engine shutdown 507 Stage 2 vernier engine shutdown 508 Stage 2 separation 509 Block DM-SLB middle adaptor jettison 517 Block DM-SLB main engine ignition #1 707 Block DM-SLB main engine shutdown #1 4534 Block DM-SLB main engine ignition #2 4864 Block DM-SLB main engine shutdown #2 23562 Block DM-SLB main engine ignition #3 23631 Block DM-SLB main engine shutdown #3

Mission-Specific Spacecraft separation



Ground Track Figure 3-4 presents the predicted ground track of injection for a generic,

representative Zenit-3SLB three-burn GTO mission.

Launch point 1-st ignition ofDM-SLB ME

LV operation phase2-nd ignition ofDM-SLB ME

3-rd ignition ofDM-SLB ME

Figure 3-4. Injection ground track for a generic Zenit-3SLB GTO mission.



3.3 Ascent Trajectory – Generic Zenit-2SLB Mission to 51.6o LEO Stage 1 Flight The two-stage Zenit-2SLB is optimized for LEO missions inclined at

51.6o, including potential flights to the International Space Station (ISS). For such missions, the Zenit Stage 1 uses the same approved launch corridor and drop zone that is used for GTO missions, along launch azimuth 64.2o (inclination 51.4 o). Liftoff occurs 3.9 seconds after ignition, upon release of the hold-downs. The roll maneuver begins at 10 seconds into flight. Main engine thrust is provided for the first 140 -150 seconds of flight, and the engine is throttled during its last seconds of operation in order to limit maximum axial acceleration. The drop zone for the first stage is 884 km down range from the launch point, within the Republic of Kazakhstan. Throughout this phase of the mission, telemetry is received by ground stations within Baikonur cosmodrome.

Stage 2 Flight The Zenit Stage 2 steering engine ignites prior to first stage separation.

Upon first/second stage separation, the first stage solid retrorockets fire and second stage main engine ignition occurs. The main engine and vernier engine continue to operate in tandem for the next four minutes of flight. Fairing jettison occurs at about 295 seconds of flight (150 seconds into second stage operation), consistent with the approved drop zone located in Siberia approximately 1924 km downrange of the launch site. At this point the free molecular heating rate has dropped to below 30 W/m2, well below the industry norm of 1,135 W/m2. After fairing jettison, the second stage performs a cross-range yaw maneuver to adjust the inclination to 51.6o. After second stage main engine cut-off, the vernier engine continues to function for an additional 500 seconds (as long as 890 seconds on other missions) to provide attitude control up through payload separation. Throughout this phase of flight, telemetry is received by the ground stations within Baikonur and Krasnoyarsk. Spacecraft separation conditions and post-separation events including collision avoidance maneuvers are described in Section 3.8.



Flight Timeline Table 3-5 provides a typical sequence of events for a representative

Zenit-2SLB mission that delivers 12,000 kg to a 51.6o-inclined, 400 km low Earth orbit, i.e. – one compatible with ISS access .

Table 3-5. Flight Timeline—Zenit-2SLB ISS Mission

Time [seconds] Event 0 Ignition

∼3.9 Liftoff 10 Begin roll maneuver 11 Begin pitch over 14 Roll to launch azimuth 60 Maximum dynamic pressure 113 Maximum axial acceleration

113 tо 132 Stage 1 engine throttle to 50% 145 Stage 2 vernier engine ignition 147 Stage 1 engine shutdown 149 Stage 1 separation 155 Stage 2 main engine ignition 295 Payload fairing jettison 397 Stage 2 main engine shutdown

893.5 Stage 2 vernier engine shutdown 893.8 Spacecraft separation pyrotechnic firing 893.86 Solid-propellant retro rocket burn



Flight Profile Figure 3-5 graphically portrays the flight profile defined in Table 3-5, along with other key trajectory events and parameters.

SC SeparationTime=894 sAltitude=400 km

Fairing JettisonTime=295sAltitude=173 kmFMH≈30 W/m2

Stage 1 SeparationTime=149 sAltitude=74 km

Maximum QTime=60 sQ=5370 kgf/m2

Stage 1 ImpactRange=884 km

Fairing ImpactRange=1924 km

Max AccelerationTime=113 sAccel=4.06 g

Stage 2 MECOTime=397 sAltitude=400 km

Figure 3-5. Typical Ascent Profile to the International Space Station Orbit at 51.6o

with Payload Mass 12000 kg



Ground Track Figure 3-6 presents the predicted ground track for a Zenit-2SLB mission

to the International Space Station.

-9 0

-8 0

-7 0

-6 0

-5 0-4 0

-3 0

-2 0

-1 0

0

1 0

2 0

3 0

4 05 0

6 0

7 0

8 0

9 0

-1 8 0 -1 5 0 -1 2 0 -9 0 -6 0 -3 0 0 3 0 6 0 9 0 1 2 0 1 5 0 1 8 0

Stages 1 and 2Operation

Figure 3-6. Flight Ground Track for a Zenit-2SLB Mission to 51.6o LEO



3.4 Payload Capability – Three Stage Zenit-3SLB Geosynchronous Transfer Orbit

The Land Launch Zenit-3SLB is a medium-lift vehicle to GTO. Employing three burns of the Block DM-SLB, it can deliver payloads weighing 3.6 metric tons to a GTO featuring a high perigee and reduced inclination, requiring 1500 m/s in additional velocity to attain geostationary or geosynchronous orbit (GEO). Performance improves rapidly for lighter satellites because correspondingly less fuel is off-loaded from the Block DM-SLB to meet a second stage drop zone constraint. Table 3-6 and Figure 3-7 show the GTO payload capability.

Table 3- 6. Zenit-3SLB Payload Capability to GTO

Delta-V to GEO [meters/second]

Inclination [degrees]

Perigee Altitude [kilometers]

Payload Systems Mass [kilograms]

0 0.00 35,786 1,600 1,000 13.0 9,430 2,830 1,500 23.2 4,100 3,600 1,800 31.0 2,120 4,120

Notes and Assumptions: • Apogee altitude of 35,786 km • Three burns of the Block DM-SLB • Mission duration approximately 6.6 hours

0

500

1000

1500

2000

2500

3000

3500

4000

4500

0 200 400 600 800 1000 1200 1400 1600 1800 2000

Delta V to target orbit, m/s

Pay

load

Mas

s, k

g

Figure 3-7. Zenit-3SLB Payload Capability to GTO



MEO, HEO, Circular and Elliptical Orbits

The Land Launch Zenit-3SLB is a heavy lift vehicle to Middle Earth and High Earth (MEO and HEO, respectively) circular and elliptical orbits that coincide with its approved launch corridors, as shown in Tables 3-7 and 3-8 and in Figures 3-8 and 3-9. MEO, HEO and elliptical orbits at other inclinations can also be obtained, typically with an additional burn of the Block DM-SLB, at a cost in performance that varies with altitude and the extent of plane change required. LEO (altitude<1000 km) and low-perigee elliptical orbits are more optimally performed by a Zenit-2SLB, as shown in a later section of this chapter. Customers are encouraged to contact Boeing Launch Services for a specific performance quotation.

Table 3-7. Zenit-3SLB Performance to Circular Orbits

Payload Capability [kg] Height [km] Inclination

51.4 o Inclination

63.9 o Inclination

98.8 o 1,000 5000 5000 5000 5,000 5000 5000 5000

10,000 4830 4340 3890 20,000 3400 3020 2570 30,000 2880 2540 2110

Note: Two burns of the Block DM-SLB main engine

2000

2500

3000

3500

4000

4500

5000

5500

0 5000 10000 15000 20000 25000 30000 35000

Orbital Height, km

Pay

load

Mas

s, k

g

i=51,4

i=63,9

i=98,8

Figure 3-8. Zenit-3SLB Performance to Circular Orbits



Table 3-8. Zenit-3SLB Performance to Elliptical Orbits Payload Capability [kg] Apogee

Height [km]

Inclination 51.4 o

Inclination 63.9 o

Inclination 98.8 o

10,000 5000 5000 5000 20,000 5000 5000 5000 30,000 5000 4850 4680 40,000 5000 4540 4320 50,000 4810 4320 4090 60,000 4650 4170 3920 70,000 4530 4050 3810

Assumptions: • Single Block DM-SLB burn • Perigee altitude of ~200 km

38003900400041004200430044004500460047004800490050005100

10000 20000 30000 40000 50000 60000 70000Height of target orbit apogee, km

Payl

oad

Mas

s, k

g

i=51,4i=63,9i=98,8

38003900400041004200430044004500460047004800490050005100

10000 20000 30000 40000 50000 60000 70000Height of target orbit apogee, km

Payl

oad

Mas

s, k

g

i=51,4i=63,9i=98,8

Figure 3-9. Zenit-3SLB Performance to Elliptical Orbits (Perigee 200 km)



High-Energy and Earth-Escape Trajectories

Table 3-9 and Figure 3-10 show the Zenit-3SLB payload capability to high-energy orbits and Earth escape. These are presented as a function of C3 (velocity-at-infinity squared).

Table 3-9. Zenit-3SLB High-Energy and Earth Escape Payload Capability

C3 [km2/s2]

Payload Capability [kg]

-20 5000 -10 4620

0 3780 15 2740 30 1900

Notes and Assumptions: • Inclination = 51.4o • Perigee altitude = 300-450 km • Single Block DM-SLB burn

0

1000

2000

3000

4000

5000

6000

-30 -20 -10 0 10 20 30 40C3, km2/s2

Payl

oad

Syst

ems M

ass,

kg

Figure 3-10. Zenit-3SLB High-Energy and Earth Escape Payload Capability



3.5 Payload Capability - Two Stage Zenit-2SLB Circular LEO Orbits

Table 3-10 and Figure 3-11 present Zenit-2SLB payload performance as a function of both circular orbit altitude and inclination.

Table 3-10 Zenit-2SLB Payload Capability for Circular Low Earth Orbits

Payload Mass [kg] Altitude [km] Inclination

51.4º Inclination

63.9º Inclination

98.8º 200 13,920 13,330 10,610 300 12,940 12,410 9,790 400 11,930 11,500 8,870 500 10,890 10,550 7,910 600 9,820 9,570 6,930 700 8,730 8,560 5,930 800 7,630 7,550 4,940 900 6,530 6,510 3,940

1,000 5,420 5,480 3,320 1,100 4,660 4,560 2,920 1,200 4,250 4,190 2,530 1,300 3,810 3,750 2,320 1,400 3,390 3,310 2,030 1,500 2,930 2,340 1,520

1000

3000

5000

7000

9000

11000

13000

15000

200 400 600 800 1000 1200 1400 1600

Circular Orbit He ight (km)

Pay

load

Mas

s (k

g)

51.4º63.9º98.8º

Figure 3-11. Zenit-2SLB Payload Capability for Circular Low Earth Orbits



Elliptical Orbits Table 3-11 and Figure 3-12 define the performance parameters for the two-stage Zenit-2SLB to various elliptical earth orbits.

Table 3-11. Zenit-2SLB performance to Elliptical Orbits

Payload Mass [kg] Apogee [km] Inclination

51.4º Inclination

63.9º Inclination

98.8º 500 13280 12730 10070

1,000 12320 11800 9250 2,000 10710 10230 7870 4,000 8290 7900 5830 6,000 6560 6290 4380 8,000 5260 5120 3310 10,000 4250 4230 2480

Note: Perigee altitude = 200 km

0

2000

4000

6000

8000

10000

12000

14000

0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Apogee altitude (km)

Payl

oad

mas

s (k

g)

51.4º63.9º98.8º

Figure 3-12. Zenit-2SLB Performance to Elliptical Orbits (Perigee 200 km)



3.6 Coast Phase Attitude Maneuvers Zenit-3SLB During coast phases the Block DM-SLB control system can provide

three axes pointing (pitch, yaw and roll) with accuracy up to ±3 deg in all three axes. The control system of the Block DM-SLB, unlike other versions of the Block DM, can also provide continuous roll around the longitudinal axis or one of the lateral axes at a rate up to 5 degrees per second. Forty minutes of any coast phase are nominally reserved for Block DM-SLB attitude maneuvers

Zenit-2SLB Zenit-2SLB missions do not feature extended coasts. Stage 2 operation immediately succeeds stage 1 operation, and payload separation occurs between 0.3 and 5 seconds after cut-off of the second stage vernier (steering) engine.



3.7 Injection Accuracy Tables 3-12 and 3-13 show 3σ orbital injection accuracy of the Land

Launch family of vehicles to representative orbits.

Table 3-12. Land Launch Zenit-2SLB and Zenit-3SLB Provide Accurate Orbital Insertion

Zenit-2SLB Zenit-3SLB Orbital Parameter Circular (1) Circular (2) Circular (3) GTO (4)

Altitude [km] ± 8 ± 9 ± 25 - Perigee [km] - - - ± 40 Apogee [km] - - - ± 100 Inclination [deg] ± 0.04 ± 0.07 ± 0.06 ± 0.1 Longitude of Ascending Node [deg] ± 0.1 ± 0.07 ± 0.2 ± 0.3

Perigee Argument [deg] - - - ± 0.2 Period [sec] ± 3.5 ± 4.5 ± 45

(1) 400 km x 400 km, inclination = 51.6° (2) 600 km x 600 km, inclination = 98° (3) 10,000 km x 10,000 km, inclination = 51.4° (4) 4,000 km x 35,786 km, inclination = 23°

Table 3-13. The Zenit-3SLB Also Provides Accurate Direct GEO Insertion

Orbit Type Orbital Altitude Inclination Period

Geostationary ± 200 km ± 0.2 deg. ± 450 s



3.8 Spacecraft Separation and Post-Separation Events 3.8.1 Zenit-3SLB Separation Event Spacecraft separation typically occurs 10-15 minutes after the final Block

DM-SLB main engine shutdown. This allows for reorientation to the required spacecraft separation attitude.

Separation Capabilities

The separation system provides a relative velocity between the Block DM-SLB and the spacecraft, typically on the order of 0.3 meters/second. The separation springs can provide a straight push-off or a transverse angular rate. Attitude and attitude rate accuracy depend heavily on spacecraft mass properties and spin rate, and may be assumed to be + 2.5 degrees and + 0.5 degrees/second in all three axes for a non-spinning separation (2.3σ). The Block DM-SLB attitude control system can provide a longitudinal spin rate up to 5 degrees per second if desired. For spacecraft requiring a transverse spin at separation, this may be provided up to 2 degrees per second within +/- 0.5 degrees per second about each axis.

CCAM After spacecraft separation, the Block DM-SLB performs a Collision and Contamination Avoidance Maneuver (CCAM), which prevents future contact with the spacecraft. The timing of this maneuver is determined for the specific mission. The Block DM-SLB then vents all residual propellant and gasses, and depletes any remaining charge in its batteries.

State Vector Delivery

The state vector at time of spacecraft separation may be delivered to the customer 35-50 minutes after the event. The format of the state vector, means of its delivery and the parameters of the spacecraft injection orbit are agreed in advance between the parties. The time of delivery of data can be updated for the specific mission.



3.8.2 Zenit-2SLB Separation Event Separation begins between 0.3 and 5 seconds after shutdown of the

second stage vernier (steering) engine.

Separation Capabilities

The Zenit-2SLB employs a typical three-axis stabilized method for payload separation along the second stage’s longitudinal axis. The actual separation is initiated by the firing of pyrotechnic ordnance charges in the spacecraft attachment assembly. The separation impulse to the spacecraft is typically provided by springs in the separation system. Nearly simultaneously, solid propellant retro-rockets on the aft end of the second stage are fired, adding to the relative separation velocity. Launch vehicle stabilization errors at the moment of spacecraft separation command generation can be kept within +/- 2 degrees for pitch and yaw and within +/- 1 degree for roll. Angular velocities at release can be kept within +/- 1.5 degrees/sec for all three axes. Table 3-14 presents typical parameters for payload motion after separation in the case of a single spacecraft, while Table 3-15 presents similar data for missions involving multiple payloads with individual masses that exceed 500 kg.

Table 3-14 Typical Spacecraft Motion After Separation - Single Payload Parameter Value

Relative separation velocity ≥ 2.8 m/s Spacecraft angular rate around any of its axes ≤ 2.5 deg/s Spacecraft attitude error ± 2 deg

Table 3-15 Typical Spacecraft Motion After Separation - Multiple Payloads (each > 500 kg)

Parameter Value Relative separation velocity ≥ 0.3 m/s Spacecraft angular rate around any of its axes ≤ 4.0 deg/s Spacecraft attitude error ± 5 deg



CCAM for second stage

Collision avoidance is achieved by firing four solid-propellant retrorockets on the aft end of the second stage for a burn time on the order of 0.5 to 1.1 seconds, slowing the second stage and moving it out of the spacecraft orbit. After a delay, the oxidizer tank is vented.

State Vector Delivery

The timing of state vector delivery depends on the mission profile as well as the location and the availability of ground stations. For a typical ascent to 51.4o, it is possible to arrange for delivery of such data to the customer between 35 and 50 minutes after payload separation.



4. SPACECRAFT ENVIRONMENTS Overview This section describes the major environments to which the spacecraft is

exposed from the time of its arrival at Baikonur cosmodrome until its separation from the launch vehicle during flight.

These environments and conditions include:

• Structural loads • Thermal

• Random vibration • Humidity

• Acoustics • Pressure venting

• Shock • Contamination

• Electromagnetic radiation

Unless otherwise noted, the payload environments presented in this sec-tion are common to both the three-stage Zenit-3SLB and the two-stage Zenit-2SLB.

Ground and Flight Environments

Those levels associated with “ground handling and transportation” ad-dress the period from the arrival of the spacecraft at Baikonur until Stage 1 ignition and liftoff.

Those levels designated as “flight” cover the subsequent period from liftoff command through spacecraft separation.

Reference Coordinate System

The coordinate system used in this section is shown in Figure 4-1. Dur-ing transfer of the spacecraft in its shipping container from the airport to the Payload Processing Facility (PPF), the +X axis coincides with the direction of travel. At other times, X coincides with the longitudinal axis of the launch vehicle. The Y axis is vertical during horizontal ground op-erations.



YVertical

XLongitudinal

ZLateral

Figure 4-1. Reference Coordinate System Used for Defining Spacecraft

Environments

Environmental Monitoring

Land Launch monitors and records spacecraft environments as specified in the Spacecraft-LV Interface Control Document and documents these results in the post-flight report to the customer. Typically, this includes:

• flight environments (accelerations, acoustics, shock, fairing thermal conditions during ascent, pressure decay, etc.);

• the temperature, humidity and cleanliness levels of the spacecraft processing and encapsulation areas while the spacecraft is present;

• the temperature, humidity and cleanliness levels of the conditioned air provided to the fairing with the spacecraft inside;

• accelerations experienced during all phases of ground processing, after the spacecraft has been removed from its shipping container

The responsibility normally resides with the customer to monitor the spacecraft environment (including accelerations) until it is unloaded from its shipping container at the cosmodrome.



4.1 Structural Loads Overview Design reference structural loading environments on spacecraft primary

and secondary structures are defined here for:

• ground transportation and handling

• flight

• spacecraft sinusoidal vibration testing

Spacecraft compliance requirements related to these environments are presented in Section 7.

Quasi-Static Load Factors, Ground Handling and Transportation

Design reference maximum acceleration levels during ground transporta-tion, handling and processing are defined in Table 4-1. The quasi-static accelerations levels are shown for the spacecraft center of gravity while in a horizontal orientation. These accelerations can be applied simultane-ously in the longitudinal, lateral and vertical directions (the axis X coin-cides with the velocity vector). During erection of the launch vehicle to a vertical position on the launch pad, the maximum acceleration of the spacecraft center of gravity is 1.5 g. In the course of combined opera-tions the spacecraft briefly transitions through various vertical orienta-tions, during which the respective axial accelerations are maintained within those limits already specified. The maximum rate of angular ac-celeration (about any axis) during crane lifts of the ILV is 0.055 radians per seconds squared.

Table 4-1. Maximum Quasi-Static Accelerations During Ground Operations

Acceleration [g] Spacecraft Processing Operation X Y Z Safety Factor

Transfer from the airport to the PPF ± 1.0 -1 ± 1.0 ± 0.4 2.0 Horizontal Combined Operations (from mating with the launcher in the PPF through completion of launcher assembly in Area 42)

Zenit-3SLB Zenit-2SLB

± 0.5 ± 0.35

-1 ± 0.5 -1 ± 0.2

± 0.4 ± 0.2

1.5 2.0

Launcher on-loading and off-loading (crane lifts in Area 42) ± 0.2 -1 ± 0.2 ± 0.2 1.5

Roll out and erection on the launch pad ± 0.35 -1 ± 0.2 ± 0.2 1.5



Quasi-Static Load Factors, Flight

From liftoff through spacecraft separation, the spacecraft is subjected to quasi-static steady-state and low-frequency dynamic accelerations. Fig-ure 4-2 provides the design reference accelerations for critical loading events. These accelerations are applied at the spacecraft center of gravity and are intended for preliminary design only. Determining the ability of specific spacecraft primary and secondary structures to withstand the dy-namic loading events during flight requires a coupled loads analysis (CLA), which will be performed for each mission. When generated and verified, CLA results supersede the generic quasi-static accelerations provided in Figure 4-2.

-3

-2

-1

0

1

2

3

4

5

-3 -2 -1 0 1 2 3

(+1, -2)

(+2, +2)

(+0.7, +4.5)(-0.7, +4.5)

(-2, +2)

(-1, -2)

Lateral (g)

Long

itudi

nal*

(g)

* positive longitudinal quasi-static accelerations are aligned with the direction of flight

Figure 4-2. Typical Quasi-Static Design (Maximum Expected) Loads in Flight



Sinusoidal Equivalent Vibration During Flight

The longitudinal and lateral low-frequency sinusoidal vibration environ-ments generated at the spacecraft separation plane during liftoff and flight phases are within the limits defined in Table 4-2. The sinusoidal vibration environment for all major flight events are specifically deter-mined for each mission during the CLA. These results determine the maximum notching in the environment spectra that can be used during spacecraft sinusoidal vibration testing.

Table 4-2. Sinusoidal Vibrations at the Spacecraft Interface

Frequency Range [Hz] Vehicle Amplitude [g] Zenit-2SLB 0.6

5 - 100 (Longitudinal and Lateral) Zenit-3SLB 0.7



4.2 Random Vibration Ground Random Vibration for Components Near the Spacecraft Interface

The spacecraft is subjected to low frequency random vibrations during transportation by rail at the cosmodrome. Table 4-3 envelopes this ran-dom vibration environment. The maximum duration of any rail transfer is six hours.

Table 4-3. Random Vibration During Ground Transport When the Spacecraft is Not in the Customer Container

Spectral Density of Power [g2/Hz]

Frequency [Hz]X-X

Longitudinal Y-Y

Vertical Z-Z

Lateral 2 0.000075 0.00015 0.00015 4 0.000575 0.0033 0.00033 8 0.002 0.0032 0.00066

10 0.0006 0.0032 0.0008 14 0.00028 0.000833 0.00033 20 0.000275 0.00015 0.00032 25 0.000275 0.00015 0.00031 30 0.000275 0.00015 0.0003 35 0.0005 0.00015 0.000185 40 0.00018 0.00015 0.000037 45 0.000125 0.00015 0.000037 50 0.000125 0.00015 0.000037

Flight Random Vibration Environment

The random vibration environment during flight at the spacecraft inter-face is enveloped in Table 4-4 and Figure 4-3.

Maximum values occur during liftoff and are closely correlated with the acoustic environment.

The environment applies to components within 0.5 m (20 inches) from the separation plane along any structural path. This environment is not to be applied to the complete spacecraft as a rigid base excitation.

Table 4-4. Random Vibration Environment During Flight

Frequency [Hz] Spectral Density [g2/Hz] 20 – 100 0.01 … 0.035 100 – 700 0.035 700 – 2000 0.035 … 0.01

Overall Level 6.8 grms



0,001

0,01

0,1

1

100 1000 10000

Frequency [Hz]

Spec

tral

Pow

er D

ensi

ty [g

2 /Hz]

10

Figure 4-3. Random Vibration Environment During Flight



4.3 Acoustics Fairing Volume Average Sound Pressure Levels

Maximum acoustic pressures occur during lift off and transonic phases of flight. Acoustic characteristics inside the Land Launch fairings are en-veloped in Table 4-5 and Figures 4-4 and 4-5.

Table 4-5. Maximum Expected Acoustic Pressure Envelope Inside Land Launch Fairings

Acoustic Pressure Level [dB] 1/3 Octave Band Center Frequency [Hz] Zenit-2SLB Zenit-3SLB

31.5 119 119 40 121 121 50 123 123 63 125 125 80 127 128

100 128 129 125 129 130 160 130 131 200 131 133 250 130 134 315 129 133 400 128 131 500 127 129 630 126 127 800 125 125

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

OASPL 140 142 Duration 40 seconds 60 seconds

Reference: dB in respect to 2 x 10-5 Pa (2.9 x 10-9 psi)



100

110

120

130

140

10 100 1000 10000

1/3 octave band center frequency, Hz

Sou

nd p

ress

ure

leve

ls, d

B

Note:Overall acoustic pressure level = 140 dB

Figure 4-4. Maximum Expected Acoustic Pressure Envelope Inside the Zenit-2SLB Fairing

100

110

120

130

140

10 100 1000 10000

1/3 octave band center frequency, Hz

Acou

stic

pre

ssur

e le

vels

, dB

Note:Overall acoustic pressure level = 142 dB

Figure 4-5. Maximum Expected Acoustic Pressure Envelope Inside the Zenit-3SLB Fairing



4.4 Shock Overview

The maximum shock at the spacecraft interface occurs at the moment of spacecraft separation. Other shock inputs, including those associated with fairing jettison and stage separations, are within this envelope.

Zenit-3SLB The maximum expected interface shock response spectrums for the 937-

mm, 1194-mm and 1666-mm diameter interfaces are presented in Table 4-6 and Figures 4-6a through 4-6c as a function of clamp band tension-ing, when using currently available Saab Ericsson Space (Saab) space-craft adapters. Maximum SC mass and center-of-gravity corresponding to these band tensions are shown in Section 5. The shock environment may differ if other adapters are used. Customers interested in other adapters are encouraged to contact BLS for further information.

Table 4-6. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB Adapters

Shock Response Spectra (g)

937 Interface 1194 Interface 1666 InterfaceBand Tension (kN) Band Tension (kN) Band Tension Frequency

(Hz) 12.5 20 30 10 20 30 40 30 kN 100 50 50 50 60 150 200 60 60 60 130 130 140 170 400 800 550 600 650 800 900 1150 1450 3,000 1300 1150 1300 1550 1500 1800 2400 3000 3,000 2000 2300 2700 3300 1850 2250 3000 3750 3,000 3000 2600 3050 3750 2300 2750 3600 4600 3,000 3500 2700 3150 3900 2500 3000 4000 5000 3,100 6000 3200 3700 4600 2500 3000 4000 5000 3,500 8000 3500 4000 5000 2500 3000 4000 5000 3,500 9000 3500 4000 5000 2500 3000 4000 5000 3,500 10000 3500 4000 5000 2500 3000 4000 5000 3,800

• Q factor of 10 • SRS levels are simultaneous in three mutually perpendicular directions • As measured at 50 mm (2 inches) from the separation plane on the spacecraft side of the interface



Figure 4-6a. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) WIth Standard SAAB 937-mm Adapter, Various Band Tensions

Figure 4-6b. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 1194-mm Adapter, Various Band Tensions



Shock levels at separation plane

10

100

1000

10000

100 1000 10000

Frequency [Hz]

Res

pons

e [g

] Q=1

0

Figure 4-6c. Zenit-3SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 1666-mm Adapter, 30 kN Band Tension



Zenit-2SLB The maximum expected interface shock response spectrums for a single

satellite using the Zenit-2 or SAAB 2624-mm interfaces are presented in Table 4-7, with the SAAB information also pictured in Figure 4-6d. The shock environment may differ if other adapters are used, or if more than one satellite is launched at a time. Customers interested in other adapters or group launches are encouraged to contact BLS for further information.

Table 4-7. Zenit-2SLB Spacecraft Shock Response Spectra (SRS)

Shock Response Spectra (g) Zenit-2 Adapter (Truss) SAAB 2624 Interface

Frequency (Hz) SRS Frequency (Hz) SRS

100-200 25-100 100-520 10-1800 200-500 100-350 520-3200 1800-5000 500-1000 350-1000 3200-10000 5000 1000-2000 1000 2000-5000 1000-3000

• Q factor of 10 • SRS levels are simultaneous in three mutually perpendicular directions • As measured at 50 mm (2 inches) from the separation plane on the space-

craft side of the interface

Shock levels at separation plane

10

100

1000

10000

100 1000 10000

Frequency [Hz]

Res

pons

e [g

] Q=1

0

Figure 4-6d. Zenit-2SLB Spacecraft Shock Response Spectra (SRS) With Standard SAAB 2624-mm Adapter



4.5 Electromagnetic Environment Overview The spacecraft will experience electromagnetic radiation stemming from:

• The background, or ambient, cosmodrome environment during ground processing;

• Ground emitters actively used during launch operations on the Zenit launch pad and during launch;

• The launch vehicle itself

Each of these sources is defined below.

Coordination It is necessary to coordinate the operation of spacecraft transmitters and other electronic equipment with emissions by the launch vehicle and sources at the cosmodrome. This is performed as part of the integration process. Allowable spacecraft emissions are described in Section 7.

Ambient Cosmodrome Electromagnetic Environment

The ambient cosmodrome electromagnetic environment varies by loca-tion, and changes over time as new equipment is introduced and older equipment is retired. Figures 4-7, 4-8 and 4-9 therefore provide prelimi-nary maximum values for electromagnetic fields levels in Land Launch facilities where the spacecraft will be present: respectively the two avail-able Payload Processing Facilities, the Launcher Assembly Building and the Launch Complex. These environments will be updated during the integration process.



40

50

60

70

80

90

100

110

120

130

140

0,01 0,1 1 10 100 1000 10000

Frequency, MHz

Fiel

d In

tens

ity, d

B µ

V/m

Figure 4-7a. Ambient Electromagnetic Environment within Payload Processing Facility

Site 254

Fiel

d In

tens

ity, d

B µ

V/m

Frequency, MHz

40

50

60

70

80

90

100

110

120

130

140

0,01 0,1 1 10 100 1000 10000

Figure 4-7b. Ambient Electromagnetic Environment within Payload Processing Facility

Area 31



40

50

60

70

80

90

100

110

120

130

140

0,01 0,1 1 10 100 1000 10000

Частота, МГц

Напря

жен

ность эл

ектр

.пол

я, дБмкВ

/м

Frequency, MHz

Fiel

d In

tens

ity, d

B µ

V/m

Figure 4-8. Ambient Electromagnetic Environment within the Launch

Vehicle Assembly Building (Area 42)

40

50

60

70

80

90

100

110

120

130

140

0,01 0,1 1 10 100 1000 10000

Частота, МГц

Напря

жен

ность эл

ектр

.пол

я, дБмкВ

/мFi

eld

Inte

nsity

, dB

µV

/m

Frequency, MHz

Figure 4-9. Ambient Electromagnetic Environment at the Zenit Launch Complex (Area 45)



Launch Vehicle Radio Equipment

The Zenit-2SLB has three sets of Sirius telemetry systems which are lo-cated on the second stage and operate in a total of five frequencies. The Zenit-3SLB uses the same three sets of Sirius systems on its second stage operating in the same five frequencies, and also uses the BITC-B teleme-try system located on the Block DM-SLB third stage that operates in two additional frequencies. Each configuration has a Glonass receiving sys-tem, located on the second stage of the Zenit-2SLB and on the third stage of the Zenit-3SLB. Characteristics of these systems are provided below in Tables 4–8, 4-9 and 4-10.

Table 4-8. Characteristics of the Sirius Transmitters (Zenit-2SLB and Zenit-3SLB)

Transmitter Characteristic Meter Band Decimeter Band

Nominal Frequency (MHz) 231.3 239.3 247.3 1010.5 1018

-0.5 dB Bandwidth (MHz) + 1.3 + 1.3

Modulation Type PCM-FM PCM-FM

Max/Min Antenna Gain Coefficient (dB) 0/-7 0/-5

Output Power 11.8 – 16 dBW (15–40 W) 10 – 14.8 dBW (10–30 W)

Reduced level relative to main signal of spurious and harmonic emissions (dB)

40 30

Table 4-9. Characteristics of BITC-B Telemetry Equipment (Zenit-3SLB Only)

Characteristic Transmitter

Nominal Frequency (MHz) 1026.5 1034.5

-0.5 dB Bandwidth (MHz) 1.25

Modulation Type TBD

Max/Min Antenna Gain Coefficient (dB) TBD

Output Power (W) 17

Reduced level relative to main signal of spurious and harmonic emissions (dB) 60



Table 4-10. Characteristics of the Glonass Receiver (Zenit-2SLB and Zenit-3SLB)

Characteristic Zenit-2SLB Zenit-3SLB

Nominal Frequency (MHz) 1575.4 + 1 1575.4 + 1

-1 dB Bandwidth (MHz) 42 50

Receiver Sensitivity at Nominal Frequency (dBW) -163 -145

Max/Min Antenna Gain Coefficient (dB) 7/-3 TBD

Radio Frequency Environment at the SC Separation Plane

The maximum field intensity levels generated by launcher systems at the spacecraft interface plane are provided in Table 4-11, and Figures 4-10 and 4-11. These account for such factors as the type and orientation of antennas, and the location of the antennas relative to the spacecraft, but do not account for fairing attenuation. The fairings for the Zenit-2SLB and Zenit-3SLB are both aluminum construction, and will attenuate field levels experienced by the SC during pre-launch preparations and after launch until fairing jettison. The degree of attenuation will depend on the size and location of RF windows (mission-specific) and will be analyzed for each mission.

Table 4-11. Maximum Field Intensity Levels Generated by the Launch Vehicle at the Spacecraft Separation Plane, Without Fairing Attenuation

Field Intensity (dB µV/m) Frequency (MHz)

Zenit-2SLB Zenit-3SLB

230.0 – 232.6 150.9 128.0

238.0 –240.6 150.9 128.0

246.0 –248.6 150.9 128.0

1009.2 – 1011.8 149.6 124.1

1016.7 – 1019.3 149.6 124.1

1025.5 – 1027.5 110.9 140 (TBR)

1033.5 – 1035.5 110.9 140 (TBR)

All Other 110.9 70



20

40

60

80

100

120

140

160

180

200

Ele

ctric

fiel

d, d

B µ

V/m

00 100 1,000 100,000

Frequency, MHz

Sirius TM system 231.3 MHz239.3 MHz247.3 MHz

Sirius TM System1,010.5 MHz1,018.0 MHz

Figure 4-10. Maximum Field Intensity Levels Generated at the Spacecraft Separation Plane by the Zenit-2SLB, Without Fairing Attenuation

20

40

60

80

100

120

140

160

180

200

Ele

ctric

fiel

d, d

B µ

V/m

00 100 1,000 100,000

Frequency, MHz

Sirius TM system 231.3 MHz239.3 MHz247.3 MHz

Sirius TM System1,010.5 MHz1,018.0 MHz

BITC-B TM System 1,026.5 MHz1,034.5 MHz

Figure 4-11. Maximum Field Intensity Levels Generated at the Spacecraft Separation Plane

by the Zenit-3SLB, Without Fairing Attenuation (BITC-B field intensity levels are TBR)



4.6 Spacecraft Thermal and Humidity Environments Introduction Described in this section are the thermal and humidity conditions to

which the spacecraft will be exposed from arrival at the Baikonur airport through separation on orbit.

4.6.1 Ground Thermal and Humidity Environments General Overview, Ground Thermal and Humidity Environments

The spacecraft thermal and humidity environment is actively controlled by facility, transportation and launch pad clean air systems from the time the spacecraft container is offloaded at Baikonur airport through lift off. This supply is also maintained in the case of a launch standby or abort. Table 4-12 provides the temperature and humidity characteristics of each processing milestone or location. Figures 4-12 and 4-13 portray the air conditioning and venting schemes for the Zenit-2SLB payload unit and the Zenit-3SLB ascent unit (shown while integrated with the launch ve-hicle).

Facility Clean Air Systems

The spacecraft is exposed to ambient facility air in the PPF from the time it is unloaded from the shipping container until it is enclosed in the fair-ing. Land Launch customers may use one of two PPF’s (described in Section 6): Site 254 and Area 31. In both locations the temperature, hu-midity and cleanliness (Class 100,000 or better per FED-STD-209E) of the ambient air is actively maintained by facility clean air systems. The Launcher Assembly Complex (Area 42) also contains a clean room with active temperature, humidity and cleanliness control that is used during mating of the Ascent Unit (or Payload Unit) to the Zenit second stage. Air supply to the fairing is shut off during this mating operation. This clean room is also available for optional customer use in case physi-cal access to the spacecraft is desired through fairing access doors.

Transportation Clean Air Systems

Clean, conditioned air is supplied to the spacecraft container by a mobile clean air system while being transported between Baikonur airport (Yubileiny) and the PPF. The mobile clean air system is also used to condition the SC enclosure during transport between the PPF and fueling area at Area 31, and to condition the fairing during all moves following payload encapsulation. There may be an interruption in the supply of conditioned air for no more than 60 minutes during loading of the fully assembled launch vehicle onto the transporter/erector inside the Launcher Assembly Complex (Area 42). Cleanliness of the encapsulated environment within the fairing is always maintained at a Class 100,000 or better level per FED-STD-209E.



Launch Pad Clean Air Systems

A clean air system at the launch pad provides conditioned air to the fair-ing until the erector is lowered and removed at 12 minutes before launch. Conditioned air is maintained from T-12 minutes continuously through launch, and after T-0 in the event of an abort until the transporter/erector and its associated conditioned air system can be reattached to the launch vehicle, by a high pressure payload fairing purge system. A Uninterruptible Power Supply (UPS) system ensures that backup power is available for the pad air supply unit such that conditioned air flow to the fairing can be resumed within one minute after failure of the primary power supply.

Impingement Ve-locity of Airflow Upon SC Surfaces

Airflow impingement upon the spacecraft surfaces is generally main-tained at or below three meters per second.



Table 4-12. Spacecraft Ground Thermal and Humidity Environment

Operations Phase/Location Acting System

Temperature [0C]

Relative Humidity

Nominal Flow Rate

SC Container, Airport to PPF Transfer Mobile Unit 15 to 30* ≤ 60% 3K-6K m3/h

PPF Processing Areas

Area 31 Area B Area 31 Room 119 Site 254

Facility Air Conditioning

15 to 28 17 to 23 18 to 25

35% - 60% 40% - 60% 30% - 60%

N/A

Transfer from PPF to HPF (Area 31) Mobile Unit 15 to 30* ≤ 60% 3K-6K m3/h

PPF Fueling Cells Area 31 Site 254


15 to 25 TBD

30% - 60% TBD N/A

PPF Encapsulation Halls

Area 31 Site 254


15 to 28 18 to 25

35% - 60% 30% - 60% N/A

Transfer to ILV Integration Area (Site 42) Mobile Unit 10 to 35* 30% - 60% 3K-6K m3/h

ILV Integration Bay (Site 42) Facility Air Conditioning 18 to 25 ≤ 80% N/A

Clean Room (Site 42) Facility Air Conditioning 21 to 26.7 30% - 60% N/A

ILV ready in Site 42 for roll-out Mobile Unit 10 to 35* < 60% ≤ 3K m3/h ILV transfer to launch complex (and from launch complex after launch abort) Mobile Unit 10 to 35* < 60% > 2250 kg/h

ILV erection, and while erect prior to LOX loading ILV de-erection if launch aborted before T-12 min

Zenit-2SLB Zenit-3SLB Pad System

10 to 35* 8 to 25*

DP ≤ -100C DP ≤ -100C

9500 m3/h 5000 m3/h

ILV erect on launch pad, from LOX loading until T-12 min

Zenit-2SLB Zenit-3SLB Pad System 10 to 35*

8 to 25* DP ≤ -300C DP ≤ -300C

9500 m3/h 5000 m3/h

ILV erect on launch pad, T-12 min to T-0 ILV erect on launch pad following launch aborted between T-12 and T-0, through ILV de-tanking and de-erection (until mo-bile unit is reconnected)

High Pressure Pad System 10 to 32* DP ≤ -550C > 2250 kg/h

Notes: • Temperatures maintained within + 2 0C of set point agreed with the customer • * Denotes temperatures as measured at the fairing (or SC container) inlet • The customer is responsible for monitoring the environment inside the SC container • DP = dew point, K = thousands



SC

Second Stage A/C Inlet(Up to 3600 m3/hr)

Fairing Vents

SeparatingScreen

Fairing A/C Inlets(to T-12 minutes)Up to 9500 m3/hr

Second Stage Vents

Fairing A/C InletHigh Pressure Pad System (from T-12 minutes)Up to 2250 kg/hr

Figure 4-12. Zenit-2SLB Ascent Unit Air-Conditioning (A/C) and Venting Scheme

Second Stage A/C Inlet(Up to 3600 m3/hr)

Third Stage A/C Inlet(Up to 4500 m3/hr)Fairing A/C Inlet

(Up to 5000 m3/hr)

Third Stage Vents

Second Stage Vents

Figure 4-13. Zenit-3SLB Ascent Unit Air-Conditioning (A/C) and Venting Scheme



4.6.2 Flight Thermal Environments General Overview, Flight Thermal Environment

After launch, the spacecraft will experience: • Heat flux radiated from internal surfaces of the fairing, before fairing

jettison. This is mitigated by insulation and ablative coatings on the fairings.

• After fairing jettison, free molecular heating and various other ther-mal influences. This is mitigated by the late timing of fairing jettison, by the short duration of the Zenit-2SLB mission, and on Zenit-3SLB by the thermal maneuvering capabilities of the Block DM-SLB

Thermal effects experienced by the spacecraft during flight are summa-rized in Table 4-13. A thermal analysis will be performed for each mis-sion, using the spacecraft thermal model provided by the customer, to assess spacecraft temperatures during all mission phases.

Table 4-13. Flight Thermal Environments

Thermal Effect Zenit-2SLB Zenit-3SLB Thermal flux radiated onto the spacecraft from fairing internal surfaces

500 W/m2 maximum 400 W/m2 maximum

Free molecular heating at fairing jettison

1135 W/m2 or less. Considerably less than 1135 W/m2 for most missions (typically around 50 W/m2) due to drop zone requirements

Free molecular heating after fairing jettison

Typically not significant due to short mission duration. Analyzed for each mission.

Dependent on mission profile, but may spike slightly at the time of the second Block DM-SLB burn as shown in Figure 4-14

Heat radiated onto space-craft surfaces by the sec-ond stage solid propellant separation thrusters

9.0 Kw-s/m2 maximum 5.1 Kw-s/m2 maximum

Solar heating, planet-reflected solar heating (al-bedo), Earth-radiated heat-ing, radiation to space

Typically not significant due to short mission duration. Analyzed for each mission.

Analyzed for each mission. For thermal management, the Block DM-SLB is designed to accommo-date preferred attitude pointing, continuous rolls, maneuvers, and orientations during coast and pre-separation phases of flight



0

50

100

150

200

250

300

0 2000 4000 6000 8000 10000 12000 14000

Time from lift-off, seconds

Dis

pers

ed fr

ee m

olec

ular

flux

, W/m

2

MEOGTO/GSO

Figure 4-14. Zenit-3SLB Free Molecular Heating Environment



4.7 Pressure Venting Overview During ascent, the payload volume is vented through a set of orifices in

the second stage equipment bay and in the third stage, as shown in Fig-ures 4-15 and 4-16.

Pressure Decay Rate

The depressurization rate, though varying somewhat by trajectory and dependent on spacecraft displaced volume, does not exceed: 0.028 kgf/cm2 per second for Zenit-2SLB 0.032 kgf/cm2 per second for Zenit-3SLB

A typical fairing cavity pressure curve for the Zenit-2SLB and Zenit-3SLB are provided in Figures 4-17 and 4-18, along with the associated pressure decay profiles shown in Figures 4-19 and 4-20. The specific predicted pressure venting rate for each launch is determined during the mission analysis phase.

Pressure Differential at Fairing Jettison

Due to the late timing of fairing jettison due to drop zone constraints, the maximum pressure differential between the pressure inside the fairing and the external pressure at fairing jettison does not exceed a very low 0.002 kgf/cm2 for both Zenit-2SLB and Zenit-3SLB.



SeparatingScreen

Second Stage Vents

Figure 4-15. Zenit-2SLB Ascent Venting Scheme

Third Stage Vents

Second Stage Vents

Figure 4-16. Zenit-3SLB Ascent Venting Scheme



-- - - Ambient Pressure-- - - Ambient Pressure-- - - Ambient Pressure-- - - Ambient Pressure

Pres

sure

, kgf

/cm

2

Time From Launch, seconds

------ Ambient Pressure

Predicted Internal Pressure

Figure 4-17. Typical Zenit-2SLB Fairing Internal Pressure Profile During Ascent

To be provided

Figure 4-18. Typical Zenit-3SLB Fairing Internal Pressure Profile During Ascent



Fairing Cavity Pressure, kgf/cm2

Pres

sure

Cha

nge

Rat

e, k

gf/c

m2

per

seco

nd

Figure 4-19. Typical Zenit-2SLB Fairing Internal Pressure Decay Profile

To be provided

Figure 4-20. Typical Zenit-3SLB Fairing Internal Pressure Decay Profile



4.8 Contamination Contamination Control During Ground Process-ing

The spacecraft is protected from contamination during ground processing by:

• Supplying a continuous flow of clean, conditioned air (class 5,000 or better per FED-STD-209E) to the SC while in its container or under the LV fairings, through launch and after T-0 in the event of a launch scrub or abort. These clean air systems, both mobile and fixed pad units, are described in more detail in Section 4.6.1 and maintain a constant overpressure inside the enclosure relative to ambient to pre-vent outside air ingress.

• Providing Class 100,000 or better per FED-STD-209E clean room facilities for all spacecraft operations (unloading, processing, fueling, encapsulation) between removal from the SC container and encapsu-lation in the LV fairing.

• Precision cleaning of the launch vehicle hardware surfaces that en-close the spacecraft, prior to placing them in proximity to the space-craft. These cleanliness levels are described in Table 4-14.

Table 4-14. Fairing Internal Surface Cleanliness Levels at Encapsulation

Particles Maximum Fairing Surface Levels Particle

Size Level 500

per Mil-Std-1246C Zenit-2SLB Zenit-3SLB Level 750

per Mil-Std-1246C >100 µm 11,900/m2 11,900/m2 30,129/m2 96,300/m2 >250 µm 281/m2 281/m2 753/m2 2,310/m2 >500 µm 10.8/m2 10.8/m2 32/m2 87.5/m2

Non-Volatile Residue

Zenit-2SLB 10 mg/m2 (Level A per Mil-Std-1246C) Zenit-3SLB TBD



Contamination Control During Flight

Potential sources of the SC contamination in flight are fairing materials, contaminants migrating from the launch vehicle equipment bays while venting during ascent, and plume impingement from the second stage solid propellant separation thrusters and, on Zenit-3SLB missions, from the Block DM-SLB steering engines and outgassing. All of these sources are addressed in the design of the Land Launch hardware and mission.

Fairing Design Features to Minimize Contamination

Materials exposed to the cavity shared by the spacecraft are selected to preclude crumbling, peeling, particle shedding, oxidation or corrosion, and with low outgassing properties that should not exceed the following values during testing according to GOST R 50109-92 (equivalent to ASTM E-595):

• Total mass loss (TML) less than 1%

• Collected Volatile Condensable Material (CVCM) less than 0.1%

Pyrotechnic devices used for fairing and satellite separation are sealed and do not release gasses or particles. The venting system is designed to preclude circulation from the upper stage (Zenit-3SLB) or second stage avionics bay (Zenit-2SLB) back into the payload unit.

Plume Impingement

The plume effect on the SC of the second stage solid propellant separa-tion thrusters is negligibly small because of their location on the aft end of the stage. On three stage missions, the Block DM-SLB performs a Contamination and Collision Avoidance Maneuver (CCAM) following satellite separation to ensure a negligibly small contaminating effect on the SC from the upper stage steering engines and stage venting. CCAM features include stage attitude control for optimum orientation relative to the SC, minimum separation distance before main engine re-start, long burns for orbit separation and fuel depletion, and fuel tank venting as a final act. The Block DM family has performed CCAM hundreds of time.


Rev. 01/2004 HPD-19000 5-1

5. SPACECERAFT INTERFACES 5.1 Mechanical Interfaces

Overview Mechanical interfaces covered in this section include:

• mass properties and modal frequencies

• fairing volumes, access doors and RF windows

• spacecraft adapters

5.1.1 Mass Properties and Modal Frequencies

Spacecraft Mass and Longitudinal Center-of Gravity Location

Table 5-1 presents the allowable spacecraft mass and longitudinal center-of- gravity (CG) limits for the Zenit-3SLB, based on the standard spacecraft adapters (SCA) currently offered by Land Launch. Customers with spacecraft having a mass / CG configuration that exceeds the defined envelopes should still consult Land Launch for a more detailed assessment.

Table 5-2 presents the analogous mass / CG envelope data for the Zenit-2SLB.

Table 5-1. Expected Spacecraft Mass and CG Limits – Zenit-3SLB

Maximum Spacecraft CG (meters) 937 SCA 1194 SCA 1666 SCA

Band Tension (kN) Band Tension (kN) Band Tension SC Mass

(kg) 12.5 20 30 10 20 30 40 30 kN

1500 2.15 2.50 2.50 2.15 2.50 2.50 2.50 2.50 2000 1.60 2.50 2.50 1.60 2.50 2.50 2.50 2.50 2500 1.30 2.10 2.50 1.30 2.50 2.50 2.50 2.50 3000 1.10 1.75 2.50 1.10 2.20 2.50 2.50 2.50 3500 0.95 1.50 2.25 0.95 1.85 2.50 2.50 2.50 4000 0.80 1.30 1.95 0.80 1.60 2.45 2.50 2.50 4500 0.70 1.15 1.75 0.70 1.45 2.15 2.50 2.40 5000 0.65 1.05 1.55 0.65 1.30 1.95 2.50 2.15

Notes and assumptions: • Mass refers to the separated mass of the spacecraft • CG refers to longitudinal distance forward of the separation plane • Lateral load factor equal to 2 g



Table 5-2. Expected Spacecraft Mass and CG Limits – Zenit-2SLB

Maximum Spacecraft CG (meters) SC Mass (kg)2624-mm SCA Zenit-2 Adapter

2500 8.00 2.52 3000 8.00 2.10 4000 8.00 1.58 5000 8.00 1.26 6000 8.00 - 7000 8.00 - 8000 8.00 - 9000 7.20 - 10000 6.50 - 11000 5.80 - 12000 5.30 - 13000 5.00 - 14000 4.60 - 15000 4.30 -

Notes and assumptions: • Mass refers to the separated mass of the spacecraft • CG refers to longitudinal distance forward of the separation plane • Lateral load factor equal to 2 g.

Spacecraft Center-of Gravity Radial Offset

The radial offset of the spacecraft CG, relative to the launch vehicle longitudinal centerline, should not exceed:

• 50-mm (Zenit-2SLB)

• 25-mm (Zenit-3SLB)

Exceptions should be brought to Land Launch for assessment.

Modal Frequencies

Table 5-3 presents guidelines for spacecraft fundamental natural frequencies on Land Launch. Exceptions should be brought to Land Launch for assessment.

Table 5-3. Recommended Spacecraft Fundamental Frequencies

Spacecraft Fundamental Frequencies Launch SystemLongitudinal Lateral

Zenit-2SLB > 15 Hz > 5 Hz Zenit-3SLB > 20 Hz > 8 Hz


Rev. 01/2004 HPD-19000 5-3

5.1.2 Payload Fairing Mechanical Interfaces

Payload Fairings Land Launch offers a 4.1 meter (outer diameter) fairing for the three-stage Zenit-3SLB vehicle (Figure 5-1) and a 3.9-meter fairing (outer diameter) for the two-stage Zenit-2SLB (Figure 5-2) vehicle. Each fairing has a demonstrated flight record. The general lay-outs of the two fairings are shown in Figures 5-3 and 5-4.

RSC Energia Photo

Figure 5-1. The Zenit-3SLB Uses the Flight-Proven 17S72 Fairing Made by NPO Lavochkin (shown attached to the Block DM)



SDO Yuznoye Photo

Figure 5-2. The Zenit-2SLB Uses a Flight-Proven Fairing Made by PO Yuzhmash


Rev. 01/2004 HPD-19000 5-5

Longitudinal SeparationJoint

Transfer Compartment

Fairing Field Joint

Spacecraft Adapter

Figure 5-3. General Lay-out of the 4.1-meter 17S72 Fairing Used on the Zenit-3SLB



Intersection Bay

Figure 5-4. General Lay-out of the 3.9-meter Fairing Used on the Zenit-2SLB


Rev. 01/2004 HPD-19000 5-7

Useable Volume Figures 5-5 and 5-6 present the spacecraft static envelopes for the Zenit-2SLB and Zenit-3SLB fairings. These envelopes account for:

• worst-case fairing manufacturing and assembly tolerances

• in-flight dynamic displacement of the fairing

• in-flight dynamic displacement of the spacecraft, conservatively assumed to be 50-mm (varies with spacecraft fundamental frequencies and subject to verification during coupled loads analysis)

These envelopes should not be exceeded by the maximum dimensions of the spacecraft, including worst case tolerances and expanded thermal blankets, under static conditions (one g longitudinal, zero g lateral). Local excursions outside these envelopes can sometimes be accommodated. Customers are encouraged to contact Land Launch for a specific assessment.

3620 975 (2624 interface)830 (Zenit-2 frame)

Spacecraft interface plane

Figure 5-5. Spacecraft Static Envelope within the Zenit-2SLB Fairing (dimensions in millimeters)



SC envelope

ø3633ø3723 ø3645

ø2103ø207

2ø3

679

ø357

7ø3

577

ø3645

600

504 (937 interface)369 (1194 interface400 (1666 interface)

ø359

5

All dimensions in millimeters

Transfer Compartment

Figure 5-6. Spacecraft Static Envelope within the Zenit-3SLB Fairing


Rev. 01/2004 HPD-19000 5-9

Useable Volume Inside Payload Structure

There is room inside the payload support structure (spacecraft adapter and transfer compartment) for satellite motor nozzles. Dimensions will be provided in a future version of this User’s Guide.

Access Doors Up to two access doors are standard on either Land Launch payload fairing, one door per fairing half. Nominally, the maximum size of each individual door can be:

• 420-mm by 420-mm on the Zenit-3SLB fairing

• 500-mm by 500-mm on the Zenit-2SLB fairing Allowable locations for the access doors are shown in figures 5-7 and 5-8. Access doors can typically be used until shortly before the ascent unit (Zenit-3SLB) or payload unit (Zenit-2SLB) leaves the PPF for the launch vehicle assembly building (Area 42), though special arrangements can also be made to open the access doors within the clean room inside Area 42.

RF Windows Both Land Launch fairings can accommodate RF windows to enable the testing of spacecraft transmitters on the launch pad. Sizes and locations will be coordinated between the customer and the Land Launch team.

Customer Insignia

There is external space on either Land Launch fairing for a customer logo or insignia. Details will be coordinated between the customer and the Land Launch team



Location of access doors to SC

Location of accessdoors to SC

Location of accessdoors to SC

Fig. 5-7. Zenit-3SLB Payload Fairing Locations for Access Doors



I II III IV I

I II III IV I6°

74°

1215

55°

Развертка цилиндрической части головного обтекателя

Зона возможного размещения люков доступа

Плоскость стыка ГО с переходным отсеком

925 40

40

6°

560

80

2495

2495

2495

I II III IV I

I II III IV I6°

74°

1215

55°

Развертка цилиндрической части головного обтекателя

Зона возможного размещения люков доступа

Плоскость стыка ГО с переходным отсеком

925 40

40

6°

560

80

2495

2495

2495

Unrolled cylinder of the Zenit-2SLB fairing

Crosshatch = Access door allowable locations

Interface plane between the fairing and the intersection bay

Figure 5-8. Zenit-2SLB Payload Fairing Locations for Access Doors

(dimensions in millimeters)

5.1.3 Spacecraft Adapters Saab Spacecraft Adapters

A full range of spacecraft adapters (SCA) made by Saab Ericsson Space is available from Land Launch. Standard offerings are shown in Table 5-4. Other adapters may also be used. Customers interested in other adapters are encouraged to contact Boeing Launch Services.

Each standard Land Launch SCA features:

• a clamp band separation system with a 100% successful demonstrated flight history

• push-off springs to impart an initial delta-velocity to the spacecraft with respect to the launch vehicle

• two electrical disconnects at the spacecraft interface



Table 5-4. Standard Saab Ericsson Space Spacecraft Adapters Available on Land Launch

SCA 937

Mass = 49 kg*

SCA 1194

Mass = 58 kg*

SCA 1666

Mass = 60 kg*

SCA 2624 (Usually Zenit-

2SLB only)

Mass = 148 kg

* On the Zenit-3SLB, Payload Systems Mass (PSM) also includes the mass of the Transfer Compartment (65 kg) and mission-unique harnessing.



Zenit-2 Adapter (Frame) for Use with Zenit-2SLB

The Zenit-2 frame adapter has been successfully used on a number of missions. The spacecraft is attached to four pyro-locks on a 2062-mm diameter. There are two electrical disconnects at the SC interface. Relative velocity between the separating spacecraft and the launch vehicle is imparted by the simultaneous firing of solid-propellant separation thrusters located on the aft end of the second stage.

Multi-Satellite Dispensers for Use with Zenit-2SLB

Features of the Zenit-2SLB that make it inherently well suited for launching groups of satellites, as well as secondary payloads, include:

• Heavy-lift LEO performance (see section 3.5)

• Large fairing (see section 5.1.2)

• A structurally robust second stage structure that is designed to support a very large mass (on Sea Launch missions it routinely carries more than 25 tonnes of combined third stage, payload and fairing mass)

Land Launch partners PO Yuzhnoye and PO Yuzhmash have extensive experience in designing and building unique, cost-effective payload structures for accommodating and separating multiple payloads. Interested customers are encouraged to contact Boeing Launch Services.



5.2 Electrical Interfaces

Overview Electrical interfaces that are covered in the following sections include:

• hard-line links (umbilical)

• radio frequency link

• in-flight commands, measurements and telemetry

• electrical power for EGSE

• bonding and grounding

5.2.1 Hard Line Links (Spacecraft Umbilical)

Umbilical Circuits Circuits capable of simultaneously supporting the functions defined in Table 5-5 are provided between the spacecraft and hook-up locations for the customer’s electrical ground support equipment (EGSE).

Table 5-5. Spacecraft Umbilical Links

Function Circuits Capability or Number & Type

Signal Up to 20 twisted shielded pairs, each at 250 mA at 50 V ac or 100 V dc

Power (External Spacecraft Power) 20 A at 110 V dc

Power (Battery Charging)

20 A at 110 V dc (from T-12 hours until T-0) 70 A at 110 V dc (all other times)

The umbilical circuits defined above are available for customer use at each stop in the launch processing flow as shown in Table 5-6.

Umbilical Use During Processing and Launch

Table 5-6. Umbilical Hook-Up Locations and Availability

Place Location for Connecting to Customer-Provided EGSE Notes

PPF Processing Cell Processing Cell Control Room Fuel Cell Control Room

After attaching the SC to the adapter

PPF Encapsulation Hall After attaching the SC and adapter to the upper stage (Zenit-3SLB)

Site 42 Clean Room and EGSE Room

Launch Complex

Bunker (power circuits) Customer Control Room (signals circuits)

While erect on the pad through launch, and after T-0 until lowered in the case of a launch abort



Umbilical Connectors

Umbilical circuits are connected to the spacecraft by separation connectors on the spacecraft adapter. Standard Land Launch spacecraft adapters each accommodate two 61-pin connectors or two 37-pin separation connectors. Other required connectors can be accommodated and will be negotiated on a case-by-case basis. Whichever separation connectors are selected, the customer or spacecraft contractor should procure them and provide the mating halves for the launch vehicle contractor to incorporate into the umbilical and the adapter.

Details of umbilical connector interfaces for customer EGSE will be coordinated in the ICD.

5.2.2 Radio Frequency Links

Direct RF communications in C-Band, Ku-Band and/or K-Band between the spacecraft in the fairing on the erect launch vehicle at the pad, and customer EGSE located in the bunker, are provided by: • RF window(s) in the fairing • A bunker roof antenna, and RF channel equipment linking the bunker

roof antenna to connectors in the bunker for customer EGSE The timing of transmissions will be coordinated by Land Launch with the Range authorities. Frequencies, RF window location, and connectors will be coordinated and documented in the ICD.

5.2.3 In-Flight Commands, Measurements and Telemetry General The launch vehicle will provide the signals and the power to initiate the

satellite separation system at the pre-determined point in the mission. No launch vehicle power or command lines will cross the spacecraft separation plane.

Separation Verification

The launch vehicle will detect spacecraft separation via two diametrically opposed separation switches at the adapter/spacecraft interface. Indication of spacecraft separation will be telemetered via redundant channels.



Satellite Environment Measurements

Lateral and longitudinal accelerations will be recorded near the interface between the spacecraft and the adapter, telemetered and received. Also recorded, telemetered and received will be information sufficient to determine the acoustics, fairing internal surface temperature, pressure decay, low frequency vibrations, high frequency vibrations and shock environments to which the satellite is exposed during launch. The launch vehicle is capable of issuing up to 2 primary and 2 redundant commands to the spacecraft during flight (in addition to commands for lift-off contact, fairing jettison, readiness for spacecraft separation and spacecraft separation). Command characteristics are as shown in Table 5-7:

Commands

Table 5-7. Characteristics of Commands from Launch Vehicle to Spacecraft

Command Feature Property Timing of Actuation Any time between lift off and satellite separation Accuracy of Actuation Timing + 35 ms

Signal Duration 64 + 8 ms Current 0.01 to 1.0 Amperes Voltage Up to 34 V



5.2.4 Electrical Power for EGSE

Ground Power Electrical power is provided to the customer at each fixed location in the processing flow, as shown in table 5-8.

Table 5-8. Electrical Power Supplies for Customer EGSE

220 V +10%/-5% 50 Hz + 2%

380 V+ 2% 50 Hz + 1%

3 Phase

120 V, 20 A, single phase

60 Hz PPF Processing Cell 80 kW 60 kW TBD

PPF Fueling Area 80 kW 60 kW TBD

Launcher Assembly Bldg 40 kW 40 kW TBD

Launch Complex Bunker 20 kW 20 kW TBD

Launch Complex Control Room 20 kW 20 kW TBD

Uninterruptible Back-up Power

Customer EGSE power is backed-up with a dedicated Uninterruptible Power System (UPS). Characteristics are noted in table 5-9. This system is fully redundant (two UPS units in tandem), each with 15 minutes of battery capability at rated full load. Both 50 Hz and 60 Hz power can be provided, but not at the same time. The UPS supplies three-phase power with a neutral which can be provided to the customer in either single or three phase receptacles.

Table 5-9. Uninterruptible Power Supply for Customer EGSE Frequency Characteristics

50 Hz 220/380 VAC, three phase "Y" at 40 KVA with power factor from unity to 80% lagging

60 Hz 120/208 VAC, three phase "Y" at 40 KVA with power factor from unity to 80% lagging



5.2.5 Bonding and Grounding

Bonding The bonding resistance across the spacecraft to adapter interface is less than 10 milliohms at a current less than 10 milliamps. The resistance between the adapter and the launch vehicle is less than 2 milliohms at a current less than 10 milliamps. Before launch, the launch vehicle is connected to ground with resistance less than 100 milliohms at a current less than 200 milliamps.

Grounding At each Land Launch facility where the spacecraft is located or customer EGSE is used during the processing flow, electrically conductive surfaces (threaded studs) are provided for connecting to facility ground.



6. LAND LAUNCH FACILITIES

Overview Land Launch has the advantage of using proven and established facilities at Baikonur Cosmodrome. These include:

• Krainy Airport for launch personnel arrival and departure • Yubileiny Airport for spacecraft and ground support equipment shipment • Two payload processing facilities: Site 31 and Site 254. Site 31 will be the

primary processing facility until Site 254 facilities improvements are completed. The Block DM-SLB is processed in Site 254.

• Zenit technical complex located at Site 42 for Zenit processing and mating of the Ascent Unit (Zenit-3SLB) or Payload Unit (PLU) with the Zenit stages followed by check-out of the integrated launch vehicle (ILV).

• Zenit launch complex located at Site 45 for launching the Zenit-3SLB ILV and the Zenit-2SLB ILV.

A map of the Land Launch Baikonur facilities is shown in figure 6-1

90 km

Syrdar-ya River

RailroadRoad

75 km

Yubileinyi airfield

Area 254Payload and Block DM

Processing

City of Baikonur

Legend

N

Site 45Zenit launch complex

Site 31 Payload Processing

Krainy Airport

Site 42 Zenit and ILVProcessing

Figure 6-1. Location of the Principal Land Launch Facilities at Baikonur Cosmodrome



6.1 Transportation of Personnel and Cargo to and from Baikonur

Krainy Airport

Personnel fly between Moscow and Baikonur via Krainy Airport (Figure 6-2), which is situated six kilometers to the west of Baikonur city. It can accommodate midsize aircraft for passenger travel throughout the year. Flights are available to and from Vnukovo-1 or Vnukovo-3 airports in Moscow on both commercially scheduled and dedicated charter flights. Land Launch will assist customer personnel in obtaining visas through the Federal Space Agency, and will provide customer representatives with access to the Cosmodrome as well as badges to the required facilities.

Figure 6-2. Krainy Airport at Baikonur

Yubileiny Airport

Yubileiny Airport is located 45 km north of Baikonur city within Baikonur Cosmodrome and is operated by Rosaviacosmos. Its runway, which is 4,500 meters long and 84 meters wide and conforms to International Civil Aviation Organization (ICAO) standards for Class 1 airports, was built to accommodate the landings of the Buran space shuttle. It handles aircraft of all classes for both freight and charter flights, including Boeing 747s and Antonov 124s. Commercial launch customers have used it many times for delivering spacecraft and associated support equipment. The airfield can operate year-round at any time of day. A typical off-load is shown in Figure 6-3. Upon arrival of aircraft, the SC container and associated equipment are offloaded from the aircraft and transferred to railcars that are located approximately 50 to 80 meters from the aircraft. Cranes, forklifts and other necessary equipment are available for these operations. The airport is connected by rail and road to all major cosmodrome facilities.



Figure 6-3. Spacecraft Off Load at Yubileiny Airport

Transportation at the Cosmodrome

Rail and road networks connect all Land Launch facilities at Baikonur. Land Launch provides the customer with all necessary transportation of equipment and people on base. Generally, equipment will move between facilities by rail while people will move by road. The spacecraft makes three major moves between facilities: from Yubileiny to the PPF, from the PPF to the launcher assembly building (Area 42) and from Area 42 to the launch complex (Area 45). Spacecraft moves are conducted by rail (Figure 6-4), inside protected enclosures (its own shipping container for the first move, and the fairing for the second and third moves) that are continuously purged with clean, conditioned air as described in Section 4.



RSC Energia Photo

Figure 6-4. Ascent Unit Transportation with Thermostating Car

6.2 Site 31 Payload Processing Facility Overview Launch Launch’s primary Payload Processing Facility (PPF) consists of the

existing Site 31 complex of buildings and facilities, which has been used previously to process numerous Western and CIS payloads. Site 254 will become the primary PPF when facility upgrades and improvements are completed. All spacecraft processing, propellant filling operations, pressurization, ordnance preparation, and payload fairing encapsulation operations are conducted here. The PPF has controlled access to ensure compliance with United States governmental security regulations as well as self-imposed customer security requirements and procedures. Major PPF features include:

• spacecraft processing areas • spacecraft fueling area • fuel storage room • oxidizer storage room • control rooms for spacecraft ground support equipment • garment change rooms with personnel airlock • an encapsulation area • office areas for spacecraft personnel

The layout of the principal buildings at Area 31 is shown in Figure 6-5.



Storage

Lavochkin NPO Storages

40Е63 380

46 45

44120

122125

105

87А 43

48Б

57

48 48А 87

4040Д

40А

124 51

40 Assembly & Test Building (ATB) 51 Support building 40А ATB Annex (Vacuum Chamber) 57 Boiler Facility 40Д ATB Annex (Clean Area) 63 Receiver Facility 40Е Ventilation Facility 87А Uninterruptable Power Supply Facility 43 Charge-Storage Battery Station 105 Transformation Station (6/0.4 kV) 44 Fueling Area 120 IAE Storage 45 Oxidizer storage 122 Refrigerating Center 46 Fuel storage 124 Laboratory Building 48 Cooling Tower 125 Cooling Tower

48А Water Recycling Pump Station 380 Electro-Diesel Station (mobile, 200 kW) 48Б Water Tank 87 Workrooms

Figure 6-5. Area 31 Partial Facility Lay-Out



Building 40/40D, PPF

Buildings 40/40D at Area 31 are used for non-hazardous payload processing. Building 40 has three principal work areas, Area A, B and C, that are shown in Figure 6-6. The SC and equipment are brought into Area C from the airport, unloaded and transitioned into building 40D, room 119 for processing. Rooms 119, 119A and 119B in Building 40D are the usual locations for SC processing and check out prior to fuelling, and are shown in Figure 6-7. After the SC is fueled in Building 44 (see below), it is returned to Building 40, Area A for the beginning of joint operations. In Area A, the SC will be mated to the spacecraft adapter and then to the Block DM in the Zenit-3SLB configuration. This unit will then be rotated to horizontal and encapsulated. For the Zenit-2SLB configuration, the SC will be mated to the spacecraft adapter and intersection bay, and then rotated to the horizontal position for encapsulation.

Building 40D Office areas,

Air-conditioned office facilities are provided at Site 31. These facilities provide private office and conference space for resident spacecraft personnel teams, including separate office space for the spacecraft manufacturer and satellite customer. International data and voice communications circuits are available.

Building 44, HPF

The SC is fueled in the Hazardous Processing Facility, Building 44, located about 300 meters from Building 40/40D. Transfer of the SC back and forth is accomplished inside a conditioned container. A layout of Building 44 is shown in Figure 6-8. Key features of Building 44 include:

• Clean tent preserving Class 100,000 conditions for the SC • Control room with blast-hardened bay window overlooking the fuel

island and clean tent • Fueling island with spill containment system, hazardous vapor

monitors and emergency egress doors • Communications, fire-fighting and emergency egress systems • Supplies of clean water, liquid nitrogen and facility air • Breathing air systems for SCAPE • Changing rooms


-

Figure 6-6. Lay-out of Buildings 40 and 40D

Building 40

Sliding gate8.4 х 10 (h)

Area - С

N

Clean Rooms

Area - B

Area - A

Building 40Д

Sanitary inspaction room Building 40А

Bridge Crane 50/10 tf (2 ps)Н(hook)= 13.74/14.63 m

Bridge Crane5 tf Н=8.1 m

Sliding gate5.5 х 7 (h)


Encapsulation BayArea – 300 sq mHeight – 18.35 mClass 100,000 cleanCrane 1: 5 t, 16.5 mCrane 2: 10 t, 14.6 mCrane 3: 50 t, 13.7 m

SC Processing RoomArea – 240 sq mHeight – 10.8 mClass 10,000 cleanCrane: 5 t, 8.1 m

SC Processing RoomArea – 300 sq mHeight – 18.35 mClass 100,000 cleanCrane 1: 10 t, 14.6 mCrane 2: 50 t, 13.7 m

Figure 6-7. SC Processing and Joint Operations Area in Buildings 40 and 40D


-

1 – Filling Hall 3 – Oxidizer Loading Area 5 – Air Lock 2 – Pressurizing Hall 4 – Changing Room 6 – Shower/Medical Station

Figure 6-8. Hazardous Processing Facility, Building 44, at Site 31



6.3 Site 254 Payload Processing Facility

Overview Site 254 will become the primary spacecraft processing facility for Land Launch after various upgrades and improvements have been completed. The upgrades include an additional processing/fuelling cell adjacent to an existing building. A layout of the existing PPF with the proposed processing cell is shown in Figure 6-9.

Encapsulation Area

Area 102

Block DMProcessing

Uninterruptible Power Supply

Sumps for Spill/ Run-Off Containment

Loading/Unloading Area Container Cleaning and Acceptance

SC Processing/Fueling Area

Customer Offices Area 101

Figure 6-9. Lay-out of SC PPF at Site 254 with proposed adjacent building

Site 254 PPF layout

The main areas of the PPF for the SC are 101, 102 and the proposed new processing cell. Upon arrival from Yubileini, the SC container and equipment are off-loaded in Area 101 that is located in the central bay of the PPF. Cleaning and acceptance of the cargo is performed in Area 101. The proposed new processing area is located adjacent to Area 102. All SC autonomous operations are performed in this cell. Integrated operations occur in Area 102 as illustrated in Figure 6-10.

Site 254 The PPF is equipped with systems to support all SC processing. Major PPF


-


PPF features technical systems include: • Power supply, 380/220 V, 50 Hz; 280/120 V, 60 Hz • Compressed gases (air, nitrogen,helium) • Conditioned air • SC processing area • SC fueling area, including remote control room • SC storage room • Oxidizer storage room • Control rooms for spacecraft ground support equipment • Garment change rooms with personnel airlock • Encapsulation area • Office areas for spacecraft personnel

RSC Energia Photo

Figure 6-10. Encapsulation Operations in Site 254 Room 102



6.4 Zenit Technical Complex Site 42 Overview The Zenit-TM technical complex located within Site 42, which includes

the launch vehicle assembly and testing Building 41 (Figure 6-11), is used for:

• Standalone integration and testing of the Zenit stages • Mating of the Zenit with the Ascent Unit (Zenit-3SLB) or with the

Payload Unit (Zenit-2SLB), to form the Integrated Launch Vehicle (ILV)

• Integrated ILV testing • ILV loading onto the transporter/erector, prior to moving to the

launch complex for launch The complex also includes office space for customer personnel, an equipment room, and a clean room.

3P111073P11107

SDO Yuzhnoye Photo

Figure 6-11. North Rail at the Zenit Technical Complex, Site 42


-


Integration Area Layout/Features

Building 41 is 120 meters long and 60 meters wide, with three parallel sets of floor-mounted rails. The center rails are used for hardware delivery into and out of the building. The rails on the north side are used for launch vehicle integration operations, while the south rails are currently used for hardware storage. Two traveling bridges each have two cranes, with 50-tonne and 10-tonne capacities. Straddling the north side rail is the clean room (Figure 6-12) that is used for mating the Ascent Unit/PLU to the Zenit second stage. The environmental parameters of this clean room are defined in Section 4. While the fairing is in the clean room the customer has the option of accessing the spacecraft through doors in the fairing. Stands and ladders are available if required.

SDO Yuzhnoye Photo

Figure 6-12. Clean Room at Area 42

Spacecraft Equipment Room

An equipment room is available for customer use in Building 41, equipped with the power supplies and the umbilical connections to the spacecraft that are defined in Section 5.



Customer Office Areas

Air-conditioned customer office facilities are provided at in Building 41, Site 42. These facilities provide private office and conference space for resident spacecraft personnel teams. The customer is provided with local and international telephone communication, internal technological communication, broadcasting communication, access to data transmission channels within Baikonur cosmodrome as well as to the international communication channels from Site 42.

6.5 Zenit Launch Complex (LC) – Site 45 Overview A general lay-out of the Zenit launch complex is shown in Figure 6-13. It

consists of two adjacent launch pads supported by shared infrastructure, including propellant tank farms, bunkered launch control complex, and control equipment. Land Launch employs the operational #1 launch pad for both Zenit-3SLB and –2SLB missions. Many features are nearly identical to the ones found on Sea Launch, including launch pad, auto-coupling and fueling systems, the transporter/erector and the control system.

Launch Complex Automated Systems

Launch operations are highly automated on Land Launch just as on Sea Launch. This has many advantages including:

• short time spent on the pad (approximately 28 hours, unless the customer needs more time for spacecraft testing)

• inherent safety to personnel, since there is no need to physically approach the launch vehicle

• high launch-on-time probability

If the launch process does experience an anomaly requiring termination, it does so automatically, assuring safety of the launch vehicle, spacecraft, and launch complex. If they are needed, launch vehicle de-fueling operations are also implemented remotely from the control post.


-


6

10N

4

3

2

8

5A 59

1

7

1 Launch pad #1 2 Launch pad #2 (not in current use) 3 Launch control block

(command center) 4 Equipment Bunker 5 Launcher storage bunkers (not in current use) 6 Kerosene storage area 7 Oxidizer storage area 8 Pressure bottle storage 9 Compressor station 10 Air conditioning plant

Land Launch

Rail Connection to Site 42

Figure 6-13. Lay-out of the Zenit Launch Complex, Area 45



Customer EGSE Room (Bunker)

Umbilical connection to the spacecraft (described in Section 5) is provided via the cable mast connected to the Zenit second stage and is disconnected at lift-off of the ILV. RF connection to the spacecraft is made through RF windows in the fairing, and also described in Section 5. The customer EGSE for connecting to these umbilical and RF links is positioned in room 114, an underground equipment room located near the launch pad (Figure 6-14). Room 114 is 10.5 meters by 5.6 meters in size. Though it is in the “unmanned area” during launch final countdown, it is well protected from the environment generated by the launch.

Room 114

Launch Pad

Figure 6-14. Location of Room 114 (Customer EGSE Room)


-


Command Center

During pre-launch and launch, SC personnel are located in the Command Center (CC), in rooms 131, 132 and/or 137 as shown in Figure 6-15. Each of these rooms is more than 60 square meters in size. Customer areas in the CC are equipped with:

• fire and environmental control systems • CCTV monitoring of the launch pad and the ILV • connections to spacecraft EGSE in the bunker, for monitoring of

SC parameters during the countdown • connections to the voice net for customer polling during

countdown The CC is two levels down inside a reinforced concrete underground building that provides protection for personnel during launch.

Entrance to CC

Room for location of Customer personnel and equipment

Room 131

Building 4

Room 132

Room 137

Building 4

Room for location of Customer personnel and equipment

Room 131

Building 4

Room 132

Room 137

Building 4

Room 131 Room 132

Room 137

Figure 6-15. Customer Location Options in the Launch Command Center



6.6 Cosmodrome Amenities Visa and Access Authorization

Land Launch supports customers in obtaining entry visas to Russia by providing the written invitations. To travel to Baikonur cosmodrome it is necessary to obtain a double or multiple visa. Land Launch also provides customer representatives with access to the Cosmodrome as well as badges to the required facilities.

Customs Clearance

Land Launch supports the customer in obtaining customs clearances at all ports of entry and exit as required for the transport of spacecraft and associated GSE. According to the existing customs regulations, the SC and associated GSE will be brought into Kazakhstan as temporary imports (for re-export) and therefore exempt from duties. Nominal administrative fees may be associated with customs clearance in Russia. If so, such fees are the responsibility of the customer. Any customs or export/licensing processes (export license authorization) in the customer’s country of origin for equipment and propellants are the responsibility of the customer. The customer is also responsible for providing all associated packing lists and invoices.

Transportation All work-related transportation of customer personnel and equipment is provided, starting from arrival at the local airport until departure from the airport. All vehicles for personnel are equipped with air-conditioning and heating systems. If necessary, additional vehicles (e.g., VIP transportation) may be rented in Baikonur. Upon request, and preferably one day in advance, Land Launch can provide transportation to meet atypical customer needs, including night shifts.

Consumables The customer will be provided in the PPF and/or HPF with de-ionized water, ethyl alcohol, compressed air for tool operation, pressurized nitrogen and helium, breathing air system for SCAPE and clean room garments. The customer should provide his own safety-critical equipment such as SCAPE.

Security Around-the-clock security is ensured to preclude access of unauthorized personnel to the SC. This coverage commences with SC arrival to the Cosmodrome Baikonur airport through launch.

Schedules Customers are provided with daily and workweek schedules. The typical workweek is six days, Monday through Saturday. Additional working time or other daily/weekly schedules can be arranged on a case-by-case basis.


-


External Communications

The customer is provided with local and international telephone/facsimile communication, e-mail and internet access, and access to allotted commutated ground and satellite international channels to transmit data between Baikonur cosmodrome and the SC customer control center. Usage fees will be coordinated in advance.

Medical Care During the launch campaign, Land Launch provides continuous access to a medical staff that can provide treatment to sick or injured personnel. Land Launch has the capability for an emergency medical evacuation to the United States or Europe if required. The medical center for providing the first treatment is located at Site 254 and at a clinic at Site 2Zh located two kilometers from Site 254.

Accommodations and Dining

Hotel accommodations are available at the Sputnik hotel (http://www.sputnikhotel.com/), located in the city of Baikonur, and on base in hotels at Site 2Zh near site 254. The Sputnik Hotel offers 120 comfortable rooms and five suites, one restaurant, a bar, a fitness center, a conference hall, offices, a swimming pool, a sauna, a gymnastic hall, a hairdresser, mountain bikes and a variety of other amenities. The hotels at 2Zh (Figure 6-16) accommodate up to 350 people in comfortable single and double rooms. Site 2Zh also features a café-canteen, a medical clinic, the Baikonur museum and the original buildings used by Yuri Gagarin and Academician Korolev – which upon special arrangement can be toured and photographed.

Figure 6-16. Hotel 1 at Site 2Zh Near the Site 254 PPF



7. SPACECRAFT DESIGN & VERIFICATION REQUIREMENTS Overview Most of the design considerations for spacecraft interested in flying on

Land Launch are defined in other sections of this User’s Guide, including:

• performance capabilities (Section 3) • ground and flight environments (Section 4) • spacecraft mechanical and electrical interfaces (Section 5) • facility interfaces (Section 6)

The first part of this section provides remaining design considerations that do not fall into the above categories, including:

• constraints on spacecraft RF transmitting and receiving • horizontal handling • safety requirements • ground support equipment considerations

The second part of this section outlines key methods and criteria for verifying that the spacecraft meets major design considerations.

7.1 Additional Spacecraft Design Considerations 7.1.1 Constraints on Spacecraft Transmitting and Receiving From launch until at least 20 seconds after spacecraft separation, the

spacecraft transmitters normally will not be used nor will commands be uplinked to the spacecraft. During integrated ground operations (after the spacecraft has been attached to launch vehicle) spacecraft transmitters should only be used at times and at frequencies that have been coordinated in advance. This normally consists of RF tests in the PPF and on the launch pad, during which the spacecraft transmitters should avoid intentional or unintentional radiation levels above 30 dB µV/m between 1570 MHz and 1630 MHz (the frequency range for the Glonass receivers on the launch vehicle) as measured one meter below the spacecraft separation plane. At all other frequencies between 10 KHz and 40GHz, spacecraft:

• unintentional emissions should not exceed 70 dB µV/m • intentional emissions should not exceed 140 dB µV/m

The maximum spacecraft intentional RF impingement on the launch vehicle is shown in Figure 7-1.



0

20

40

60

80

100

120

140

160

0 5000 10000 15000 20000 25000 30000 35000

Glonass Receivers 1570-1630 MHz

Elec

tric

field

, dB

µV

/m

Frequency, MHz

Figure 7-1. Maximum Intentional Spacecraft Electric Field Impingement on the Launch Vehicle (one meter below the separation plane)

7.1.2 Horizontal Handling Spacecraft systems and procedures must be compatible with the

spacecraft being placed in a horizontal attitude for several days (approximately seven), between encapsulation in the fairing inside the PPF and erection of the fully assembled launch vehicle at the launch pad, and again after T-0 in the unlikely event of a launch abort. While it is horizontal the spacecraft will be subjected to random vibration and accelerations during hoisting, transportation and launch vehicle assembly operations. These environments are defined in Section 4.



7.1.3 Safety Design Considerations Pressurized systems

Design and verification of spacecraft flight hardware pressurized vessels, structures, and components of the SC pressurized systems must be in accordance with recognized aerospace industry design guidelines. These pressurized systems should also be consistent with the Land Launch system payload environment. The design of the pressurized systems 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 both the horizontal and vertical orientations should be considered. Design details may be required as a portion of documentation to support regulatory agency requirements for the mission.

Ordnance systems Ordnance systems aboard spacecraft for operation of propulsion, separation, and mechanical systems must be designed in accordance with recognized standards and regulations. These systems must preclude inadvertent firing when subjected to Land Launch specified shock, vibration, thermal, or electromagnetic environments. Ordnance devices must be classified in accordance with applicable government codes and meet applicable 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 may be required to support regulatory agency requirements for the mission.

7.1.4 Ground Support Equipment (GSE) Considerations Customer GSE and checkout equipment which will be used at the

Baikonur launch base should be in accordance with the recognized safety requirements, and be capable of functioning with the facility interfaces and under the conditions (temperature/ humidity mode, cleanliness class, power supply parameters, gas supply, etc.), that are defined elsewhere in this User’s Guide. Design details may be required as a portion of the documentation used to satisfy regulatory agency requirements for the mission.



7.2 Spacecraft Verification Requirements Flexibility to meet customer needs

The design verification processes and criteria defined below are guidelines. They can be individually tailored to reflect the specific requirements of each Land Launch customers.

7.2.1 Spacecraft Structural Capability Factors of safety

The minimum factors of safety and test levels for several test options are shown in Table 7-1.

Table 7-1. Factors of Safety and Test Options

Factors of safety

Test option Yield Ultimate Test level Test success criteria 1.0 - No detrimental

deformation at 1.0 Qualification test (test of dedicated article to ultimate loads)

1.0 1. 3

1.25 - No failure at 1.25 Proto-qualification test (test of article used subsequently for flight or system test)

1.25 1.4 1.25 No detrimental deformation or misalignment

Qualification by analysis (test of article not required)

1.6 2.0 N/A N/A

Acceptance test (performed on each flight article)

1.1 1.3 1.1 No detrimental deformation

Test-verified model required for final CLA

Preliminary loads for quasi-static events are calculated using the Land Launch quasi-static load factors for ground handling, transportation, and flight that are defined in Section 4. Preliminary operational loads for transient flight events are calculated by Land Launch in a coupled loads analysis (CLA) using a preliminary spacecraft model provided by the customer. The final CLA operational loads used for verification must be generated using a test-verified spacecraft model. Verification of the spacecraft model can be performed either by modal survey or sine test.



Test requirements

Land Launch requires verification of a spacecraft’s structure load-carrying capability. The Land Launch qualification requirements on the spacecraft reflect standard practice, with appropriate tailoring to accommodate specific spacecraft and mission-specific characteristics. Structural testing on the spacecraft generally depends on the design heritage. Unique qualification tests can be developed by the spacecraft customer to account for design heritage. Land Launch will work with the spacecraft customer by evaluating customer-proposed testing to support the spacecraft integration process. The following tests are accepted by Land Launch for demonstrating structural compliance:

- modal survey test; - static loads test; - sine vibration testing; - acoustic testing; - shock qualification.

If a candidate qualification approach is not addressed here, Land Launch is open to proposed alternatives for ensuring spacecraft compatibility.

Modal survey test

The objective of a modal survey test is to determine the dynamic characteristics of the spacecraft structure. Following the test, the spacecraft mathematical model is adjusted. The adjusted model is categorized as “test verified.”

Static loads test Spacecraft static load testing is one option for validating the spacecraft structural strength. The extent of testing depends on the heritage of the spacecraft structure, but compliance with Table 7-1 is the objective.



Sine vibration testing

The sine vibration test levels for qualification, proto-qualification and acceptance are shown in Table 7-2, and are derived from the operational levels defined in Table 4-2. Qualification testing is for a dedicated test article. For proto qualification testing, the test article will be the first flight unit of a spacecraft series. Acceptance testing is generally performed to demonstrate workmanship and is an option available to the spacecraft customer for minor spacecraft design changes on proven spacecraft designs that have not previously flown on Land Launch. The extent of the testing depends on the heritage of the spacecraft structure. Spacecraft compliance must be demonstrated as described above. Test envelope “notching” (decrease of sine environment amplitudes) may be employed to prevent excessive loading of the spacecraft structure. However, the resulting sine vibration environment with notching should not be less than a test amplification factor level times the equivalent sine vibration level determined by CLA and provided by Land Launch. The test amplification factor levels depend on the test options chosen and are described above. Table 7-2. Sine Vibration Amplitudes and Sweep Rates

Test Frequencies Vehicle Qualification Proto-qualification AcceptanceZenit-2SLB 0.75 g 0.75 g 0.6 g 5 – 100

(Longitudinal and Lateral) Zenit-3SLB 0.88 g 0.88 g 0.7 g Test frequency sweep rate (octaves/min) 2 4 4



Acoustic testing Land Launch requires that each new spacecraft design undergo acoustic

qualification or proto-qualification testing. The maximum expected flight acoustic levels are provided in Section 4. Acoustic levels and durations for qualification, proto-qualification and acceptance testing are shown in Table 7-3.

Table 7-3. Spacecraft Acoustic Margins and Test Durations

Test Levels Duration [sec] Qualification + 3 dB over levels of

maximum expected acoustic environment

120

Proto-qualification + 3 dB over levels of maximum expected acoustic environment

60

Acceptance Maximum expected acoustic environment

60

Shock qualification

The spacecraft should be compatible with the shock environment in Section 4 with a 3-dB margin. The shock environment in Section 4 represents the maximum expected environment (operational levels) with no added margin. Analysis, similarity, and/or test can demonstrate qualification. A shock test using the spacecraft flight article, spacecraft adapter and flight equivalent separation system can be performed to qualify the spacecraft.



7.2.2 Matchmate Test Land Launch generally uses off-the-shelf spacecraft adapters with a flight

history on other launchers. For spacecraft types lacking prior flight experience with the adapter/separation system to be used, Land Launch requires a matchmate test between the spacecraft and the adapter flight hardware (see Figure 7-2). This test is usually performed at the spacecraft manufacturer’s production facility. It includes mechanical and electrical mate and checkout. It can also include the firing of the spacecraft separation ordnance to define the shock environment at the SC interface. Repeat missions of a spacecraft type typically do not require a matchmate. Instead, an adapter fit-check will typically be performed in the PPF at Baikonur at the start of the processing flow.

265807J3-033

Figure 7-2. Electrical and Mechanical Matchmate Test



8. MISSION INTEGRATION AND OPERATIONS Overview This section describes mission management, mission integration and

launch operations for a typical Land Launch mission.

The detailed topics are:

• Mission management.

• Mission documentation and schedule.

• Mission analysis.

• Operations planning.

• Launch Campaign.

• Post flight operations.

• Safety.

• Quality Assurance.

The Launch Services Agreement (LSA) for a Land Launch mission con-tains a Statement of Work (SOW) that details the services shown in this section and identifies the tasks and deliverables. The mission integration master schedule will contain these tasks and deliverables.

8.1 Mission Management Mission Manager To provide the customer with efficient mission integration and launch

operations processes, Land Launch assigns a Mission Manager to be the direct point of contact between the customer and the Land Launch or-ganization throughout the program. The Mission Manager will be re-sponsible for ensuring the timely and satisfactory completion of all as-pects of the mission including technical analyses, documents, launch campaign and schedule.

The Mission Manager organizes and chairs all meetings and reviews ac-cording to the Mission Integration Schedule, of which an example is pro-vided in Section 8.2. Land Launch and the customer will meet as often as necessary to support the mission.



Mission Team Roles and Responsibilities

The mission team is responsible for all mission analyses, operations planning, and launch campaign activities. On Land Launch the mission team is comprised of the Mission Manager, SIS (see Section 1 for a de-scription of SIS) and Boeing’s technical staff. SIS will perform the mis-sion analyses in Section 8.3 as well as the launch campaign. Boeing’s technical staff will be used in an oversight role.

8.2 Mission Documentation and Schedule Overview Land Launch will prepare all the documentation required for the launch

service. Land Launch will maintain configuration control of all signed mission documentation. Changes to signed documentation will be made via formal change proposals.

Integration Docu-mentation

At the beginning of the integration process, documentation needed for spacecraft integration will be coordinated with the customer and will be based on the Statement of Work (SOW) contained in the Launch Ser-vices Agreement (LSA).

Spacecraft/Land Launch System Interface Control Document

ICD Verification Matrix

Land Launch will create the spacecraft/Land Launch system Interface Control Document (ICD) that defines the technical interface require-ments for the launch service. It will be based on the customer Interface Requirements Document (IRD), any other spacecraft requirements, and the launch vehicle and launch site characteristics. Once signed by Land Launch and by the customer, this document will be under configuration control.

The ICD verification matrix defines the process by which the ICD re-quirements and functions are to be verified. The matrix is included in the ICD. All requirements must be met and verified prior to launch.



Mission Integra-tion Schedule

Land Launch will develop the mission specific integration schedule based on the launch date contained in the LSA. The schedule will contain all mile-stones necessary for a successful completion of the contract. A typical sched-ule for a non-recurring mission is provided in Table 8-1.

Table 8-1. Typical Mission Integration Schedule Event/Milestone Date Program Management Program Kickoff Meeting L - 18 Months Program Management Plan L - 17 Months Quarterly Management Report L - 15, 12, 9, 6, 3 Months Interface Activities Spacecraft Interface Requirements Document (IRD) L - 18 Months Spacecraft Environmental Test Plan L - 15 Months Preliminary Interface Control Document (ICD) L - 15 Months Preliminary Design Review (PDR) L - 11 Months Final ICD L - 10 Months Matchmate/Shock Test Plan L - 10 Months Critical Design Review (CDR) L - 6 Months Matchmate/Shock Test (first of a type) L - 5 Months Safety Spacecraft Safety Data Package L - 12 & L- 6 Months Launch Campaign Activities Ground Operations Working Group (GOWG # 1) L - 12 Months Launch Operations Plan L - 8 Months Spacecraft Preshipment Review (GOWG #2) L - 2 Months Spacecraft arrival at Baikonur L - 1 Month Combined Operations Readiness Review L – 1 Month Transfer Readiness Review (rollout to pad) L – 3 days Launch Readiness Review (LRR) L - 7 hours Launch L Post Launch Activities Spacecraft Orbit Data L + 45 Minutes Spacecraft GSE Shipment from Baikonur L + 3 days Flight Data Report L + 2 Months



8.3 Mission Analyses Mission Analyses Mission analyses are performed to verify that the spacecraft requirements are

satisfied and to demonstrate the compatibility of the spacecraft with the launch vehicle. Typically all mission analyses will be performed to support the Pre-liminary Design Review (PDR), and then repeated with updated spacecraft in-formation for the Critical Design Review (CDR). The typical analyses include:

• Mission design

• Spacecraft separation analysis

• Thermal analysis

• Coupled Loads Analysis (CLA)

• Clearance analysis

• Venting analysis

• RF link and EMC analyses

• Contamination analysis

The mission design will include the trajectory, orbit and dispersions, flight se-quences through spacecraft separation, verification of attitude requirements through separation and collision avoidance.

The spacecraft separation analysis determines the relative velocity between the spacecraft and the launch vehicle upon separation, the angular velocity of the spacecraft, and the clearances.

Thermal analyses are performed to demonstrate the spacecraft compatibility with thermal environments to which it will be exposed, during ground opera-tion and during launch through spacecraft separation. A spacecraft thermal model is required from the customer.

The CLA is performed to determine spacecraft loads, accelerations and dis-placements for the critical flight events. The CLA will allow the customer to determine structural margins. The CLA will determine the permissible notch-ing during spacecraft sine vibration testing. A spacecraft dynamic model is required from the customer.

The clearance analysis is performed to determine the clearances between the spacecraft and fairing during encapsulation and during flight. Manufacturing tolerances are included in this analysis.

The venting analysis determines the fairing depressurization rate during flight.



The RF link analysis is performed to demonstrate that spacecraft RF require-ments have been satisfied. The EMC analysis demonstrates RF compatibility between the spacecraft and launch vehicle at the launch base and during flight.

The contamination analysis is performed to verify that accumulated contami-nation during operations and flight will not exceed requirements.

8.4 Operations Planning Launch Campaign Planning

Launch site operations planning will be conducted as necessary during the mission integration phase and will be concluded before the beginning of the launch campaign. During the operations planning phase, an Opera-tions Plan and overall schedule will be produced.

The Operations Plan will be based on the spacecraft autonomous opera-tions requirements, the launch vehicle autonomous requirements, the combined operations requirements and the ICD. This plan will include a description of all operations from spacecraft arrival through launch in-cluding SC GSE departure after launch. The plan will include a list of reviews and major milestones.

The operations overall schedule will be based on the Operations Plan and serve as the baseline schedule for the launch campaign. The schedule will include all operations and reviews identified in the plan and will be the basis for the daily schedules produced at the launch site. When nec-essary, the schedule will be modified to reflect the current status of ac-tivities. A typical launch campaign schedule is shown in Table 8-2.

Table 8-2. Typical Launch Campaign Schedule Event/Milestone Completion Launch +/- time Spacecraft (SC) arrival at Baikonur, delivery to PPF L - 30 days SC to fueling area L - 20 days SC fueling complete L - 15 days SC mate to adapter L - 14 days Mate of SC/adapter stack to Block DM (Zenit-3SLB) or Intersection Bay (Zenit-2SLB) L - 12 days

Encapsulation L - 9 days Mate of Ascent Unit (Zenit-3SLB) or PLU (Zenit-2SLB) to Zenit stages in Area 42 L - 6 days Transfer Readiness Review L - 3 days Roll-out and erection on pad L – 2 days Launch Readiness Review (LRR) L – 7 hours Launch L

SC GSE leaves Baikonur L + 3 days



8.5 Launch Campaign Overview Land Launch makes use of existing facilities and processes for spacecraft

autonomous, launch vehicle autonomous and combined operations. Land Launch will support the customer during spacecraft autonomous opera-tions. Land Launch is responsible for all launch vehicle and combined operations.

The launch campaign begins with delivery of the spacecraft and GSE to the Baikonur Cosmodrome in Kazakhstan and is concluded with the shipment of the spacecraft GSE from Baikonur. The Spacecraft Preship-ment Review verifies the readiness of the SC and GSE for shipment to Baikonur.

Spacecraft Arrival and Transport

The SC and GSE arrive at the Yubileiny airfield within the confines of Baikonur Cosmodrome. The SC and GSE are offloaded from the aircraft and loaded onto rail cars on an adjacent railhead. The aircraft offloading area is located approximately 50 meters from the railway loading area. Land Launch will support the offload and provide equipment to load the SC and GSE onto rail cars, and to supply conditioning air to the SC con-tainer during rail transport. Land Launch will transfer the SC and GSE from the airfield to the Payload Processing Facility (PPF), either area 31 or 254 as previously agreed with the customer, where the containers are offloaded in an airlock.

Spacecraft Processing

The customer will perform the necessary standalone spacecraft opera-tions and tests in the PPF to prepare the spacecraft for fueling. These op-erations may include:

• Unpacking and visual inspection.

• Assembly and functional tests.

• Electrical checkout.

Land Launch provides the customer with the use of the PPF during standalone spacecraft operations, and will transfer the SC and GSE from the PPF to the Hazardous Processing Facility (HPF). At Area 31 where the PPF and HPF are in separate buildings, the SC will be enclosed in a container and supplied with conditioned air during transfer.

Spacecraft Fueling Spacecraft fueling is performed in the HPF. Land Launch and the cus-tomer will perform the following operations that include:

• Setup of SC.

• Setup of equipment.



• Setup of propellant tanks.

• Fill of the SC propellant and pressure tanks.

Hazardous operations are performed from a control room that is either safely remote from the fueling cell (Area 254 PPF) or blast-hardened (Area 31 PPF). Land Launch oversees these hazardous operations and ensures their safety.

Land Launch will transfer the fueled SC from the HPF to the PPF inte-gration bay.

Spacecraft Mating with Launch Vehicle Elements in the PPF Integration Bay

After SC fueling and return to the integration bay, the following opera-tions are performed:

• Final preparations of the SC and spacecraft adapter.

• SC mate on the adapter and attachment of the separation system.

• Mate of the spacecraft umbilical connectors.

• Electrical check.

• Mate of the SC/adapter stack with the Block DM-SLB (Zenit-3SLB) or Intersection Bay (Zenit-2SLB).

• Electrical check.

• Rotation of the stack to the horizontal position.

• Enclosure of the SC with the fairing.

• Integrated testing and checkout.

• Preparation for transfer to Area 42, the LV integration building. The Ascent Unit (Zenit-3SLB) or PLU (Zenit-2SLB) is installed on a rail-road car and a thermal conditioning unit is connected to the Ascent Unit/PLU. The environmental characteristics provided by condition-ing unit are shown in Section 4.

The Ascent Unit (or PLU) is then transferred to Area 42 via rail, with an air supply system providing the fairing volume with clean, conditioned air whose characteristics are described in Section 4.

Launch Vehicle Autonomous Processing

Autonomous processing of the Zenit stages and the Block DM-SLB (Ze-nit-3SLB missions) is completed before the start of combined operations.



Mating with Zenit Stages

Integration of the Ascent Unit (AU) or PLU with the Zenit stages takes place inside a clean room within Area 42. The AU/PLU is delivered by rail, then transferred to a rail-mounted integration trolley for horizontal mate to the Zenit second stage, much like assembly of the Sea Launch Zenit-3SL on the center rail of the Assembly and Command Ship. Zenit electrical connections are fully verified end-to-end beforehand.

Integrated Testing Integrated testing is performed following PLU/AU mating with the Zenit. The customer may perform SC battery charging prior to the ILV transfer to the pad. The clean room in Area 42 that is used for mating the AU/PLU to the Zenit may also be used to enclose the access door areas on the fairing if the customer needs to physically access the SC. The complete Integrated Launch Vehicle (ILV) is fully checked out prior to transfer to the launch pad (Area 45).

Transfer Readi-ness Review (TRR) and Transfer to the Launch Pad

A TRR is held prior to the transfer to the pad. This review verifies the readiness of the ILV, ground systems and launch base. The TRR author-izes the ILV transfer to the pad and subsequent final launch preparations.

In Area 42, the ILV is loaded onto the transporter-erector railcar (famil-iar to users of Sea Launch). This hoisting operation (no more than 1 hour) is the only time the payload enclosure is not either in the clean room or purged with clean, conditioned air. The ILV is transferred to the pad via rail, a distance of about four kilometers, in about 15 minutes.

Launch Pad Operations

Typical launch pad operations include:

• Erection of the ILV.

• Automated mating of all connectors (electrical, pneumatic, propel-lants, etc.).

• LV and SC checkouts and rehearsals.

• Integrated launch countdown rehearsals.

• Launch day countdown and launch.

Because of the highly automated nature of the Zenit, pad operations are typically very brief and launch can occur as little as 26-hours after arrival of the ILV at the launch pad unless the customer needs additional time for SC tests. The ILV can remain erect on the pad for up to four days, if necessary. Clean air to the fairing is maintained throughout.



Launch Readiness Review (LRR)

The LRR is held during launch day and prior to LV fueling, at approxi-mately L-7 hours. The purpose of the review is to status the readiness of all systems including the LV, SC and the range. This review authorizes the fueling of the LV and continuation of the launch countdown.

Propellant Loading

Loading the ILV with compressed gasses begins 25 hours before launch. Propellant loading begins at L-3 hr after a poll of launch management and the customer gives the “go” for launch. From this point forward until T – 50 seconds, launch holds of up to 20 minutes (or the length of the launch window, if less than 20 minutes) are possible. If a launch abort is called, a recycle to a second launch attempt can normally be accom-plished in two days.

Second “Go” Poll and Launch

The second “go” poll of launch management and the customer is con-ducted before disconnecting the transporter/erector from the launch vehi-cle at approximately L-12 minutes.

Launch Control Center

During all phases of launch processing and flight, the launch control cen-ter has ultimate responsibility and authority for all decisions and com-mands affecting the ILV. The Master Countdown Procedure as de-scribed in the SOW defines the communication architecture and protocol to be used by all parties during launch operations.



8.6 Post-Flight Activities Land Launch will provide the customer with orbital data at spacecraft

separation within 50 minutes after separation. Land Launch will support the customer in transferring GSE and containers from the work sites to Yubileiny airfield for shipment to the customer facility. Normally this ac-tivity is completed within three days after launch. Land Launch will transport the customer’s propellant containers to the port of entry into Russia for shipment from there to the customer facility.

8.7 Safety A safety evaluation of the spacecraft and launch processes is necessary to

demonstrate that equipment and operations are safe for a Baikonur launch campaign. The customer will need to provide necessary data to assess all potential hazards introduced by the spacecraft processing during the launch campaign.

Land Launch will define the data required and will conduct reviews for new launch customers or first of a kind spacecraft. Follow-on missions with similar spacecraft and operations use an abbreviated version of this approach in which changes to the previously approved mission are identi-fied. A safety demonstration will be accomplished through the submission of documents describing and defining all hazardous items and their proc-essing. Submission documents are prepared by the customer and are iden-tified in the customer SOW. The details of the necessary documents will be a topic at the program kick-off meeting.

A safety compliance certificate will be issued after Land Launch has veri-fied that the customer has demonstrated that the spacecraft and procedures are in compliance for a launch from Baikonur.

8.8 Quality Assurance

General Quality assurance, comprising qualification, acceptance, configuration management, quality, and reliability, falls under the purview of the Sea Launch Mission Assurance process. Sea Launch Mission Assurance is performed by the Mission Assurance organization and comprises the integrated set of systems and processes that ensures mission success and ever-increasing product reliability. Quality assurance is a continuous cycle that starts with the review of qualified flight-proven hardware and continues through the mission postflight data review and implementation of corrective actions and lessons learned for the next mission. Between



these points, the Sea Launch quality control and change management functions ensure that the vehicle reliability and system performance characteristics are maintained through a contolled set of launch vehicle processes and procedures

Hardware Review

A distinguishing feature of Sea Launch as an international launch services partnership is the hardware quality review that is performed at the factory level by Boeing engineers. This adds an extra layer of quality control and oversight to the rigorous systems already in place at each sub-contractor, and contributes to the outstanding launch success record at Sea Launch. This proven and existing hardware review system will be used to provide the same quality control and oversight for hardware that is destined for Land Launch missions.


Rev. Initial Release HPD-19000 A-1

APPENDIX A

User Questionnaire

When completed, this questionnaire provides the Sea Launch Company, LLC with the preliminary data necessary to begin evaluation of the compatibility of the Land Launch system with a new spacecraft type and to start the integration process. Companies considering the Land Launch system should complete this questionnaire and return it to:


One World Trade Center

Suite 950

Long Beach, CA 90831

USA (Attention: Ms. Paula Korn)

Sea Launch Company, LLC will treat this data as customer proprietary information and will not disclose any part of the information contained herein to any entity outside the Sea Launch Limited Partnership without your expressed written permission.

For further information contact:

Ms. Paula Korn

Phone: 562-499-4700

Fax: 562-499-4755

Email: [email protected]


A-2 HPD-19000 Rev. Initial Release

APPENDIX A

User Questionnaire Spacecraft Physical Characteristics SI Units English Units

Maximum height above I/F plane mm in Spacecraft protrusions below I/F plane (drawing) Maximum cross-sectional area mm in Static envelope drawing Adapter interface drawing Spacecraft volumetric displacement m3 ft3 Spacecraft free air volume m3 ft3 Spacecraft coordinate system (drawing) Mass properties* at launch (and separation, if different) Dry/wet Dry/wet Mass kg lb Center of gravity (origin on centerline, at I/F plane) X axis (X assumed to be longitudinal axis) mm in Y axis mm in Z axis mm in Moments and products of inertia about the CG Ixx kg-m2 slg-ft2 Iyy kg-m2 slg-ft2 Izz kg-m2 slg-ft2 Ixy kg-m2 slg-ft2 Ixz kg-m2 slg-ft2 Iyz kg-m2 slg-ft2 PLF access hatches (drawing)

Size Location Purpose When are they used?

* Identify values as maximum, minimum, or nominal and list tolerances as appropriate.



Spacecraft Orbit Parameters SI Units English Units

Final orbit apogee km mi Final orbit perigee km mi Final orbit inclination deg RAAN deg Argument of perigee deg Launch date Launch window min

Guidance Parameters

Maximum angular rate at separation Roll/spin deg/sec

Tolerance ± deg/sec

Pitch and yaw deg/sec

Tolerance ± deg/sec

Separation attitude

Separation velocity m/sec ft/sec

Maximum pointing error (cone angle) deg

Maximum tip-off rate deg/sec

Minimum tip-off rate (if applicable) deg/sec

Maximum angular acceleration deg/sec2



Electrical Interface

In-flight separation connectors MS part number (flight) MS part number (EGSE) Electrical power requirements External power for spacecraft bus Voltage V Current A Battery charging Battery voltage V Maximum current A Hard-line link Link required Y / N Remote bus voltage sense Y / N Baseband command rate bps Command baseband modulation Baseband telemetry PCM code Baseband telemetry rate bps RF link RF telemetry link required? Y / N If Yes, complete the following table.



Effective isotropic radiated power (EIRP) (dBW) Maximum - Nominal Minimum Antenna gain (referenced to boresight) (dB)

0

±10°

±20°

±30°

±40°

±60°

±80°

Antenna location (mm) Xsc Ysc Zsc


Antenna boresight

Polarization

Bandwidth Transmit frequency Subcarrier frequency Data rate Subcarrier modulation Carrier modulation Carrier modulation index Required link error rate (BER)



RF command link required? Y / N If Yes, complete the following table

Required Flux Density at SC Antenna (dBW/m2)

Maximum Nominal Minimum Antenna gain (refers to boresight)

0°

±10°

±20°

±30°

±40°

±60°

±80°


Antenna boresight Polarization

Frequency (only one at a time) Carrier modulation Data rate Required carrier-to-noise density at spacecraft antenna

External communications Ethernet RS 422 Brewster link In-flight interfaces Serial PCM? Y / N Number of channels PCM code Data rate Analog telemetry? Y / N Number of channels Samples per second Digital telemetry (discrete) Number of channels



In-flight discrete commands Number of commands Contact closure or voltage pulse Times of occurrence Duration Status requirements Separation break wires Numbers required Electromagnetic radiation curve

Spacecraft missions Spacecraft requirements

Does the spacecraft have any special EMC concerns (e.g., lightning or RF protection)?

What frequencies of LV intentional emitters are of special interest to the spacecraft?

What are the spacecraft sensitivities to magnetic fields?

When in launch configuration, which antennas will be radiating?



Thermal Environment SI Units English Units

Spacecraft allowable air temperature range Ground processing (pre-PLF encapsulation) to °C to °F After encapsulation to °C to °F

Prelaunch to °C to °F

EGSE allowable air temperature range Ground processing at the PPF to °C to °F LV integration building, area 42 to °C to °F

Launch complex, area 45 to °C to °F

Allowable humidity range Spacecraft processing % to % RH After encapsulation in PLF % to % RH Prelaunch % to % RH EGSE % to % RH Maximum ascent heat flux (pre-PLF jettison) Maximum free molecular heat flux At PLF jettison W/m2 W/ft2 Following PLF jettison W/m2 W/ft2 Maximum ascent depressurization rate Pa/s psi/s Maximum ascent differential pressure Pa psi Maximum differential at PLF jettison Pa psi Thermal maneuver requirements in flight Orientation requirements in flight Heat dissipation Spacecraft processing W After encapsulation W Prelaunch W What thermal analyses are required of Sea Launch?



Dynamic Environment Allowable acceleration loads Vertical Cantilevered Lateral g g

Axial g g

Allowable vibration curve at interface Sinusoidal Random Allowable acoustics environment Allowable shock curve Fundamental natural frequencies Axial Hz Lateral Hz Maximum allowable air impingement meters/

sec



Ground Processing Requirements SI Units English Units

Match mate test? Y / N Test location Payload separation system (PSS) firing required? Y / N SCA separation distance mm in Transportation to Baikonur Spacecraft Transport method Point of entry (Moscow airport name) Spacecraft container dimensions Height m ft

Width m ft

Length m ft

Container weight (with spacecraft) kg lb

CG of shipping container Environmental control and monitoring equipment Special handling considerations GSE Number of GSE shipping containers

Maximum container dimensions m x m x m ft x ft x ft

Maximum container weight (loaded) kg lb

Delivered with spacecraft? Y / N Transport method

Location delivered to

Environmental control and monitoring equipment

Special handling considerations

Payload Processing Facility (PPF) Spacecraft dimensions with handling fixture m x m x m ft x ft x ft

Spacecraft weight with handling fixture kg lb

Time required in processing facility days



Ground Processing (continued) Standard facility services provided in Baikonur are outlined in section 6 of the user’s guide. Any special services required by the user should be described below. Spacecraft Handling

Electrical power

Temperature/humidity

Grounding, ESD control

RF

Cleanroom

Consumables or commodities

Environmental monitoring

Hazardous materials storage

Ordnance storage

Other

EGSE Handling

Electrical power


Other

Operations Communication requirements

Storage

Security

Office space

Access (disabilities, equipment)

Other



Ground Processing (continued) ILV integration, area 42

How long can the spacecraft remain encapsulated and in horizontal orientation?

Describe spacecraft access requirements in the LV integration facility

Standard facility services provided in area 42 are outlined in section 6 of the user’s guide. Any special services required by the user should be described below. Spacecraft

Electrical power

PLF temperature, humidity, and cleanliness

RF control

PLF access (frequent environmental controls)

Purges

Consumables


Other

EGSE (size and weight)

Electrical power


Other

Operations

Communications

Storage

Security

Personnel accommodations

Office space

Other



Ground Processing (continued) Launch Complex, area 45 Standard facility services provided at the launch complex are outlined in section 6 of the user’s guide. Any special services required by the user should be described below. Spacecraft Electrical power

PLF temperature, humidity, and cleanliness

RF control

Purges

Consumables


Access

Other

EGSE (size and weight)

Electrical power


Other

Operations Communications

Storage

Security

Personnel accommodations

Office space

Other



Contamination Control Requirements Facility environments (indicate any applicable environmental limits for your payload): Airborne particulates (per FED STD 209)

Airborne hydrocarbons ppm

Nonvolatile residue rate mg/m2 - month mg/ft2 - month

Particle fallout rate % obscuration/day

What are your prelaunch spacecraft cleanliness requirements and verification methods (e.g., visibly clean, % obscuration, NVR wipe, etc.)? Do you have launch phase allocations for contamination from the booster? If yes, indicate allocations: Molecular

(µg/cm2) Particulate (% obscuration)

Thermal control surfaces

Solar array

Optical surfaces

Star tracker or sensors

Other

Do you require a gaseous spacecraft purge? ______________ Type and quality of gas _______________________________________________________ Continuous or intermittent _________________________________________________________

What are your specific cleanroom garment requirements (e.g., smocks, coveralls, caps, hoods, gloves, or booties)?

Do you have any specific cleanroom process restrictions (e. g., use of air tools, air pallets, welding, or drilling)?

Other special contamination control requirements:


Rev. Initial Release HPD-19000 B-1

APPENDIX B. LAND LAUNCH STANDARD OFFERINGS AND OPTIONS

This appendix lists Land Launch standard offerings and options for

• Standard-offering hardware.

• Standard-offering launch vehicle performance.

• Standard-offering launch services.

• Standard-offering facilities and support services.

• Optional services.

STANDARD-OFFERING HARDWARE The following lists the basic hardware items and associated services that

will be provided for a Land Launch mission.

Launch Vehicle • Zenit-2SLB two-stage booster

or

• Zenit-3SLB three-stage vehicle that includes the Block DM-SLB upper stage.

Payload Fairing (PLF) and Spacecraft Adapter (SCA)

• PLF with up to two access doors, RF window and a customer logo.

• SCA with compatible mechanical interfaces and separation system. Current standard offering adapters are SCA2624 and Zenit-2 adapter for the Zenit-2SLB and SCA937, SCA1194, and SCA1666 for the Zenit-3SLB. Contact Boeing Launch Services for adapters not listed.


B-2 HPD-19000 Rev. Initial Release

Electrical Interfaces

RF interfaces RF (command and telemetry) links between spacecraft and spacecraft electrical ground support equipment (EGSE) during operations when the integrated launch vehicle is erect at the launch site.

Hard-line interfaces

• Two in-flight disconnect connectors for telemetry, command, and power interfaces to the spacecraft.

• Hard-line (command and telemetry) umbilical links.

• In-flight disconnect break wires as needed.

STANDARD-OFFERING LAUNCH VEHICLE PERFORMANCE The following lists the basic performance that will be provided for a two

stage or three stage Land Launch mission.

Orbit and Mass • 12,000 kg payload systems mass to a 51.60 inclined, 400 km low earth orbit for Zenit-2SLB two-stage booster.

• 3,600 kg payload systems mass to GTO requiring a net velocity change of 1500 meters per second to GSO for Zenit-3SLB three-stage vehicle that includes the Block DM-SLB upper stage.

Orbit Accuracy • See Section 3 for the injection accuracy for the Land Launch family of vehicles to representative orbits.

STANDARD-OFFERING LAUNCH SERVICES

The following sections list the basic launch services to be provided for a Land Launch mission.

Mission Management

Mission Management includes program-level services, such as contracts, scheduling, program management, mission integration and launch operations

• Program Management Plan



Meetings and Reviews

• Customer kickoff meeting.

• Preliminary and final ICD review.

• Preliminary and Critical Design Reviews (PDR & CDR).

• System readiness review.

• Ground operations working group and technical interchange meetings as required.

• Ground operations readiness review (ready to receive spacecraft).

• Combined operations readiness review (ready to mate spacecraft to SCA).

• Ascent unit (or payload unit for the Zenit-2SLB) readiness review (ready to transfer ascent unit to LV integration facility).

• Transfer readiness review (ready to transport the ILV to the launch pad).

• Launch readiness review (ready to launch).

• Post launch flight data review.

Documentation The following documents will be provided to each customer according to the mission integration schedule (developed by Sea Launch with customer concurrence after mission identification):

• Mission integration schedule.

• Program Management Plan.

• Spacecraft/Land Launch ICD.

• Launch Operations Plan.

• Master Countdown Procedure.

• Launch commit criteria.

• Matchmate plan (if matchmate is required)



Mission Integration

Analyses The following analyses will be conducted. Unless noted otherwise, each analysis will involve two cycles (preliminary and final) for first-time integration, and one cycle for repeat integration.

• Mission design.

• Spacecraft shock analysis (one cycle for both first-time and repeat integration).

• Spacecraft acoustic analysis (one cycle for both first-time and repeat integration).

• Spacecraft separation analysis.

• Critical clearance analysis (dynamic).

• Coupled loads analysis.

• Pre-launch environmental control system analysis (normal and abort scenarios).

• Flight thermal analysis.

• Venting analysis.

• Contamination assessment.

• EMC assessment.

• Spacecraft on-pad RF link analysis (one cycle for both first-time and repeat integration).

• Spacecraft/PLA clearance analysis (static).

• Post flight analysis (one cycle for both first-time and repeat integration).

Verification Land Launch will perform verification of spacecraft requirements as defined in the ICD.

Separation state vector A spacecraft separation state vector is normally provided to the customer within 50 min of separation.

Interface test For each first-of-type spacecraft, Land Launch will perform a matchmate test at the spacecraft factory that includes the SCA and a flight equivalent separation system. Matchmate tests consist of mechanical and electrical interface testing between spacecraft and SCA.



STANDARD-OFFERING FACILITIES AND SUPPORT SERVICES

The following sections list the facilities and support services available to each Land Launch customer for the processing, assembly, installation, transportation, fueling and launch of their spacecraft. See Section 6 for details.

Payload Processing Facilities (PPF)

Land Launch will assist the customer with unloading the spacecraft and GSE from the aircraft at the cosmodrome airport. Land Launch will provide for the transportation of this hardware and miscellaneous equipment from the airfield to the PPF. The PPF facilities are described in detail in the Land Launch Users Guide. Facilities at the PPF are available to the customer for approximately five weeks during the launch campaign. The activities and services at the PPF include:

• Fit check: a spacecraft-to-spacecraft adapter fit check

• Spacecraft checkout and fueling

• Encapsulation

• Ascent Unit (Payload Unit for the Zenit-2SLB) checkout

• Control rooms

• Remote, or blast hardened, control room

• Offices

• Conference room

• Environmentally controlled areas.

Customer operations may be performed and scheduled on a 24-hr basis with access to the customer office building and the PPF. All customer equipment must be removed from Land Launch facilities within three (3) workdays after the launch is completed. Following the launch, Land Launch will assist the customer and provide for the transportation of the GSE from the PPF to the local airfield.



PPF Communications

The following communication services will be available to customer personnel during PPF operations:

• Intercom

• TV monitors and cameras

• Telephones with international access

• Launch video

• Countdown net

• Hand-held radios

PPF Security • Twenty four hour perimeter security

• Security devices on all internal and external doors leading to the payload processing areas

• Twenty four hour electronic monitoring for ingress and egress, and smoke and fire detection

PPF Support Services

• Separate technical and facility electrical power

• Use of clean room compatible mechanical support equipment (e.g., man-lifts and forklifts)

• Photographic services

• Protective garments for clean rooms and hazardous operations

• Negotiated quantities of gaseous nitrogen, liquid nitrogen, gaseous helium, isopropyl alcohol, and other general-purpose cleaning agents and solvents

• Monitoring system for detection of hazardous vapors

• Receipt and staging of spacecraft propellants

• Receipt and storage of ordnance components

• Use of calibrated weights

• A photocopier and a facsimile machine



Launch Vehicle Integration Facility, Area 42

Land Launch will provide facilities for personnel and spacecraft GSE in the LV integration facility, Area 42, for spacecraft check out and battery charging if required. Land Launch will provide power and conditioned air to the spacecraft. Land Launch will assist the customer during movement and installation of spacecraft GSE from the PPF to Area 42. Typical spacecraft tasks are:

• Spacecraft battery charging

• Spacecraft/Launch Vehicle umbilical line validation

• Spacecraft check out

Launch Complex (LC) Facilities, Area 45

Land Launch will provide LC facilities for customer personnel and spacecraft GSE. Land Launch will provide the following services:

• Power

• Gas conditioning

• RF link from the spacecraft to spacecraft GSE

• Assistance in installing GSE in a customer GSE room

• Assistance in SC/LV umbilical line validation

• Launch complex safety training

Launch Complex Communications

The following communication links are available within the launch complex, area 45:

• Intercom (including countdown net)

• TV monitors and cameras (including launch video)

• Dedicated telephone lines to international circuits

Launch Complex, Area 45, Security

• Full time security for the spacecraft and GSE

• The facilities have 24-hr electronic monitoring for smoke and fire detection



Environmental Controls

Land Launch will maintain spacecraft environments within the limits specified in the Spacecraft/ Launch Vehicle Interface Control Document (ICD) while the spacecraft is in the Payload Processing Facility, encapsulated within the ascent unit or payload unit, during transit to and at the launch vehicle integration complex, transit on the ILV to the launch pad, erection to vertical, and final countdown and launch.

Range Services Land Launch will coordinate with the customer and submit all required range documentation for the launch of the customer’s spacecraft on the Zenit Launch Vehicle. Land Launch will provide all necessary telemetry assets to verify launch vehicle functions and conditions from lift-off to spacecraft separation.

Logistics Support • Accommodations and meal services: Land Launch will make hotel reservations, and make meal services available for the customer team during working hours on base

• Medical Services: clinic staffed by a medical doctor • Emergency evacuation to Western Europe • Travel reservations • Daily weather forecasts • Daily transport between hotel and Cosmodrome facilities • Bilingual Secretary • Translators

OPTIONAL SERVICES

The following services can be provided by Land Launch on customer request. These services are considered optional and are subject to negotiation. Requests for other services, not described in this document, will be considered by Land Launch on a case-by-case basis.

Mission Analysis Repeating any analysis listed or any additional analysis or design work due to changes made by customer special request.



Interface Tests • Matchmate testing at the spacecraft factory for repeat of spacecraft type.

• Separation shock test.

Support Services • Additional secretaries and/or translators.

• Charter airplane for personnel and/or spacecraft parts.

Facilities • Additional office space depending on availability.

• Use of any facilities beyond the time frame allocated to the customer.

Materials • Gaseous nitrogen.

• Oxygen.

• Helium.

• Liquid nitrogen.

• Isopropyl alcohol.

• De-ionized water.

• Solvents and gases in addition to negotiated quantities.

section 1 introduction - geo ring

Documents