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					                                 Land Launch
User’s Guide
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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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                             Abbreviations and Acronyms

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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                             Abbreviations and Acronyms

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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                              Abbreviations and Acronyms

W                   watt(s)

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




xxii                                                       Initial Release
                                                                        Land Launch User’s Guide Section 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.




Rev. Initial Release                         HPD-19000                                               1-1
                                                                   Land Launch User’s Guide Section 1




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




1-2                                       HPD-19000                          Rev. Initial Release
                                                                         Land Launch User’s Guide Section 1




What the sys-      The standard Land Launch system comprises:
tem includes
                   •   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      Land Launch provides:
the Customer
                   •   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




Rev. Initial Release                       HPD-19000                                                  1-3
                                                                                         Land Launch User’s Guide Section 1




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
                             18 mo                        12 mo                        6 mo                          0 mo


                                     Sign Contract

                                                 Mission Analysis

                                                                                  Operations Planning

                                                                               S/C Delivery and processing
                        New Mission
                                                                                                        Launch Ops




                              12 Month Integration Timeline
                                                          12 mo                        6 mo                          0 mo


                                                               Sign Contract

                                                                    Mission Analysis

                                                                                  Operations Planning

                                                                               S/C Delivery and processing
                        Repeat Mission
                                                                                                        Launch Ops




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




1-4                                             HPD-19000                                               Rev. Initial Release
                                                                        Land Launch User’s Guide Section 1




Baikonur               The Zenit Launch Complex is located at the Baikonur Cosmodrome in
Cosmodrome             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.




Rev. Initial Release                       HPD-19000                                                 1-5
                                                              Land Launch User’s Guide Section 1




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




1-6                                   HPD-19000                         Rev. Initial Release
                                                                               Land Launch User’s Guide Section 1




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.



                                            Sea Launch                Federal Space Agency
                                           Company, LLC                       (FSA)

                   - Program Management                                    - CIS Licensing
                   - Contracts and Legal Support                           - CIS Governmental Interfaces
                   - Insurance Interface
                   - US Licensing


                                                               Space International
                         Boeing                                  Services, Ltd.
                                                                     (SIS)
                   - Sales & Marketing
                   - Mission management                           Yuzhnoye SDO
                   - Quality and Technical Oversight              Yuzhmash PO
                   - Hardware Acceptance Review                   KBTM
                   - Customer Safety and Logistics Support        TsENKI
                                                                  RSC Energia
                                                                  NPO Lavochkin

                                                                   - Launch Services
                                                                   - Mission Integration
                                                                   - Baikonur Operations




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




Rev. Initial Release                               HPD-19000                                                1-7
                                                                 Land Launch User’s Guide Section 1




Sea Launch      Sea Launch Company, LLC, headquartered in Long Beach, California, is
Company, LLC    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      The Boeing Company is enlisted by Sea Launch to provide the following
Company         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           Space International Services, Ltd (SIS) is a company comprised of SDO
International   Yuzhnoye, PO Yuzhmash, RSC Energia, the Center for Ground Space
Services, Ltd   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




1-8                                HPD-19000                               Rev. Initial Release
                                                                    Land Launch User’s Guide Section 1




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




Rev. Initial Release                      HPD-19000                                              1-9
                                                                  Land Launch User’s Guide Section 1




     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




1-10                                   HPD-19000                            Rev. Initial Release
                                                                      Land Launch User’s Guide Section 1




Design Bureau of       The principal offices of KBTM are located in Moscow, Russia. In its
Transport              fifty year history KBTM has designed numerous launch complexes in-
Machinery              cluding the Zenit complex at Baikonur (Figure 1-6), and also performs
(KBTM)                 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




Rev. Initial Release                      HPD-19000                                              1-11
                                                                     Land Launch User’s Guide Section 1




Center for Ground   The principal offices of TsENKI are located in Moscow, Russia in close
Space Infrastruc-   association with the Russian Space Agency. TsENKI is responsible for
ture Operations     operation of the ground aerospace infrastructure facilities at Baikonur
(TsENKI)            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




1-12                                   HPD-19000                           Rev. Initial Release
                                                                         Land Launch User’s Guide Section 1




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




Rev. Initial Release                      HPD-19000                                                 1-13
                                                                 Land Launch User’s Guide Section 1




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   Land Launch enjoys the support of the Federal Space Agency which will
Agency          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.




1-14                               HPD-19000                           Rev. Initial Release
                                                                                                Land Launch User’s Guide Section 2




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

                                                   58.65 m (192.4 ft)

               Fairing             Block      Zenit Stage 2                           Zenit Stage 1
           10.4 m (34.1 ft)       DM-SLB     10.4 m (34.1 ft)                        32.9 m (107.9 ft)




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



                                             Figure 2-1. The Zenit-3SLB

                                                     Intersection Bay
                                                      0.35 m (1.1 ft)
                                                                            57.4 m (188.3 ft)

                      Fairing                 Zenit Stage 2                                 Zenit Stage 1
                  13.7 m (44.8 ft)           10.4 m (34.1 ft)                              32.9 m (107.9 ft)




                                                                        Ø3.9 m (12.8 ft)



                                             Figure 2-2. The Zenit-2SLB




Rev. Initial Release                                     HPD-19000                                                           2-1
                                                                 Land Launch User’s Guide Section 2




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            The original Zenit-2 was first launched in 1985 from Baikonur Cos-
Flight History   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         From its introduction in 1974 through March 2004 the Block DM has
Flight History   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).




2-2                                  HPD-19000                             Rev. Initial Release
                                                                       Land Launch User’s Guide Section 2




Flight Success         Tables 2-1 and 2-2 list flight records for each of the three Zenit-3SL
Ratios                 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
                               Year            Versions          Cumulative Flight
            Stage
                            Introduced          Flown               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
                                 Year                                       Reliability
            Stage                                  Flight Record
                              Introduced                                    Estimate
      Zenit Stage 1
      Zenit Stage 2*
                                 1999
                                 1999
                                                      12 of 12
                                                      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.




Rev. Initial Release                       HPD-19000                                                2-3
                                                                    Land Launch User’s Guide Section 2




2.1. The Land Launch Zenit

Design Heritage   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.




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




2-4                                     HPD-19000                             Rev. Initial Release
                                                                         Land Launch User’s Guide Section 2




Changes Made           Significant configuration differences between the heritage Zenit-2 and the
for Sea Launch         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                Zenit specifications and performance parameters are shown in Table 2-3.
Specifications         Stage 1 and Stage 2 configurations are pictured in Figure 2-4. With pro-
and Configura-         pellant mass fractions exceeding 90%, the designs of both stages rank
tions                  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.




Rev. Initial Release                         HPD-19000                                                2-5
                                                                    Land Launch User’s Guide Section 2




                               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)
                            One RD-171                   One RD-120 Main Engine
 Engines
                       (four thrust chambers)   One RD-8 Vernier Engine (four thrust chambers)
                            740,000 kgf
 Thrust (sea level)                                             Not applicable
                         (1.63 million lbf)
                            806,400 kgf             Main Engine: 93,000 kgf (205,028 lbf)
 Thrust (vacuum)
                         (1.78 million lbf)         Vernier Engine: 8,100 kgf (17,857 lbf)
 Specific impulse
                              309.5 s                           Not applicable
 (sea level)
 Specific impulse                                            Main Engine 350 s
                              337.2 s
 (vacuum)                                                   Vernier Engine 342.8 s
 Attitude Control     Nozzle gimbal + 6.3 deg     Vernier engine nozzle gimbal + 33 degrees




2-6                                         HPD-19000                         Rev. Initial Release
                                                                                             Land Launch User’s Guide Section 2




                                                                        Avionics bay

                                                                        Liquid oxygen tank
                                              Stage 2
                                           10.4 m (34 ft)
                                                                         Kerosene tank
                                                                         Main engine
                                                                          SPST (4)
                        Separation plane                                 Steering engine

                                                                         Interstage frame



                                                                    ∅3.9 m
                                                                    (12.8 ft)




                                       Stage 1
                                   32.9 m (108 ft)                      Liquid oxygen tank




 Solid-propellant separation                                            Kerosene tank
 thrusters (SPST) (4)

 Main                                                                    Single turbopump
 engine
 nozzles (4)




                                                            А   А
                     View A-А


               Figure 2-4. Land Launch Uses the Same Zenit Stages that are Used
                                 on the Sea Launch Zenit-3SL




Rev. Initial Release                                        HPD-19000                                                     2-7
                                                                   Land Launch User’s Guide Section 2




2.1.1        Zenit Stage 1

Overall          The Land Launch Zenit Stage 1 (Figure 2-5) features an aluminum primary
Configuration    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




2-8                                     HPD-19000                            Rev. Initial Release
                                                                       Land Launch User’s Guide Section 2




RD-171M Engine         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.




                                                                           265807J3-038


                        Figure 2-6. The RD-171M is the Most Powerful Liquid Rocket Engine
                                             Presently in Operation




Rev. Initial Release                       HPD-19000                                                2-9
                                                                    Land Launch User’s Guide Section 2




2.1.2        Zenit Stage 2

Overall Configu-     Like the first stage, the Zenit second stage (Figure 2-7) features an inte-
ration               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



2-10                                     HPD-19000                            Rev. Initial Release
                                                                     Land Launch User’s Guide Section 2




RD-120                 The 2nd stage main engine is a single-chamber, fixed nozzle liquid-
Main Engine            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                   The RD-8 vernier engine mounted in the aft end of Stage 2 provides
Vernier Engine         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.




Rev. Initial Release                      HPD-19000                                              2-11
                                                                 Land Launch User’s Guide Section 2




2.2    The Block DM-SLB Upper Stage

Overall          The Block DM-SLB (Figure 2-8) used on the Zenit-3SLB is closely de-
Configuration    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           The Block DM-SLB upper stage is powered by the 11D58M engine,
Main 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         Three–axis stabilization and attitude control during coast periods, includ-
Control/Ullage   ing continuous rolls, are provided by two attitude control/ullage engines
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     Significant configuration differences between the heritage Block DM and
for Sea Launch   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



2-12                                  HPD-19000                            Rev. Initial Release
                                                                           Land Launch User’s Guide Section 2




 Block DM-SLB          The main differences between the Block DM-SLB and the Block DM-SL
 Versus the            are:
 Block DM-SL              • 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
                                 1
                       Length                                          5.93 m (19.4 ft)
                       Diameter (primary)                              3.7 m (12.1 ft)
                                                  2, 3
                       Maximum Launch Mass               (fueled)      17,800 kg (39,240 lb)
                                     3
                       Maximum 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




Rev. Initial Release                            HPD-19000                                              2-13
                                                     Land Launch User’s Guide Section 2




                                                              Avionics Container




                                                              Avionics/Payload Truss




                                                              Coolant Piping




                                                               Upper Adapter




                                                                LOX Tank


                                                                Middle Adapter



                                                                Kerosene Tank




                                                                 Attitude Control/Ullage Engine


                                                                 Lower Adapter



                                                                 Main Engine




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




2-14                        HPD-19000                           Rev. Initial Release
                                                                                        Land Launch User’s Guide Section 2




2.3      Zenit-3SLB Ascent Unit

Components and Inte-            The Zenit-3SLB Ascent Unit (figure 2-9) consists of the spacecraft,
gration                         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.


                                                       PSS/BDM
                                  Spacecraft           Interface PLF/BDM                              Stage 2/Stage 3
                                  Interface Plane                Interface Plane   Block DM-SLB (BDM) Interface Plane
   Payload         Spacecraft
   Fairing (PLF)                  Payload Support
                                  Structure (PSS)




                                  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         Once inside the PLF, physical access to the spacecraft is gained through
Characteristics        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.


Rev. Initial Release                                HPD-19000                                                           2-15
                                                                  Land Launch User’s Guide Section 2




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

Fairing Thermal   The internal and external thermal insulation of the PLF nose cone protects
Protection        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   The payload support structure for the Zenit-3SLB is provided by RSC
Structure         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.




2-16                                   HPD-19000                            Rev. Initial Release
                                                                                 Land Launch User’s Guide Section 2




2.4    The Zenit-2SLB Payload Unit

Components and         The Zenit-2SLB payload unit (PLU), shown in Figure 2-10, consists of
Integration            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.

                                                   Spacecraft Adapter      Interface Truss
           Fairing
                        Ø3.9 m (12.8 ft)




                                                                                               Intersection
                                                                                               Bay


                                             13.6 m ( 44.8 ft)

                                              14.0 m ( 45.9 ft)




                                           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         Access doors up to 500 mm x 500 mm (19.7 inches x 19.7 inches) can be
Characteristics        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.




Rev. Initial Release                                       HPD-19000                                         2-17
                                                                   Land Launch User’s Guide Section 2




Conditioned Air     Clean, conditioned air is provided to the payload fairing volume from en-
Supply to the       capsulation until launch including during transport between facilities.
Fairing             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     External thermal insulation protects the payload structure from overheat-
Protection          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          For dedicated launches of a single spacecraft on the Zenit-2SLB, Land
Adapters            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   For spacecraft that have unique interface and separation requirements,
and Multi-          Land Launch can examine other heritage spacecraft adapter designs, in-
Spacecraft          cluding those that incorporate bolt-type attachment and separation mecha-
Launches            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.




2-18                                    HPD-19000                            Rev. Initial Release
                                                                     Land Launch User’s Guide Section 2




                                                    КА
                                                    SC




                                                    2-я ступень РН
                                                    LV Stage 2




                                       O2062




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




Rev. Initial Release                  HPD-19000                                                  2-19
                                             Land Launch User’s Guide Section 2




       This page intentionally left blank.




2-20                  HPD-19000                        Rev. Initial Release
                                                                   Land Launch User’s Guide Section 3



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




Rev. Initial Release                      HPD-19000                                                3-1
                                                                Land Launch User’s Guide Section 3



Performance    Performance data in this section is based on the following set of ground
Ground Rules   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.




3-2                                  HPD-19000                            Rev. Initial Release
                                                                      Land Launch User’s Guide Section 3



Launch Window          The launch vehicle and associated ground systems can support a launch
Availability           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                 Zenit-2SLB: 20 m/s (45 miles/hour)
                       (at 10m above ground surface)        Zenit-3SLB: 18 m/s (40 miles/hour)
                       Pad Turn-around Time Between
                                                            10 Days
                       Launches
                       Nominal Turn-around Time After       1 Day (if scrub precedes LV fueling)
                       Launch Scrub                         < 3 Days (scrub after LV is fueled)
                       Maximum Annual Launch Rate           Twelve (of which no more than seven
                       (Factory Limited)                    Zenit-3SLB)
                                                            Zenit-2SLB: 98%
                       Launch-on-Time Probability
                                                            Zenit-3SLB: 97%




Rev. Initial Release                        HPD-19000                                                 3-3
                                                                       Land Launch User’s Guide Section 3



3.1 Launch Site and Accessible Orbits

Site Location        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
                              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.




                Approved Azimuth
                Potential Azimuth
                                                                A=35.00, i=63.00


                                                                       A=64.20, i=51.40


                                                                         A=82.10, i=46.20
                                        Baikonur




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




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




3-4                                      HPD-19000                               Rev. Initial Release
                                                                     Land Launch User’s Guide Section 3




Accessible Orbits      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
                                           Accessible                     Usual Plane Change
                       Orbit Type                           Vehicle
                                          Inclinations                          Method
                                         46.2 < i < 71 o
                       LEO                                 Zenit-2SLB Second Stage Yaw
                                        84.0 < i < 105 o
                       MEO, HEO,                                         Third Stage Perigee,
                       GTO, Elliptical,                                  Apogee or Post-
                                         0.0 < i < 110 o Zenit-3SLB
                       escape                                            Perigee Burn
                       trajectories                                      (mission-specific)

                       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.




Rev. Initial Release                       HPD-19000                                                 3-5
                                                                       Land Launch User’s Guide Section 3



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 burn
                                                                       Injection into transfer orbit

                                  Block DM-SLB first burn
                                  Injection into parking orbit


               Transfer orbit
               HP = 200 km                   Zenit Stage I/II
               HA = 35,950 km
               Inc = 48.6 deg
                                       Parking orbit
                                       HP = 180 km
                                       HA = 417 km
                                       Inc = 51.4 deg

      Block DM-SLB third burn                                                Target transfer orbit
      Injection into target transfer orbit                                   HP = 4,100 km
                                                                             HA = 35,786 km
                                                                             Inc = 23.2 deg




      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)




3-6                                                  HPD-19000                   Rev. Initial Release
                                                                          Land Launch User’s Guide Section 3




          Launch point




                                                  Fairing
                                                  L=1924 km



                         First Stage
                         L=∼884 km




                                                                                                   Second Stage
                                                                                                   L=6850 km




     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.




Rev. Initial Release                            HPD-19000                                                 3-7
                                                                   Land Launch User’s Guide Section 3




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.




3-8                                     HPD-19000                            Rev. Initial Release
                                                                        Land Launch User’s Guide Section 3



Block DM-SLB           At approximately 65 to 90 seconds after second stage main engine
Powered Flight         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.




Rev. Initial Release                          HPD-19000                                                 3-9
                                                                 Land Launch User’s Guide Section 3




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




3-10                                    HPD-19000                          Rev. Initial Release
                                                                        Land Launch User’s Guide Section 3




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 of
                                                                                          DM-SLB ME




                                    2-nd ignition of         LV operation phase
                                    DM-SLB ME



                                                                  3-rd ignition of
                                                                  DM-SLB ME




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




Rev. Initial Release                          HPD-19000                                                      3-11
                                                               Land Launch User’s Guide Section 3



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.




3-12                                 HPD-19000                           Rev. Initial Release
                                                                   Land Launch User’s Guide Section 3




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




Rev. Initial Release                      HPD-19000                                              3-13
                                                                                Land Launch User’s Guide Section 3



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



                                                             Stage 2 MECO          SC Separation
                                                             Time=397 s            Time=894 s
                                                             Altitude=400 km       Altitude=400 km
                              Fairing Jettison
                              Time=295s
                              Altitude=173 km
                              FMH≈30 W/m2


             Stage 1 Separation
             Time=149 s
             Altitude=74 km

        Max Acceleration
        Time=113 s
        Accel=4.06 g


       Maximum Q
       Time=60 s
       Q=5370 kgf/m2

                                            Stage 1 Impact     Fairing Impact
                                            Range=884 km       Range=1924 km




         Figure 3-5. Typical Ascent Profile to the International Space Station Orbit at 51.6o
                                   with Payload Mass 12000 kg




3-14                                              HPD-19000                               Rev. Initial Release
                                                                                   Land Launch User’s Guide Section 3




Ground Track                       Figure 3-6 presents the predicted ground track for a Zenit-2SLB mission
                                   to the International Space Station.



  90
  80
                                                                 Stages 1 and 2
  70
                                                                 Operation
  60
  50
  40
  30
  20
  10
   0
 -1 0
 -2 0
 -3 0
 -4 0
 -5 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      30        60      90       120      150      180


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




Rev. Initial Release                                    HPD-19000                                                3-15
                                                                                           Land Launch User’s Guide Section 3




3.4 Payload Capability – Three Stage Zenit-3SLB

Geosynchronous                          The Land Launch Zenit-3SLB is a medium-lift vehicle to GTO.
Transfer Orbit                          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          Inclination           Perigee Altitude        Payload Systems
[meters/second]          [degrees]              [kilometers]          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
                       4500

                       4000

                       3500

                       3000
    Payload Mass, kg




                       2500

                       2000

                       1500

                       1000

                       500

                         0
                              0   200       400     600     800       1000       1200     1400    1600      1800     2000
                                                           Delta V to target orbit, m/s

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



3-16                                                       HPD-19000                                 Rev. Initial Release
                                                                                    Land Launch User’s Guide Section 3



MEO, HEO,                            The Land Launch Zenit-3SLB is a heavy lift vehicle to Middle Earth and
Circular and                         High Earth (MEO and HEO, respectively) circular and elliptical orbits that
Elliptical Orbits                    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
                                       Inclination          Inclination          Inclination
                           [km]                 o                   o
                                          51.4                 63.9                 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

                          5500
                                                                                               i=51,4
                          5000
                                                                                               i=63,9
                          4500
                                                                                               i=98,8
      Payload Mass, kg




                          4000

                          3500

                          3000

                          2500

                          2000
                                 0         5000    10000     15000    20000      25000    30000      35000
                                                            Orbital Height, km


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




Rev. Initial Release                                       HPD-19000                                              3-17
                                                                          Land Launch User’s Guide Section 3



               Table 3-8. Zenit-3SLB Performance to Elliptical Orbits
        Apogee                      Payload Capability [kg]
        Height        Inclination         Inclination          Inclination
                              o                   o
         [km]            51.4                63.9                 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

                          5100
                          5000                                                         i=51,4
                          4900                                                         i=63,9
                          4800                                                         i=98,8
                          4700
       Payload Mass, kg




                          4600
                          4500
                          4400
                          4300
                          4200
                          4100
                          4000
                          3900
                          3800
                                 10000 20000 30000 40000 50000 60000 70000
                                           Height of target orbit apogee, km



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




3-18                                          HPD-19000                             Rev. Initial Release
                                                                                          Land Launch User’s Guide Section 3




High-Energy and                              Table 3-9 and Figure 3-10 show the Zenit-3SLB payload capability to
Earth-Escape                                 high-energy orbits and Earth escape. These are presented as a function of
Trajectories                                 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

                                     6000
          Payload Systems Mass, kg




                                     5000

                                     4000

                                     3000

                                     2000

                                     1000

                                        0
                                            -30     -20      -10         0        10    20        30         40
                                                                   C3,   km2/s2
          Figure 3-10. Zenit-3SLB High-Energy and Earth Escape Payload Capability




Rev. Initial Release                                               HPD-19000                                            3-19
                                                                                                 Land Launch User’s Guide Section 3



3.5 Payload Capability - Two Stage Zenit-2SLB

Circular LEO                                Table 3-10 and Figure 3-11 present Zenit-2SLB payload performance as
Orbits                                      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                  Inclination               Inclination
                                                      51.4º                        63.9º                     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
                              15000



                              13000



                              11000                                                                    51.4º
                                                                                                       63.9º
       Paylo ad Mas s (kg )




                                                                                                       98.8º
                               9000



                               7000



                               5000



                               3000



                               1000
                                      200    400          600          800            1000   1200         1400         1600

                                                            Circula r Orbit He ig ht (km)


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

3-20                                                                HPD-19000                              Rev. Initial Release
                                                                                     Land Launch User’s Guide Section 3



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          Inclination              Inclination
                                                   51.4º               63.9º                    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


                         14000
                         12000                                                                     51.4º
  Paylo ad mas s (kg )




                         10000                                                                     63.9º
                                                                                                   98.8º
                         8000

                         6000
                         4000
                         2000
                            0
                                 0   1000    2000    3000    4000    5000    6000    7000      8000      9000      10000
                                                             Apogee altitude (km)

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




Rev. Initial Release                                         HPD-19000                                             3-21
                                                             Land Launch User’s Guide Section 3



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.




3-22                              HPD-19000                            Rev. Initial Release
                                                                        Land Launch User’s Guide Section 3



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




Rev. Initial Release                        HPD-19000                                                  3-23
                                                                 Land Launch User’s Guide Section 3



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         The separation system provides a relative velocity between the Block
Capabilities       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       The state vector at time of spacecraft separation may be delivered to the
Delivery           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.




3-24                                   HPD-19000                           Rev. Initial Release
                                                                     Land Launch User’s Guide Section 3



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             The Zenit-2SLB employs a typical three-axis stabilized method for
Capabilities           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




Rev. Initial Release                       HPD-19000                                               3-25
                                                                 Land Launch User’s Guide Section 3




CCAM for second   Collision avoidance is achieved by firing four solid-propellant
stage             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      The timing of state vector delivery depends on the mission profile as well
Delivery          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.




3-26                                  HPD-19000                            Rev. Initial Release
                                                                  Land Launch User’s Guide Section 4




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      Those levels associated with “ground handling and transportation” ad-
Environments           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         The coordinate system used in this section is shown in Figure 4-1. Dur-
Coordinate System 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.




Rev. Initial Release                       HPD-19000                                                   4-1
                                                                 Land Launch User’s Guide Section 4




                                                Y
                                             Vertical




                                                             X
                                                        Longitudinal
                                    Z
                                  Lateral


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




Environmental   Land Launch monitors and records spacecraft environments as specified
Monitoring      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.




4-2                                  HPD-19000                                 Rev. Initial Release
                                                                     Land Launch User’s Guide Section 4




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      Design reference maximum acceleration levels during ground transporta-
Factors, Ground        tion, handling and processing are defined in Table 4-1. The quasi-static
Handling and           accelerations levels are shown for the spacecraft center of gravity while
Transportation         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]                   Safety
      Spacecraft Processing Operation
                                                      X          Y          Z                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      Zenit-3SLB       ± 0.5      -1 ± 0.5       ± 0.4           1.5
in the PPF through completion of Zenit-2SLB          ± 0.35     -1 ± 0.2       ± 0.2           2.0
launcher assembly in Area 42)
Launcher on-loading and off-loading (crane lifts
                                                     ± 0.2      -1 ± 0.2       ± 0.2           1.5
in Area 42)
Roll out and erection on the launch pad              ± 0.35     -1 ± 0.2       ± 0.2           1.5




Rev. Initial Release                       HPD-19000                                                      4-3
                                                                                              Land Launch User’s Guide Section 4




Quasi-Static Load        From liftoff through spacecraft separation, the spacecraft is subjected to
Factors, Flight          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.

                                                                          5
                                      (-0.7, +4.5)                                 (+0.7, +4.5)

                                                                          4
                                                      Longitudinal* (g)
                                                                          3

                        (-2, +2)                                                                         (+2, +2)
                                                                          2


                                                                          1

                                    Lateral (g)
                                                                          0
                   -3          -2               -1                             0      1              2              3
                                                                          -1


                                                                          -2
                                     (-1, -2)                                             (+1, -2)

                                                                          -3


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

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




4-4                                                  HPD-19000                                                 Rev. Initial Release
                                                                  Land Launch User’s Guide Section 4




Sinusoidal             The longitudinal and lateral low-frequency sinusoidal vibration environ-
Equivalent             ments generated at the spacecraft separation plane during liftoff and
Vibration During       flight phases are within the limits defined in Table 4-2. The sinusoidal
Flight                 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]
                                5 - 100                 Zenit-2SLB                     0.6
                       (Longitudinal and Lateral)       Zenit-3SLB                     0.7




Rev. Initial Release                       HPD-19000                                                   4-5
                                                                Land Launch User’s Guide Section 4




4.2 Random Vibration

Ground Random     The spacecraft is subjected to low frequency random vibrations during
Vibration for     transportation by rail at the cosmodrome. Table 4-3 envelopes this ran-
Components Near   dom vibration environment. The maximum duration of any rail transfer is
the Spacecraft    six hours.
Interface
                     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             Y-Y              Z-Z
                                            Longitudinal      Vertical         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     The random vibration environment during flight at the spacecraft inter-
Vibration         face is enveloped in Table 4-4 and Figure 4-3.
Environment
                  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


4-6                                  HPD-19000                                Rev. Initial Release
                                                                                           Land Launch User’s Guide Section 4




                                     1




                                    0,1
   Spectral Power Density [g2/Hz]




                                    0,01




      0,001

                                          10                   100                      1000                     10000

                                                                      Frequency [H z]


                                               Figure 4-3. Random Vibration Environment During Flight




Rev. Initial Release                                                 HPD-19000                                                  4-7
                                                                  Land Launch User’s Guide Section 4




4.3 Acoustics

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

      Table 4-5. Maximum Expected Acoustic Pressure Envelope Inside Land Launch Fairings
                1/3 Octave Band                Acoustic Pressure Level [dB]
              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
                                             -5            -9
           Reference: dB in respect to 2 x 10 Pa (2.9 x 10 psi)




4-8                                       HPD-19000                             Rev. Initial Release
                                                                                              Land Launch User’s Guide Section 4




                                       140
      Sound pressure levels, dB


                                       130



                                       120



                                       110



                                       100
                                             10   100                                      1000                           10000

                                                    1/3 octave band center frequency, Hz
  Note:
  Overall acoustic pressure level = 140 dB


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

                                       140
        Acoustic pressure levels, dB




                                       130




                                       120




                                       110




                                       100
                                             10   100                                 1000                             10000
                                                    1/3 octave band center frequency, Hz

  Note:
  Overall acoustic pressure level = 142 dB




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




Rev. Initial Release                                      HPD-19000                                                                4-9
                                                                      Land Launch User’s Guide Section 4




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 Interface
Frequency
                 Band Tension (kN)                    Band Tension (kN)                 Band Tension
   (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




4-10                                         HPD-19000                              Rev. Initial Release
                                                              Land Launch User’s Guide Section 4




           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

Rev. Initial Release                    HPD-19000                                             4-11
                                                                                        Land Launch User’s Guide Section 4




                                                     Shock levels at separation plane

       Response [g] Q=10   10000




                           1000




                            100




                              10
                                100                                  1000                                    10000
                                                                Frequency [Hz]


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




4-12                                                          HPD-19000                               Rev. Initial Release
                                                                                          Land Launch User’s Guide Section 4




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

                           10000
       Response [g] Q=10




                            1000




                             100




                              10
                                100                                   1000                                      10000
                                                                 Frequency [Hz]


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



Rev. Initial Release                                           HPD-19000                                                  4-13
                                                              Land Launch User’s Guide Section 4




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           The ambient cosmodrome electromagnetic environment varies by loca-
Cosmodrome        tion, and changes over time as new equipment is introduced and older
Electromagnetic   equipment is retired. Figures 4-7, 4-8 and 4-9 therefore provide prelimi-
Environment       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.




4-14                                  HPD-19000                             Rev. Initial Release
                                                                                     Land Launch User’s Guide Section 4




                                 140


                                 130
    Field Intensity, dB µV/m


                                 120


                                 110


                                 100


                                     90


                                     80


                                     70


                                     60


                                     50


                                     40
                                       0,01     0,1       1         10         100              1000            10000

                                                              Frequency, MHz
   Figure 4-7a. Ambient Electromagnetic Environment within Payload Processing Facility
                                        Site 254
                               140


                               130


                               120
    Field Intensity, dB µV/m




                               110


                               100


                               90


                               80


                               70


                               60


                               50


                               40
                                 0,01         0,1     1            10          100              1000            10000

                                                              Frequency, MHz
   Figure 4-7b. Ambient Electromagnetic Environment within Payload Processing Facility
                                        Area 31



Rev. Initial Release                                          HPD-19000                                                 4-15
                                                                                                      Land Launch User’s Guide Section 4




                                        140


                                        130


                                        120
    Напряженно сть э лектр.по dB µV/m
    Field Intensity, ля, д БмкВ/м



                                        110


                                        100


                                         90


                                         80


                                         70


                                         60


                                         50


                                         40
                                           0,01           0,1          1             10         100              1000            10000

                                                                           Frequency, MHz
                                                                                Часто та, МГц



                                                  Figure 4-8. Ambient Electromagnetic Environment within the Launch
                                                                 Vehicle Assembly Building (Area 42)

                                        140


                                        130


                                        120
   Field Intensity, dBля, д БмкВ/м
      Напряженно сть э лектр.по µV/m




                                        110


                                        100


                                         90


                                         80


                                         70


                                         60


                                         50


                                         40
                                           0,01           0,1          1             10         100              1000            10000

                                                                           Frequency, MHz
                                                                                Часто та, МГц




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


4-16                                                                         HPD-19000                              Rev. Initial Release
                                                                    Land Launch User’s Guide Section 4




Launch Vehicle         The Zenit-2SLB has three sets of Sirius telemetry systems which are lo-
Radio Equipment        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
                                                   0/-7                                0/-5
Coefficient (dB)
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                    40                                   30
emissions (dB)


         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
                                                                      TBD
     Coefficient (dB)
     Output Power (W)                                                   17
     Reduced level relative to main signal of
                                                                        60
     spurious and harmonic emissions (dB)




Rev. Initial Release                         HPD-19000                                              4-17
                                                                    Land Launch User’s Guide Section 4




         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
                                                 -163                           -145
       Nominal Frequency (dBW)
       Max/Min Antenna Gain
                                                 7/-3                           TBD
       Coefficient (dB)




Radio Frequency         The maximum field intensity levels generated by launcher systems at the
Environment at          spacecraft interface plane are provided in Table 4-11, and Figures 4-10
the SC Separation       and 4-11. These account for such factors as the type and orientation of
Plane                   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




4-18                                        HPD-19000                             Rev. Initial Release
                                                                                                    Land Launch User’s Guide Section 4




                                  200

                                  180                                                                    Sirius TM System
                                                      Sirius TM system
                                                                                                         1,010.5 MHz
                                                      231.3 MHz
                                  160                                                                    1,018.0 MHz
                                                      239.3 MHz
                                                      247.3 MHz
                                  140
Electric field, dB µV/m




                                  120

                                  100

                                         80

                                         60

                                         40

                                         20

                                         0
                                              0                   100                        1,000                              100,000
                                                                         Frequency, MHz


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

                                         200
                                                                         Sirius TM system   Sirius TM System
                                         180                             231.3 MHz          1,010.5 MHz
                                                                         239.3 MHz          1,018.0 MHz
                                         160                             247.3 MHz
                                                                                                         BITC-B TM System
                                         140                                                             1,026.5 MHz
               Electric field, dB µV/m




                                                                                                         1,034.5 MHz
                                         120

                                         100

                                          80

                                          60

                                          40

                                          20

                                             0
                                                  0                100                      1,000                             100,000
                                                                         Frequency, 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)




Rev. Initial Release                                                     HPD-19000                                                  4-19
                                                                 Land Launch User’s Guide Section 4




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,    The spacecraft thermal and humidity environment is actively controlled
Ground Thermal       by facility, transportation and launch pad clean air systems from the time
and Humidity         the spacecraft container is offloaded at Baikonur airport through lift off.
Environments         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   The spacecraft is exposed to ambient facility air in the PPF from the time
Systems              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, conditioned air is supplied to the spacecraft container by a mobile
Clean Air Systems    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.


4-20                                     HPD-19000                             Rev. Initial Release
                                                                   Land Launch User’s Guide Section 4




Launch Pad Clean       A clean air system at the launch pad provides conditioned air to the fair-
Air Systems            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-        Airflow impingement upon the spacecraft surfaces is generally main-
locity of Airflow      tained at or below three meters per second.
Upon SC Surfaces




Rev. Initial Release                       HPD-19000                                               4-21
                                                                   Land Launch User’s Guide Section 4




               Table 4-12. Spacecraft Ground Thermal and Humidity Environment
                                              Acting        Temperature       Relative         Nominal
       Operations Phase/Location
                                              System            [0C]         Humidity         Flow Rate
SC Container, Airport to PPF Transfer        Mobile Unit      15 to 30*        ≤ 60%         3K-6K m3/h
                      Area 31 Area B                          15 to 28       35% - 60%
PPF Processing                               Facility Air
                      Area 31 Room 119                        17 to 23       40% - 60%            N/A
Areas                                        Conditioning
                      Site 254                                18 to 25       30% - 60%
Transfer from PPF to HPF (Area 31)           Mobile Unit      15 to 30*        ≤ 60%         3K-6K m3/h
                      Area 31                Facility Air     15 to 25       30% - 60%
PPF Fueling Cells                                                                                 N/A
                      Site 254               Conditioning       TBD            TBD
PPF Encapsulation Area 31                    Facility Air     15 to 28       35% - 60%
                                                                                                  N/A
Halls                 Site 254               Conditioning     18 to 25       30% - 60%
Transfer to ILV Integration Area (Site 42)   Mobile Unit      10 to 35*      30% - 60%       3K-6K m3/h
                                             Facility Air
ILV Integration Bay (Site 42)                                 18 to 25         ≤ 80%              N/A
                                             Conditioning
                                             Facility Air
Clean Room (Site 42)                                         21 to 26.7      30% - 60%            N/A
                                             Conditioning
ILV ready in Site 42 for roll-out            Mobile Unit      10 to 35*        < 60%          ≤ 3K m3/h
ILV transfer to launch complex (and from
                                           Mobile Unit        10 to 35*        < 60%         > 2250 kg/h
launch complex after launch abort)
ILV erection, and while
erect prior to LOX loading Zenit-2SLB                         10 to 35*     DP ≤ -100C        9500 m3/h
                                            Pad System
ILV de-erection if launch      Zenit-3SLB                      8 to 25*     DP ≤ -100C        5000 m3/h
aborted before T-12 min
ILV erect on launch pad,
                               Zenit-2SLB                     10 to 35*     DP ≤ -300C        9500 m3/h
from LOX loading until T-                   Pad System
                               Zenit-3SLB                      8 to 25*     DP ≤ -300C        5000 m3/h
12 min
ILV erect on launch pad, T-12 min to T-0
ILV erect on launch pad following launch
                                           High Pressure
aborted between T-12 and T-0, through                         10 to 32*     DP ≤ -550C       > 2250 kg/h
                                            Pad System
ILV de-tanking and de-erection (until mo-
bile unit is reconnected)
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




4-22                                         HPD-19000                           Rev. Initial Release
                                                              Land Launch User’s Guide Section 4




                             Second Stage Vents


     Fairing Vents                                           Fairing A/C Inlet
                                                             High Pressure Pad System
                                                             (from T-12 minutes)
                                 SC
                                                             Up to 2250 kg/hr


    Fairing A/C Inlets
    (to T-12 minutes)                                          Second Stage A/C Inlet
    Up to 9500 m3/hr          Separating                       (Up to 3600 m3/hr)
                              Screen



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




                             Third Stage A/C Inlet          Second Stage A/C Inlet
       Fairing A/C Inlet
                             (Up to 4500 m3/hr)             (Up to 3600 m3/hr)
       (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




Rev. Initial Release                      HPD-19000                                           4-23
                                                                   Land Launch User’s Guide Section 4




4.6.2 Flight Thermal Environments

General Overview,     After launch, the spacecraft will experience:
Flight Thermal        • Heat flux radiated from internal surfaces of the fairing, before fairing
Environment               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        500 W/m2 maximum                   400 W/m2 maximum
internal surfaces
Free molecular heating at      1135 W/m2 or less. Considerably less than 1135 W/m2 for most
fairing jettison               missions (typically around 50 W/m2) due to drop zone requirements
Free molecular heating         Typically not significant due Dependent on mission profile, but
after fairing jettison         to short mission duration.     may spike slightly at the time of the
                               Analyzed for each mission.     second Block DM-SLB burn as
                                                              shown in Figure 4-14
Heat radiated onto space-
craft surfaces by the sec-
                                 9.0 Kw-s/m2 maximum                 5.1 Kw-s/m2 maximum
ond stage solid propellant
separation thrusters
Solar heating, planet-       Typically not significant due    Analyzed for each mission. For
reflected solar heating (al- to short mission duration.       thermal management, the Block
bedo), Earth-radiated heat- Analyzed for each mission.        DM-SLB is designed to accommo-
ing, radiation to space                                       date preferred attitude pointing,
                                                              continuous rolls, maneuvers, and
                                                              orientations during coast and pre-
                                                              separation phases of flight




4-24                                        HPD-19000                            Rev. Initial Release
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                                                300
        Dis pers ed free molecular flux, W/m2

                                                250
                                                                                                                       MEO
                                                                                                                       GTO/GSO
                                                200




                                                150




                                                100




                                                 50




                                                  0

                                                      0        2000     4000     6000           8000           10000        12000       14000

                                                                                 Time from lift-off, seconds




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




Rev. Initial Release                                                              HPD-19000                                                            4-25
                                                               Land Launch User’s Guide Section 4




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     The depressurization rate, though varying somewhat by trajectory and
Rate               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           Due to the late timing of fairing jettison due to drop zone constraints, the
Differential at    maximum pressure differential between the pressure inside the fairing
Fairing Jettison   and the external pressure at fairing jettison does not exceed a very low
                   0.002 kgf/cm2 for both Zenit-2SLB and Zenit-3SLB.




4-26                                   HPD-19000                             Rev. Initial Release
                                                             Land Launch User’s Guide Section 4




                           Second Stage Vents




                                     Separating
                                     Screen


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




                                Third Stage Vents
                                          Second Stage Vents



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




Rev. Initial Release                    HPD-19000                                            4-27
                                                                   Land Launch User’s Guide Section 4




       Pressure, kgf/cm2




                                                  ------ Ambient Pressure
                                                             -- - - Ambient Pressure
                                                        Predicted Internal Pressure




                               Time From Launch, seconds



       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




4-28                                     HPD-19000                                 Rev. Initial Release
                                                                                                    Land Launch User’s Guide Section 4




     Pressure Change Rate, kgf/cm2 per second




                                                                        Fairing Cavity Pressure, kgf/cm2

                                                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




Rev. Initial Release                                                          HPD-19000                                             4-29
                                                                   Land Launch User’s Guide Section 4




4.8 Contamination


Contamination         The spacecraft is protected from contamination during ground processing
Control During        by:
Ground Process-
                      •   Supplying a continuous flow of clean, conditioned air (class 5,000 or
ing
                          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
Particle        Level 500            Maximum Fairing Surface Levels                  Level 750
  Size      per Mil-Std-1246C        Zenit-2SLB         Zenit-3SLB               per Mil-Std-1246C
>100 µm         11,900/m2             11,900/m 2
                                                         30,129/m2                   96,300/m2
>250 µm          281/m2                281/m 2
                                                          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




4-30                                       HPD-19000                             Rev. Initial Release
                                                                   Land Launch User’s Guide Section 4




Contamination          Potential sources of the SC contamination in flight are fairing materials,
Control During         contaminants migrating from the launch vehicle equipment bays while
Flight                 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         Materials exposed to the cavity shared by the spacecraft are selected to
Features               preclude crumbling, peeling, particle shedding, oxidation or corrosion,
to Minimize            and with low outgassing properties that should not exceed the following
Contamination          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                  The plume effect on the SC of the second stage solid propellant separa-
Impingement            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. Initial Release                       HPD-19000                                               4-31
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4-32              HPD-19000                                Rev. Initial Release
                                                                   Land Launch User’s Guide Section 5




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     Table 5-1 presents the allowable spacecraft mass and longitudinal center-
and Longitudinal    of- gravity (CG) limits for the Zenit-3SLB, based on the standard
Center-of           spacecraft adapters (SCA) currently offered by Land Launch. Customers
Gravity Location    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)
SC Mass             937 SCA                           1194 SCA                      1666 SCA
  (kg)         Band Tension (kN)                Band Tension (kN)                  Band Tension
             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




Rev. 01/2004                              HPD-19000                                                5-1
                                                                  Land Launch User’s Guide Section 5




                         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- The radial offset of the spacecraft CG, relative to the launch vehicle
of Gravity Radial  longitudinal centerline, should not exceed:
Offset
                      • 50-mm (Zenit-2SLB)
                         •   25-mm (Zenit-3SLB)
                      Exceptions should be brought to Land Launch for assessment.


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


                          Table 5-3. Recommended Spacecraft Fundamental Frequencies
                                               Spacecraft Fundamental Frequencies
                      Launch System
                                              Longitudinal              Lateral
                        Zenit-2SLB              > 15 Hz                  > 5 Hz
                        Zenit-3SLB              > 20 Hz                  > 8 Hz




5-2                                      HPD-19000                            Rev. Initial Release
                                                                 Land Launch User’s Guide Section 5




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)




Rev. 01/2004                            HPD-19000                                                5-3
                                                             Land Launch User’s Guide Section 5




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




5-4                                   HPD-19000                          Rev. Initial Release
                                                                 Land Launch User’s Guide Section 5




                                                      Longitudinal Separation
                                                      Joint




                                                         Fairing Field Joint

                                                          Spacecraft Adapter



                                                       Transfer Compartment




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




Rev. 01/2004                           HPD-19000                                                 5-5
                                                             Land Launch User’s Guide Section 5




                                                          Intersection Bay



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




5-6                                  HPD-19000                           Rev. Initial Release
                                                                          Land Launch User’s Guide Section 5




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.




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




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




Rev. 01/2004                                HPD-19000                                                     5-7
                                                                                Land Launch User’s Guide Section 5




                                                                                 SC envelope




                                               600
                                               504 (937 interface)
                                                     ø3645
                                               369 (1194 interface                    Transfer Compartment
 ø2072




                                               400 (1666 interface)
                      ø2103
                                                 All dimensions in millimeters
                                      øø3577




                                                                              ø3595
         ø3679




                                       3577




                              ø3723                                   ø3633                               ø3645



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




5-8                                                  HPD-19000                              Rev. Initial Release
                                                                 Land Launch User’s Guide Section 5




Useable Volume   There is room inside the payload support structure (spacecraft adapter and
Inside Payload   transfer compartment) for satellite motor nozzles. Dimensions will be
Structure        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         There is external space on either Land Launch fairing for a customer logo or
Insignia         insignia. Details will be coordinated between the customer and the Land
                 Launch team




Rev. 01/2004                           HPD-19000                                                 5-9
                                                        Land Launch User’s Guide Section 5




                                         Location of access
                                         doors to SC


                                                 Location of access
                                                 doors to SC




                                                                     Location of access
                                                                     doors to SC



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




5-10                            HPD-19000                           Rev. Initial Release
                                                                                  Land Launch User’s Guide Section 5




                            Р а зве ртка цилиндриче ской ча сти головного обте ка те ля
                                 Unrolled cylinder of the Zenit-2SLB fairing
                                З она возможного ра зме ще ния люков доступа
                               Crosshatch = Access door allowable locations
                                                  6°

                        I                    II            III               IV             I




                                                                                                560 1215
                 2495
                 2495
                 2495




                                                                                                  80
                                                                                  40
                 925




                        I               II                  III         40   IV             I
                                   6°                             55°
                                                                 74°


                  Плоскость стыка ГО с пе ре ходным отсе ком
                     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         A full range of spacecraft adapters (SCA) made by Saab Ericsson Space
Adapters                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


Rev. Initial Release                                   HPD-19000                                                5-11
                                                                Land Launch User’s Guide Section 5




   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.




5-12                                    HPD-19000                           Rev. Initial Release
                                                                    Land Launch User’s Guide Section 5




Zenit-2 Adapter        The Zenit-2 frame adapter has been successfully used on a number of
(Frame) for Use        missions. The spacecraft is attached to four pyro-locks on a 2062-mm
with Zenit-2SLB        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        Features of the Zenit-2SLB that make it inherently well suited for
Dispensers for Use     launching groups of satellites, as well as secondary payloads, include:
with Zenit-2SLB
                          •   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.




Rev. Initial Release                       HPD-19000                                              5-13
                                                                   Land Launch User’s Guide Section 5




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
                                            Up to 20 twisted shielded pairs,
                     Signal
                                            each at 250 mA at 50 V ac or 100 V dc
                     Power (External
                                           20 A at 110 V dc
                     Spacecraft Power)
                     Power (Battery        20 A at 110 V dc (from T-12 hours until T-0)
                     Charging)             70 A at 110 V dc (all other times)
Umbilical Use        The umbilical circuits defined above are available for customer use at
During Processing    each stop in the launch processing flow as shown in Table 5-6.
and Launch
                               Table 5-6. Umbilical Hook-Up Locations and Availability
                                   Location for Connecting to
                     Place                                          Notes
                                   Customer-Provided EGSE
                                   Processing Cell
                                                                    After attaching the SC to
                     PPF           Processing Cell Control Room
                                                                    the adapter
                                   Fuel Cell Control Room
                                                                    After attaching the SC and
                     PPF           Encapsulation Hall               adapter to the upper stage
                                                                    (Zenit-3SLB)
                     Site 42       Clean Room and EGSE Room
                                                                    While erect on the pad
                                   Bunker (power circuits)
                     Launch                                         through launch, and after
                                   Customer Control Room
                     Complex                                        T-0 until lowered in the
                                   (signals circuits)
                                                                    case of a launch abort


5-14                                      HPD-19000                            Rev. Initial Release
                                                                    Land Launch User’s Guide Section 5




Umbilical              Umbilical circuits are connected to the spacecraft by separation
Connectors             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             The launch vehicle will detect spacecraft separation via two diametrically
Verification           opposed separation switches at the adapter/spacecraft interface.
                       Indication of spacecraft separation will be telemetered via redundant
                       channels.




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                                                            Land Launch User’s Guide Section 5




Satellite      Lateral and longitudinal accelerations will be recorded near the interface
Environment    between the spacecraft and the adapter, telemetered and received. Also
Measurements   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.


Commands       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:


                   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
                                      + 35 ms
               Actuation Timing
               Signal Duration        64 + 8 ms
               Current                0.01 to 1.0 Amperes
               Voltage                Up to 34 V




5-16                               HPD-19000                            Rev. Initial Release
                                                                     Land Launch User’s Guide Section 5




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%         380 V+ 2%            120 V, 20 A,
                                             50 Hz + 2%           50 Hz + 1%           single phase
                                                                    3 Phase               60 Hz
                       PPF Processing
                                                 80 kW               60 kW                 TBD
                       Cell
                       PPF Fueling
                                                 80 kW               60 kW                 TBD
                       Area
                       Launcher
                                                 40 kW               40 kW                 TBD
                       Assembly Bldg
                       Launch Complex
                                                 20 kW               20 kW                 TBD
                       Bunker
                       Launch Complex
                                                 20 kW               20 kW                 TBD
                       Control Room

Uninterruptible        Customer EGSE power is backed-up with a dedicated Uninterruptible
Back-up Power          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
                                    220/380 VAC, three phase "Y" at 40 KVA with power
                         50 Hz
                                    factor from unity to 80% lagging
                                    120/208 VAC, three phase "Y" at 40 KVA with power
                         60 Hz
                                    factor from unity to 80% lagging




Rev. Initial Release                       HPD-19000                                               5-17
                                                             Land Launch User’s Guide Section 5




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.




5-18                                HPD-19000                            Rev. Initial Release
                                                                     Land Launch User’s Guide Section 6




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
         N                                         Yubileinyi airfield
                                                              Site 31 Payload Processing

                                                               Site 42 Zenit and ILV
                                                                    Processing




                                                                Site 45              75 km
                        Area 254                        Zenit launch complex
                  Payload and Block DM
                       Processing



                                                          Syrdar-ya River
                                                     City of Baikonur

                 Krainy Airport                             Legend
                                                               Railroad
                                                               Road


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




Rev. Initial Release                     HPD-19000                                              6-1
                                                                  Land Launch User’s Guide Section 6




6.1 Transportation of Personnel and Cargo to and from Baikonur

Krainy      Personnel fly between Moscow and Baikonur via Krainy Airport (Figure 6-2),
Airport     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   Yubileiny Airport is located 45 km north of Baikonur city within Baikonur
Airport     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.




6-2                                      HPD-19000                           Rev. Initial Release
                                                                       Land Launch User’s Guide Section 6




                       Figure 6-3. Spacecraft Off Load at Yubileiny Airport

Transportation at      Rail and road networks connect all Land Launch facilities at Baikonur.
the Cosmodrome         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.




Rev. Initial Release                       HPD-19000                                              6-3
                                                            Land Launch User’s Guide Section 6




                                                                          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.

6-4                                  HPD-19000                           Rev. Initial Release
                                                                                Land Launch User’s Guide Section 6




                                                       125
                                               122


                                                                          124            51
          46                       45
                                               120
                                                                                 40А
                       44
                                                                     40

                                                               40Е               40Д
                                                63       380                           48Б     48 48А     87

                                                        Storage

                                                                                 87А    43
                                                      Lavochkin NPO
                                                         Storages
                                                                                       105           57




 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




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                                                                Land Launch User’s Guide Section 6




Building        Buildings 40/40D at Area 31 are used for non-hazardous payload
40/40D, PPF     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    Air-conditioned office facilities are provided at Site 31. These facilities
Office areas,   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,    The SC is fueled in the Hazardous Processing Facility, Building 44, located
HPF             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




6-6                                     HPD-19000                            Rev. Initial Release
Land Launch User’s Guide Section 6


                                                                                      -




N
                                                                                          Building 40                    Sanitary inspaction room   Building 40А


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


                                                                       Clean Rooms




                   Sliding gate
                   8.4 х 10 (h)
                                                                                                           Area - A                                 Bridge Crane
                                                                                                                                                    5 tf Н=8.1 m




                                                       Area - С                                         Area - B



                                                                                                          Sliding gate
                                                                                                          5.5 х 7 (h)                                Building 40Д




                                                                Figure 6-6. Lay-out of Buildings 40 and 40D
                                                                                        Land Launch User’s Guide Section 6




Encapsulation Bay                                                                                          SC Processing Room
Area – 300 sq m                                                                                            Area – 240 sq m
Height – 18.35 m                                                                                           Height – 10.8 m
Class 100,000 clean                                                                                        Class 10,000 clean
Crane 1: 5 t, 16.5 m                                                                                       Crane: 5 t, 8.1 m
Crane 2: 10 t, 14.6 m
Crane 3: 50 t, 13.7 m




SC Processing Room
Area – 300 sq m
Height – 18.35 m
Class 100,000 clean
Crane 1: 10 t, 14.6 m
Crane 2: 50 t, 13.7 m




                        Figure 6-7. SC Processing and Joint Operations Area in Buildings 40 and 40D
Land Launch User’s Guide Section 6


                                                              -




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
                                                                     Land Launch User’s Guide Section 6


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.


Customer Offices
                                                                               Area 101




                                             Encapsulation Area

                                      Area 102


                               Block DM
                               Processing
                                            Loading/Unloading Area      Container Cleaning and
                                                                             Acceptance



                                                              SC Processing/Fueling Area
             Uninterruptible Power
             Supply

                                                                         Sumps for Spill/
                                                                         Run-Off Containment



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


Site 254             The main areas of the PPF for the SC are 101, 102 and the proposed
PPF layout           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

6-10                                         HPD-19000                          Rev. Initial Release
                                                                   Land Launch User’s Guide Section 6


                                                     -
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




Rev. Initial Release                    HPD-19000                                              6-11
                                                           Land Launch User’s Guide Section 6


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.




                                                                          3P11107

                                                                       SDO Yuzhnoye Photo


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




6-12                               HPD-19000                          Rev. Initial Release
                                                                      Land Launch User’s Guide Section 6


                                                       -
Integration Area       Building 41 is 120 meters long and 60 meters wide, with three parallel
Layout/Features        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             An equipment room is available for customer use in Building 41,
Equipment Room         equipped with the power supplies and the umbilical connections to the
                       spacecraft that are defined in Section 5.




Rev. Initial Release                     HPD-19000                                                6-13
                                                               Land Launch User’s Guide Section 6



Customer Office   Air-conditioned customer office facilities are provided at in Building 41,
Areas             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    Launch operations are highly automated on Land Launch just as on Sea
Automated         Launch. This has many advantages including:
Systems
                     •   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.




6-14                                  HPD-19000                           Rev. Initial Release
                                                                 Land Launch User’s Guide Section 6


                                                    -

                                                                    6

                           N
                                                                10
                                                  1                                      2
                         Land Launch                                        4
                                                                            3
       1      Launch pad #1
                                                                8
       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)                                              5A
       6      Kerosene storage area                9                5
       7      Oxidizer storage area
       8      Pressure bottle storage
       9      Compressor station
      10      Air conditioning plant




                                                                        7

   Rail Connection
   to Site 42


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




Rev. Initial Release                   HPD-19000                                             6-15
                                                          Land Launch User’s Guide Section 6



Customer EGSE   Umbilical connection to the spacecraft (described in Section 5) is
Room (Bunker)   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)




6-16                               HPD-19000                         Rev. Initial Release
                                                                               Land Launch User’s Guide Section 6


                                                            -
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            Room 132
                                                Room 131            Room 132




                                                         Room 137
                                              Room 137




                                                                        Building 4
                              Building 4



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




Rev. Initial Release                          HPD-19000                                                    6-17
                                                                Land Launch User’s Guide Section 6


6.6 Cosmodrome Amenities

Visa and Access   Land Launch supports customers in obtaining entry visas to Russia by
Authorization     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           Land Launch supports the customer in obtaining customs clearances at all
Clearance         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.




6-18                                   HPD-19000                           Rev. Initial Release
                                                                        Land Launch User’s Guide Section 6


                                                         -
External               The customer is provided with local and international telephone/facsimile
Communications         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 Hotel accommodations are available at the Sputnik hotel
and Dining     (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


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                                                                         Land Launch User’s Guide Section 7




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.




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                                160



                                140



                                120
      Electric field, dB µV/m



                                100
                                                                  Glonass Receivers
                                 80
                                                                  1570-1630 MHz

                                 60



                                 40



                                 20



                                  0
                                      0   5000      10000     15000     20000      25000     30000         35000

                                                            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.




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7.1.3 Safety Design Considerations

Pressurized            Design and verification of spacecraft flight hardware pressurized vessels,
systems                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.




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                                                                    Land Launch User’s Guide Section 7




7.2 Spacecraft Verification Requirements

Flexibility to meet   The design verification processes and criteria defined below are
customer needs        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
                                                   safetyTest
             Test option                  Yield Ultimate level                Test success criteria
Qualification test (test of                1.0       1. 3     1.0         - No detrimental
dedicated article to ultimate loads)                                      deformation at 1.0
                                                              1.25        - No failure at 1.25
Proto-qualification test (test of          1.25      1.4      1.25        No detrimental
article used subsequently for flight or                                   deformation or
system test)                                                              misalignment
Qualification by analysis (test of         1.6       2.0      N/A         N/A
article not required)
Acceptance test (performed on each         1.1       1.3      1.1         No detrimental
flight article)                                                           deformation




Test-verified      Preliminary loads for quasi-static events are calculated using the Land
model required for Launch quasi-static load factors for ground handling, transportation, and
final CLA          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.




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                                                                           Land Launch User’s Guide Section 7




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.




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                                                                  Land Launch User’s Guide Section 7




 Sine vibration      The sine vibration test levels for qualification, proto-qualification and
 testing             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 Acceptance
           5 – 100             Zenit-2SLB        0.75 g            0.75 g         0.6 g
 (Longitudinal and Lateral) Zenit-3SLB           0.88 g            0.88 g         0.7 g
 Test frequency sweep rate (octaves/min)            2                 4             4




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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            120
                                              maximum expected acoustic
                                              environment

                     Proto-qualification      + 3 dB over levels of                     60
                                              maximum expected acoustic
                                              environment

                     Acceptance               Maximum expected acoustic                 60
                                              environment



Shock                The spacecraft should be compatible with the shock environment in Section 4
qualification        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.




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                                                                        Land Launch User’s Guide Section 7




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




    7-8                                    HPD-19000                                   Rev. Initial Release
                                                                          Land Launch User’s Guide Section 8




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.




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                                                                      Land Launch User’s Guide Section 8




Mission Team        The mission team is responsible for all mission analyses, operations
Roles and           planning, and launch campaign activities. On Land Launch the mission
Responsibilities    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-   At the beginning of the integration process, documentation needed for
mentation           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     Land Launch will create the spacecraft/Land Launch system Interface
Launch System       Control Document (ICD) that defines the technical interface require-
Interface Control   ments for the launch service. It will be based on the customer Interface
Document            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.


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




8-2                                     HPD-19000                           Rev. Initial Release
                                                                          Land Launch User’s Guide Section 8




Mission Integra-       Land Launch will develop the mission specific integration schedule based on
tion Schedule          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




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                                                                         Land Launch User’s Guide Section 8




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.




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                       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         Launch site operations planning will be conducted as necessary during
Planning                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
                                                                                    L - 12 days
                        Intersection Bay (Zenit-2SLB)
                        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




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                                                                        Land Launch User’s Guide Section 8




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   The SC and GSE arrive at the Yubileiny airfield within the confines of
and Transport        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           The customer will perform the necessary standalone spacecraft opera-
Processing           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.


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                                                                           Land Launch User’s Guide Section 8




                       •   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      After SC fueling and return to the integration bay, the following opera-
with Launch            tions are performed:
Vehicle Elements
                       •   Final preparations of the SC and spacecraft adapter.
in the PPF
Integration Bay        •   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 of the Zenit stages and the Block DM-SLB (Ze-
Autonomous             nit-3SLB missions) is completed before the start of combined operations.
Processing




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                                                                         Land Launch User’s Guide Section 8




Mating with Zenit    Integration of the Ascent Unit (AU) or PLU with the Zenit stages takes
Stages               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-      A TRR is held prior to the transfer to the pad. This review verifies the
ness Review (TRR)    readiness of the ILV, ground systems and launch base. The TRR author-
and Transfer to      izes the ILV transfer to the pad and subsequent final launch preparations.
the Launch Pad       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           Typical launch pad operations include:
Operations
                     •   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.




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                                                                         Land Launch User’s Guide Section 8




Launch Readiness       The LRR is held during launch day and prior to LV fueling, at approxi-
Review (LRR)           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 the ILV with compressed gasses begins 25 hours before launch.
Loading                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       The second “go” poll of launch management and the customer is con-
and Launch             ducted before disconnecting the transporter/erector from the launch vehi-
                       cle at approximately L-12 minutes.


Launch Control         During all phases of launch processing and flight, the launch control cen-
Center                 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.




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                                                                   Land Launch User’s Guide Section 8




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


8-10                                 HPD-19000                           Rev. Initial Release
                                                                         Land Launch User’s Guide Section 8




                       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               A distinguishing feature of Sea Launch as an international launch services
Review                 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.




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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:



                                   Sea Launch Company, LLC

                                    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: Paula.Korn@sea-launch.com




Rev. Initial Release                    HPD-19000                                           A-1
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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.




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       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




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       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.




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                   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 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)




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        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 location                                        (mm)
                               Xsc
                               Ysc
                             Zsc
                     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


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        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?




Rev. Initial Release                          HPD-19000                                          A-7
Land Launch User’s Guide




       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?




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       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




Rev. Initial Release                            HPD-19000                                       A-9
Land Launch User’s Guide




       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                     mx            mx          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               mx             mx         m           ft x          ft x        ft
            Spacecraft weight with handling fixture                                            kg                   lb
            Time required in processing facility                              days




A-10                                         HPD-19000                                    Rev. Initial Release
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       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
                 Temperature/humidity
                 Other
            Operations
                 Communication requirements
                 Storage
                 Security
                 Office space
                 Access (disabilities, equipment)
                 Other




Rev. Initial Release                          HPD-19000                                                        A-11
Land Launch User’s Guide




       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
                  Environmental monitoring
                  Other
            EGSE (size and weight)
                  Electrical power
                  Temperature/humidity
                  Other
            Operations
                  Communications
                  Storage
                  Security
                  Personnel accommodations
                  Office space
                  Other




A-12                                          HPD-19000                                       Rev. Initial Release
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       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
                 Environmental monitoring
                 Access
                 Other
            EGSE (size and weight)
                 Electrical power
                 Temperature/humidity
                 Other
            Operations
                 Communications
                 Storage
                 Security
                 Personnel accommodations
                 Office space
                 Other




Rev. Initial Release                          HPD-19000                                                       A-13
Land Launch User’s Guide




       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      Particulate (% obscuration)
                                                                (µg/cm2)
                                    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:




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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 with up to two access doors, RF window and a customer logo.
(PLF) and
                           •   SCA with compatible mechanical interfaces and separation system.
Spacecraft
                               Current standard offering adapters are SCA2624 and Zenit-2 adapter
Adapter (SCA)
                               for the Zenit-2SLB and SCA937, SCA1194, and SCA1666 for the
                               Zenit-3SLB. Contact Boeing Launch Services for adapters not listed.




Rev. Initial Release                         HPD-19000                                            B-1
Land Launch User’s Guide




Electrical                 RF interfaces
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                        Mission Management includes program-level services, such as
Management                     contracts, scheduling, program management, mission integration and
                               launch operations
                           •   Program Management Plan




B-2                                        HPD-19000                               Rev. Initial Release
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Meetings and               •   Customer kickoff meeting.
Reviews
                           •   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)




Rev. Initial Release                       HPD-19000                                                 B-3
Land Launch User’s Guide




Mission                    Analyses
Integration                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.




B-4                                        HPD-19000                               Rev. Initial Release
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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                     Land Launch will assist the customer with unloading the spacecraft and
Processing                  GSE from the aircraft at the cosmodrome airport. Land Launch will
Facilities (PPF)            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.




Rev. Initial Release                       HPD-19000                                                 B-5
Land Launch User’s Guide




PPF                        The following communication services will be available to customer
Communications             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                   •   Separate technical and facility electrical power
Services
                              •   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




B-6                                        HPD-19000                                 Rev. Initial Release
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Launch Vehicle                 Land Launch will provide facilities for personnel and spacecraft GSE
Integration                    in the LV integration facility, Area 42, for spacecraft check out and
Facility, Area 42              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                 Land Launch will provide LC facilities for customer personnel and
(LC) Facilities,               spacecraft GSE. Land Launch will provide the following services:
Area 45
                                    •    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             The following communication links are available within the launch
Communications             complex, area 45:
                                •       Intercom (including countdown net)
                                •       TV monitors and cameras (including launch video)
                                •       Dedicated telephone lines to international circuits


Launch Complex,            •    Full time security for the spacecraft and GSE
Area 45, Security
                           •    The facilities have 24-hr electronic monitoring for smoke and fire
                                detection




Rev. Initial Release                             HPD-19000                                           B-7
Land Launch User’s Guide




Environmental               Land Launch will maintain spacecraft environments within the limits
Controls                    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.




B-8                                        HPD-19000                               Rev. Initial Release
Land Launch User’s Guide




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.




Rev. Initial Release                       HPD-19000                                                  B-9

				
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