proton launch vehicle mission planner's guide
proton launch vehicle mission planner's guide
proton launch vehicle mission planner's guide
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Revision 4 March 1999<br />
PROTON LAUNCH VEHICLE<br />
MISSION PLANNER’S GUIDE<br />
REVISION NOTICE<br />
This document supersedes the Proton Launch Vehicle User’s<br />
Guide - Revision 3, Issue 1 dated February 1997<br />
DISCLOSURE OF DATA LEGEND<br />
These technical data are exempt from the licensing requirement<br />
of International Traffic in Arms Regulation (ITAR) under ITAR<br />
Part 125.4(b)(13) and are cleared for public release.<br />
LKEB-9812-1990<br />
© 1999 International Launch Services<br />
International Launch Services<br />
101 West Broadway, Suite 2000<br />
San Diego, California 92101 USA
Revision 4 March 1999<br />
Anatoli I. Kiselev<br />
General Director<br />
Khrunichev State Research and<br />
Production Space Center<br />
Anatoli K. Nedaivoda<br />
General Designer<br />
Salyut Design Bureau<br />
Khrunichev State Research and<br />
Production Space Center<br />
PROTON LAUNCH VEHICLE<br />
MISSION PLANNER’S GUIDE<br />
Wilbur C. Trafton<br />
President<br />
International Launch Services<br />
Eric F. Laursen<br />
Chief Engineer<br />
International Launch Services<br />
International Launch Services<br />
101 West Broadway, Suite 2000<br />
San Diego, California 92101 USA<br />
Page i
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
PREFACE<br />
The Proton Mission Planner’s Guide is intended to provide information to potential Customers and spacecraft<br />
suppliers, concerning spacecraft design criteria, Proton <strong>launch</strong> capability, available <strong>mission</strong> analysis and custom<br />
engineering support, documentation availability and requirements, and program planning. It is intended to serve as an<br />
aid to the planning of future <strong>mission</strong>s but should not be construed as a contractual commitment.<br />
The units of measurement referred to in this document are based on the International System of Units (SI), with<br />
English units given in parenthesis and all identified dimensions shown should be considered as approximate. In the<br />
event that one or more dimensions are critical to a specific payload intregration or processing operation, the SC<br />
Customer should obtain accurate dimensions from International Launch Services (ILS).<br />
This Guide will be updated or revised periodically. All comments and suggestions for additional information are<br />
hereby solicited and will be greatly appreciated.<br />
Those Customers wishing to receive revisions and updates to this manual, or who wish to submit comments or<br />
suggestions, are asked to kindly contact:<br />
International Launch Services<br />
101 West Broadway, Suite 2000<br />
San Diego, California 92101 USA<br />
Telephone: (619) 645-6400<br />
Facsimile: (619) 645-6500<br />
http:// www.lmco.com/ILS/<br />
Page ii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
REVISION HISTORY<br />
Revision Date Revision No. Change Description Approval<br />
15 December 1993 1, Issue 1 Eric Laursen<br />
Chief Engineer, LKEI<br />
December 1995 2, Issue 1 Eric Laursen<br />
Chief Engineer, LKEI<br />
Proton Division, ILS<br />
February 1997 3, Issue 1 Section 1<br />
• Updated Integration Schedule<br />
• Minor typographical corrections<br />
Section 2<br />
• Proton M fairing dimension update<br />
• Launch history update and corrections<br />
• Failure/Corrective Action update<br />
Section 3<br />
• Addition of Proton K/Block DM<br />
performance with use of standard kerosene<br />
Section 4<br />
• Ground Ops Instrumentation measurement<br />
capabilities update<br />
• Updated Proton LV radiated em<strong>mission</strong>s<br />
• Updated flight instrumentations capabilities<br />
• Updated flight loads environments<br />
• Updated flight acoustics<br />
Section 6<br />
• Updated Proton/BlockDM usable fairing<br />
envelopes with standard adapters<br />
Section 7<br />
• Updated Mission Integration schedule<br />
• Updated analysis, meetings and<br />
documentation schedules<br />
March 1999 4, Issue 1 • Complete rewrite/update of document to<br />
reflect flight measured environments,<br />
interfaces and performance<br />
• Addition of Proton M/Breeze M <strong>vehicle</strong><br />
data<br />
• Discussion of Baikonur payload processing<br />
and Launch operations facilities<br />
Eric Laursen<br />
Chief Engineer, LKEI<br />
Proton Division, ILS<br />
Eric Laursen<br />
Proton Chief Engineer,<br />
ILS<br />
Rich Waterman<br />
Manager Mission<br />
Development, ILS<br />
Page iii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
FOREWORD<br />
International Launch Services (ILS), is pleased to offer one of the most capable commercial <strong>launch</strong> <strong>vehicle</strong>s, and the<br />
most comprehensive <strong>launch</strong> services, available today. The Proton’s services are now available to worldwide Customers<br />
at a most competitive price.<br />
ILS is the exclusive marketing agent for commercial sales of the Proton <strong>launch</strong> <strong>vehicle</strong> worldwide, and is supported in<br />
its operations by full access to the incomparable technological expertise of its parent companies; Lockheed Martin<br />
Corporation (LMC), Khrunichev State Research and Production Space Center (KhSC), and Russian Space Complex<br />
Energia ( Energia). ILS provides customers with a single point of contact for all <strong>mission</strong> analyses, custom engineering,<br />
and <strong>launch</strong> support tasks involved in using the Proton <strong>launch</strong> <strong>vehicle</strong>. Both individually and collectively, the members<br />
of the ILS team are committed to providing the most cost-effective <strong>launch</strong> services available in the world-from initial<br />
program planning to successful spacecraft <strong>launch</strong>.<br />
This document provides performance, environments and interfaces for Proton <strong>vehicle</strong>s that have already been qualified<br />
to manage payloads up to 4.8 mt in mass. For SC exceeding 4.8 mt, parameters referred to in this document should be<br />
considered as PRELIMINARY and will be further defined/refined considering the geometry, mass properties and<br />
other physical characteristics of a particular spacecraft.<br />
Page iv
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
TABLE OF CONTENTS<br />
PREFACE ................................................................................................................................................... II<br />
REVISION HISTORY.................................................................................................................................III<br />
FOREWORD ...............................................................................................................................................IV<br />
TABLE OF CONTENTS................................................................................................................................ V<br />
LIST OF TABLES........................................................................................................................................XI<br />
LIST OF FIGURES......................................................................................................................................XI<br />
ABBREVIATIONS AND ACRONYMS........................................................................................................ XIX<br />
1. PROTON LAUNCH SERVICES.............................................................................................................. 1-1<br />
1.1 CONSTITUENT ILS COMPANIES........................................................................................................... 1-2<br />
1.2 ILS CONSTITUENT COMPANY EXPERTISE......................................................................................... 1-3<br />
1.3 ADVANTAGES OF USING THE PROTON LAUNCH VEHICLE .......................................................... 1-4<br />
1.4 VEHICLE DESCRIPTION ......................................................................................................................... 1-5<br />
1.5 GENERAL DESCRIPTION OF THE PROTON FAMILY ....................................................................... 1-7<br />
1.6 PROTON THREE - STAGE BOOSTER..................................................................................................... 1-8<br />
1.6.1 Proton First Stage ................................................................................................................................. 1-10<br />
1.6.2 Proton Second Stage.............................................................................................................................. 1-10<br />
1.6.3 Proton Third Stage................................................................................................................................ 1-10<br />
1.6.4 Proton Flight Control System ............................................................................................................... 1-11<br />
1.7 BLOCK DM FOURTH STAGE ................................................................................................................ 1-11<br />
1.8 BREEZE M FOURTH STAGE................................................................................................................. 1-15<br />
1.9 PAYLOAD FAIRINGS .............................................................................................................................. 1-18<br />
1.10 BAIKONUR INFRASTRUCTURE AND SERVICES ........................................................................... 1-23<br />
1.10.1 Baikonur Infrastructure ....................................................................................................................... 1-23<br />
1.10.2 Proton Launch Campaign.................................................................................................................... 1-25<br />
1.11 PLANNED ENHANCEMENTS ............................................................................................................. 1-27<br />
1.12 PROTON PRODUCTION AND OPERATIONS-SLIDES.................................................................... 1-27<br />
2. VEHICLE PERFORMANCE .................................................................................................................. 2-1<br />
2.1 OVERVIEW .................................................................................................................................................. 2-1<br />
2.2 PROTON LAUNCH SYSTEM CAPABILITIES......................................................................................... 2-1<br />
2.2.1 The Baikonur Launch Site....................................................................................................................... 2-2<br />
2.2.2 Launch Availability ................................................................................................................................. 2-2<br />
2.2.3 Payload Fairings and Adapters ................................................................................................................ 2-3<br />
2.2.4 Upper Stage Capabilities ......................................................................................................................... 2-3<br />
2.3 PROTON ASCENT PROFILE .................................................................................................................... 2-4<br />
2.3.1 Proton Booster Ascent ............................................................................................................................. 2-4<br />
2.3.2 Block DM Trajectory and Sequence ........................................................................................................ 2-8<br />
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2.3.3 Breeze M Trajectory Sequence ...............................................................................................................2-12<br />
2.3.4 Collision and Contamination Avoidance Maneuver............................................................................... 2-13<br />
2.4 PERFORMANCE GROUNDRULES....................................................................................................... 2-14<br />
2.4.1 Payload Systems Mass Definition.......................................................................................................... 2-14<br />
2.4.2 Payload Fairings.................................................................................................................................... 2-15<br />
2.4.3 Mission Analysis Groundrules ............................................................................................................... 2-15<br />
2.4.4 Performance Confidence Levels ............................................................................................................ 2-15<br />
2.5 DIRECT INJECTION LEO MISSIONS................................................................................................... 2-15<br />
2.6 GEOSYNCHRONOUS TRANSFER MISSIONS..................................................................................... 2-16<br />
2.6.1 Launch to Geosynchronous Transfer Orbit ............................................................................................ 2-16<br />
2.7 ORBIT INJECTION ACCURACY.............................................................................................................2-23<br />
2.8 SPACECRAFT ORIENTATION AND SEPARATION.............................................................................2-24<br />
2.9 LAUNCH VEHICLE TELEMETRY DATA ............................................................................................. 2-25<br />
2.10 MISSION OPTIMIZATION/ PERFORMANCE ENHANCEMENTS..................................................2-29<br />
2.10.1 Nonstandard Mission Designs ..............................................................................................................2-29<br />
2.10.2 Subsynchronous Transfer......................................................................................................................2-29<br />
2.10.3 Super Synchronous Transfer .................................................................................................................2-29<br />
3. SPACECRAFT ENVIRONMENTS .......................................................................................................... 3-1<br />
3.1 THERMAL/HUMIDITY ............................................................................................................................ 3-1<br />
3.1.1 Ground Thermal Environment................................................................................................................ 3-1<br />
3.1.2 Ascent ..................................................................................................................................................... 3-6<br />
3.1.3 Orbit ....................................................................................................................................................... 3-7<br />
3.1.4 Humidity ................................................................................................................................................ 3-7<br />
3.1.5 Air Impingement Velocity....................................................................................................................... 3-7<br />
3.2 CONTAMINATION ENVIRONMENT...................................................................................................... 3-8<br />
3.2.1 Ground Contamination Control.............................................................................................................. 3-8<br />
3.2.2 In Flight Contamination Control ............................................................................................................ 3-8<br />
3.3 PRESSURE .................................................................................................................................................. 3-9<br />
3.3.1 Payload Compartment Venting ............................................................................................................... 3-9<br />
3.4 MECHANICAL LOADS............................................................................................................................ 3-10<br />
3.4.1 Quasi-Static Loads................................................................................................................................ 3-10<br />
3.4.2 Sine and Random Vibration Loads ........................................................................................................ 3-14<br />
3.4.3 Acoustic Loads ...................................................................................................................................... 3-19<br />
3.4.4 Shock Loads.......................................................................................................................................... 3-21<br />
3.4.5 Environmental Test Requirements ........................................................................................................ 3-23<br />
3.5 ELECTROMAGNETIC COMPATIBILITY ............................................................................................ 3-24<br />
3.5.1 EMI Safety Margin (EMISM) ............................................................................................................. 3-24<br />
3.5.2 Radiated E<strong>mission</strong>s............................................................................................................................... 3-24<br />
3.5.3 RF Transmitter/Receiver Systems EMC............................................................................................... 3-29<br />
4. SPACECRAFT INTERFACES................................................................................................................. 4-1<br />
4.1 MECHANICAL INTERFACES .................................................................................................................. 4-1<br />
4.1.1 Structural Interfaces ................................................................................................................................ 4-1<br />
4.1.2 General SC Structural and Load Requirements ....................................................................................... 4-1<br />
4.1.3 Fairing Interfaces .................................................................................................................................... 4-2<br />
4.1.4 GN2/Dry Air Purge Option .................................................................................................................. 4-23<br />
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Issue 1, Revision 4, March 1, 1999<br />
4.2 ELECTRICAL INTERFACES .................................................................................................................. 4-24<br />
4.2.1 Airborne Interfaces................................................................................................................................ 4-24<br />
4.2.2 Launch Pad EGSE Interfaces................................................................................................................ 4-40<br />
4.2.3 Telemetry/Command Links.................................................................................................................. 4-41<br />
4.2.4 Electrical Grounding............................................................................................................................. 4-47<br />
4.2.5 Electrical Bonding ................................................................................................................................ 4-47<br />
4.2.6 SC/LV Lightning Protection ................................................................................................................. 4-47<br />
4.2.7 Static Discharge .................................................................................................................................... 4-47<br />
4.3 FITCHECK OF MECHANICAL/ELECTRICAL INTERFACES........................................................... 4-47<br />
5. MISSION INTEGRATION AND MANAGEMENT .................................................................................. 5-1<br />
5.1 MANAGEMENT PROVISIONS................................................................................................................. 5-1<br />
5.1.1 Key Personnel ......................................................................................................................................... 5-1<br />
5.1.2 Interface Control Document (ICD) ........................................................................................................ 5-1<br />
5.1.3 Schedule Monitoring............................................................................................................................... 5-1<br />
5.1.4 Documentation Control and Delivery ..................................................................................................... 5-4<br />
5.1.5 Meetings and Reviews ............................................................................................................................. 5-4<br />
5.1.6 DTRA Oversight ..................................................................................................................................... 5-8<br />
5.1.7 Quarterly Report..................................................................................................................................... 5-8<br />
5.1.8 Quality Provisions ................................................................................................................................... 5-8<br />
5.1.9 Launch License And Permits................................................................................................................... 5-8<br />
5.2 ILS DELIVERABLES.................................................................................................................................. 5-8<br />
5.2.1 ICD Development................................................................................................................................... 5-9<br />
5.2.2 Preliminary and Critical Design.............................................................................................................. 5-9<br />
5.2.3 Spacecraft Testing - Fitcheck/ Shock Test Plan, and Report ................................................................. 5-11<br />
5.2.4 Safety.................................................................................................................................................... 5-11<br />
5.2.5 Launch Campaign and Launch.............................................................................................................. 5-11<br />
5.2.6 Management and Reports...................................................................................................................... 5-11<br />
5.3 CUSTOMER DELIVERABLES................................................................................................................ 5-11<br />
5.3.1 ICD Development ................................................................................................................................ 5-13<br />
5.3.2 Preliminary and Critical Design............................................................................................................ 5-13<br />
5.3.3 Spacecraft Testing ................................................................................................................................. 5-13<br />
5.3.4 Required Safety Data and Certificates................................................................................................... 5-13<br />
5.3.5 Launch Campaign and Launch.............................................................................................................. 5-14<br />
5.4 SPECIFIC CUSTOMER RESPONSIBILITIES....................................................................................... 5-14<br />
5.4.1 Campaign Duration .............................................................................................................................. 5-14<br />
5.4.2 Spacecraft And Associated Ground Equipment ..................................................................................... 5-14<br />
5.4.3 Final Spacecraft Data............................................................................................................................ 5-15<br />
5.4.4 Spacecraft Readiness ............................................................................................................................. 5-15<br />
5.4.5 Removal of SC Support Equipment ...................................................................................................... 5-15<br />
5.4.6 Evaluation Of Launch Vehicle And Associated Services......................................................................... 5-15<br />
5.4.7 Spacecraft Propellants........................................................................................................................... 5-15<br />
5.4.8 Connectors ............................................................................................................................................ 5-15<br />
5.5 ILS SERVICES AND MATERIAL SPECIFICALLY EXCLUDED ........................................................ 5-16<br />
6. SPACECRAFT AND LAUNCH FACILITIES ........................................................................................... 6-1<br />
6.1 FACILITIES OVERVIEW ........................................................................................................................... 6-1<br />
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Issue 1, Revision 4, March 1, 1999<br />
6.1.1 Yubeleini Airport..................................................................................................................................... 6-1<br />
6.1.2 Building 92A-50 ..................................................................................................................................... 6-3<br />
6.1.3 Area 31 Facilities .................................................................................................................................... 6-3<br />
6.1.4 Building 92-1 and the Proton Launch Zone (Area 81) ............................................................................ 6-4<br />
6.1.5 Hotels ..................................................................................................................................................... 6-4<br />
6.2 SPACECRAFT PROCESSING FACILITIES............................................................................................. 6-4<br />
6.2.1 Facility 92A-50....................................................................................................................................... 6-5<br />
6.2.2 Area 31, Buildings 40/40D General Description ................................................................................... 6-14<br />
6.2.3 Area 31, Bldg 44 - General Description ................................................................................................ 6-20<br />
6.2.4 Building 254 (TBD).............................................................................................................................. 6-23<br />
6.2.5 Area 92, Building 92-1 - General Description....................................................................................... 6-23<br />
6.3 LAUNCH COMPLEX FACILITIES......................................................................................................... 6-24<br />
6.3.1 Area 81, Launch Pad 23 - General Description ..................................................................................... 6-24<br />
6.3.2 Facility Layout & Area Designations..................................................................................................... 6-26<br />
7. LAUNCH CAMPAIGN ......................................................................................................................................7-1<br />
7.1 ORGANIZATIONAL RESPONSIBILITIES.............................................................................................. 7-1<br />
7.1.1 Khrunichev.............................................................................................................................................. 7-1<br />
7.1.2 Strategic Rocket Forces ........................................................................................................................... 7-1<br />
7.1.3 Energia (Block DM <strong>launch</strong>es only) ......................................................................................................... 7-1<br />
7.1.4 ILS.......................................................................................................................................................... 7-2<br />
7.1.5 Spacecraft Customer ............................................................................................................................... 7-2<br />
7.2 CAMPAIGN ORGANIZATION ................................................................................................................. 7-2<br />
7.2.1 Contractual and Planning Organization................................................................................................... 7-2<br />
7.2.2 Organization During Combined Operations ............................................................................................ 7-3<br />
7.2.3 Planning Meetings................................................................................................................................... 7-4<br />
7.3 COUNTDOWN ORGANIZATION ............................................................................................................ 7-5<br />
7.4 ABORT CAPABILITY ................................................................................................................................. 7-6<br />
7.5 LAUNCH OPERATIONS OVERVIEW....................................................................................................... 7-7<br />
7.5.1 Launch Vehicle Processing .................................................................................................................... 7-12<br />
7.5.2 Spacecraft Preparations Through Arrival ............................................................................................... 7-12<br />
7.5.3 Area 31 (Buildings 40, 40D, 44) - Spacecraft Testing, Fueling, and Ascent Unit integration ................ 7-12<br />
7.5.4 Area 92 (Building 92A-50) - Spacecraft Testing, Fueling, and Ascent Unit Integration ........................ 7-14<br />
7.6 LAUNCH VEHICLE INTEGRATION (BUILDING 92-1) THRU LAUNCH PAD OPERATIONS.... 7-15<br />
7.7 LAUNCH PAD OPERATIONS................................................................................................................. 7-16<br />
8. PROTON LAUNCH SYSTEM ENHANCEMENTS.................................................................................. 8-1<br />
8.1 HIGH VOLUME PAYLOAD FAIRINGS................................................................................................... 8-3<br />
8.2 TANDEM LAUNCH SYSTEM ................................................................................................................... 8-5<br />
8.3 AVIONICS SYSTEM MASS UPGRADES.................................................................................................. 8-7<br />
8.4 NEW CYROGENIC UPPER STAGE ......................................................................................................... 8-7<br />
8.5 SUMMARY.................................................................................................................................................. 8-7<br />
A. PROTON LAUNCH SYSTEM HISTORY................................................................................................A-1<br />
A.1 BACKGROUND AND HISTORY..............................................................................................................A-1<br />
A.2 PROTON FLIGHT HISTORY ...................................................................................................................A-2<br />
A.3 DETAILED FLIGHT HISTORY ...............................................................................................................A-4<br />
A.4 FAILURES CAUSES AND CORRECTIVE ACTION.............................................................................A-10<br />
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Issue 1, Revision 4, March 1, 1999<br />
B1. QUALITY SYSTEM ...........................................................................................................................B1-1<br />
C1. GENERAL INFORMATION...............................................................................................................C1-1<br />
C2. INPUT DOCUMENTS....................................................................................................................... C2-2<br />
C3. INTERFACES.................................................................................................................................... C3-3<br />
C3.1 MECHANICAL INTERFACES .............................................................................................................C3-3<br />
C3.2 ELECTRICAL INTERFACES................................................................................................................C3-4<br />
C3.3 ENVIRONMENTAL INTERFACES......................................................................................................C3-5<br />
C3.4 FLIGHT DESIGN ..................................................................................................................................C3-6<br />
C3.5 OPERATIONS.........................................................................................................................................C3-7<br />
C3.6 OPERATIONS (CONTINUED).............................................................................................................C3-8<br />
D1. 1194AX-500 ADAPTER SYSTEM...................................................................................................... D1-1<br />
D1.1 INTRODUCTION................................................................................................................................. D1-1<br />
D1.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY ................................................. D1-1<br />
D1.2.1 Interface Ring Characteristics............................................................................................................ D1-1<br />
D1.2.2 Structural Capability ......................................................................................................................... D1-3<br />
D1.3 USABLE VOLUME ............................................................................................................................... D1-5<br />
D1.4 SEPARATION SYSTEM ..................................................................................................................... D1-11<br />
D1.5 ELECTRICAL INTERFACE .............................................................................................................. D1-14<br />
D1.6 INSTRUMENTATION ....................................................................................................................... D1-14<br />
D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS..................................................................... D1-15<br />
D2. 1194AX-625 ADAPTER SYSTEM .......................................................................................................D2-1<br />
D2.1 INTRODUCTION ................................................................................................................................. D2-1<br />
D2.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY.................................................. D2-1<br />
D2.2.1 Interface Ring Characteristics............................................................................................................ D2-1<br />
D2.2.2 Structural Capability ......................................................................................................................... D2-3<br />
D2.3 USABLE VOLUME ............................................................................................................................... D2-5<br />
D2.4 SEPARATION SYSTEM ..................................................................................................................... D2-21<br />
D2.5 ELECTRICAL INTERFACE .............................................................................................................. D2-24<br />
D2.6 INSTRUMENTATION ....................................................................................................................... D2-24<br />
D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS..................................................................... D2-25<br />
D3. 1666V-1000 ADAPTER SYSTEM .......................................................................................................D3-1<br />
D3.1 INTRODUCTION................................................................................................................................. D3-1<br />
D3.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY ................................................. D3-1<br />
D3.2.1 Interface Ring Characteristics............................................................................................................ D3-1<br />
D3.2.2 Structural Capability ......................................................................................................................... D3-3<br />
D3.3 USABLE VOLUME ............................................................................................................................... D3-5<br />
D3.4 SEPARATION SYSTEM ..................................................................................................................... D3-11<br />
D3.5 ELECTRICAL INTERFACE .............................................................................................................. D3-13<br />
D3.6 INSTRUMENTATION ....................................................................................................................... D3-13<br />
D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS...................................................................... D3-14<br />
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D4. 1666A-1150 ADAPTER SYSTEM ...................................................................................................... D4-1<br />
D4.1 INTRODUCTION................................................................................................................................. D4-1<br />
D4.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY ................................................. D4-1<br />
D4.2.1 Interface Ring Characteristics............................................................................................................ D4-1<br />
D4.2.2 Structural Capability ......................................................................................................................... D4-3<br />
D4.3 USABLE VOLUME ............................................................................................................................... D4-5<br />
D4.4 SEPARATION SYSTEM ..................................................................................................................... D4-11<br />
D4.5 ELECTRICAL INTERFACE .............................................................................................................. D4-13<br />
D4.6 INSTRUMENTATION ....................................................................................................................... D4-13<br />
D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS ..................................................................... D4-14<br />
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Issue 1, Revision 4, March 1, 1999<br />
LIST OF FIGURES<br />
FIGURE 1-1: ILS LAUNCH SERVICES CHARTER ...................................................................................................... 1-1<br />
FIGURE 1.1-1: ILS CORPORATE PARENTAGE.......................................................................................................... 1-2<br />
FIGURE 1.1-2: ILS/LKE TASK PARTITIONS............................................................................................................ 1-3<br />
FIGURE 1.2-1: LAUNCH EXPERIENCE ..................................................................................................................... 1-4<br />
FIGURE 1.3-1: BENEFITS TO THE SPACECRAFT DESIGNER AND OWNER................................................................ 1-5<br />
FIGURE 1.4-1: PROTON LAUNCH VEHICLES FLIGHT PROVEN HARDWARE.............................................................. 1-6<br />
FIGURE 1.6-1: PROTON K AND M MAJOR HARDWARE ELEMENTS ......................................................................... 1-9<br />
FIGURE 1.7-1: PROTON K/BLOCK DM MAJOR HARDWARE ELEMENTS................................................................ 1-12<br />
FIGURE 1.7-2: PROTON DM UPPER STAGE WITHIN ITS AERODYNAMIC SHROUDS .............................................. 1-13<br />
FIGURE 1.7-3: BLOCK DM AS IT APPEARS IN FLIGHT.......................................................................................... 1-14<br />
FIGURE 1.8-1: PROTON M/BREEZE M MAJOR HARDWARE ELEMENTS ................................................................ 1-16<br />
FIGURE 1.8-2: BREEZE M (WITH AUXILIARY PROPELLANT TANK) ....................................................................... 1-17<br />
FIGURE 1.8-3: THE BREEZE M AND BREEZE M CORE AS THEY APPEAR IN FLIGHT............................................ 1-17<br />
FIGURE 1.9-1: PROTON K AND M PAYLOAD FAIRINGS ......................................................................................... 1-19<br />
FIGURE 1.9-2: BLOCK DM PAYLOAD FAIRING FOR SINGLE SPACECRAFT ............................................................ 1-20<br />
FIGURE 1.9-3: BREEZE-M PAYLOAD FAIRING (STANDARD) FOR SINGLE SPACECRAFT....................................... 1-21<br />
FIGURE 1.9-4: BREEZE M PAYLOAD FAIRING (LONG) ..........................................................................................1-22<br />
FIGURE 1.10.1-1: OVERALL LAYOUT OF THE BAIKONUR COSMODROME.............................................................. 1-23<br />
FIGURE 1.10.1.2-1: BUILDING 92A-50 SPACE VEHICLE PROCESSING FACILITY ................................................... 1-24<br />
FIGURE 1.10.2-1: PROTON LAUNCH CAMPAIGN FOR 92A-50 ............................................................................... 1-26<br />
FIGURE S-1: TANK COMPONENT FABRICATION AT KHRUNICHEV USES AUTOMATED MACHINING CENTERS........ S-1<br />
FIGURE S-2: FIRST STAGE PROPELLANT TANK AUTOMATED WELD-UP................................................................. S-1<br />
FIGURE S-3: FIRST, SECOND, AND THIRD STAGE SUB-ASSEMBLIES AWAITING INTEGRATION.............................. S-2<br />
FIGURE S-4: PROTON INTERSTAGE COMPONENT FABRICATION............................................................................. S-2<br />
FIGURE S-5: FIRST STAGE BUILD UP ON PROTON FIXTURE................................................................................... S-3<br />
FIGURE S-6: INTERSTAGE JOINING SECOND AND THIRD STAGES........................................................................... S-3<br />
FIGURE S-7: RD-253 HIGH-PRESSURE ENGINE ON THE FIRST STAGE EXTERNAL FUEL TANK............................S-5<br />
FIGURE S-8: PROTON FINAL ASSEMBLY HALL AT KHRUNICHEV ............................................................................S-5<br />
FIGURE S-9: ASSEMBLED FIRST STAGE SHOWING HOLD DOWN POINTS AND AFT END SERVICES<br />
CONNECTORS .................................................................................................................................................. S-6<br />
FIGURE S-10: END-TO END TEST OF ASSEMBLED PROTON AT KHRUNICHEV FACTORY ....................................... S-6<br />
FIGURE S-11: THE BLOCK DM UNDERGOING FINAL ASSEMBLY AND TESTING AT THE ENERGIA PLANT IN<br />
KOROLEV, NEAR MOSCOW............................................................................................................................... S-7<br />
FIGURE S-12: THE COMPLETED BLOCK DM STAGE, BEFORE ATTACHMENT OF THE EXTERNAL SHROUD ........... S-7<br />
FIGURE S-13: FINISHED BLOCK DM STAGES IN THEIR AERODYNAMIC SHROUDS, AWAITING SHIPMENT TO<br />
BAIKONUR ....................................................................................................................................................... S-8<br />
FIGURE S-14: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING EXTERNAL VIEW .......................................... S-8<br />
FIGURE S-15: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING INTERNAL VIEW........................................... S-9<br />
FIGURE S-16: BREEZE M STAGE STRUCTURAL COMPONENTS IN MANUFACTURE ................................................ S-9<br />
FIGURE S-17: BREEZE M CORE STAGE MANUFACTURING .................................................................................. S-10<br />
FIGURE S-18: BREEZE M AVIONICS BAY ASSEMBLY ............................................................................................. S-10<br />
FIGURE S-19: BREEZE M STAGE FINAL ASSEMBLY .............................................................................................. S-11<br />
FIGURE S-20: SPACECRAFT ARRIVAL AT BAIKONUR.............................................................................................. S-11<br />
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FIGURE S-21: SPACECRAFT PROCESSING IN AREA 92 (SVPF).............................................................................. S-12<br />
FIGURE S-22: SPACECRAFT PROCESSING AND ENCAPSULATION IN AREA 31........................................................ S-13<br />
FIGURE S-23: THE COMPLETE PROTON LAUNCH VEHICLE BEING LIFTED ONTO ITS<br />
TRANSPORTER/ERECTOR .............................................................................................................................. S-13<br />
FIGURE S-24: LAUNCH VEHICLE INTERFACES AT A PROTON LAUNCH PAD, BAIKONUR....................................... S-14<br />
FIGURE S-25: PROTON ERECTION AT THE PAD ..................................................................................................... S-14<br />
FIGURE S-26: MOBILE SERVICE TOWER (MST).................................................................................................... S-15<br />
FIGURE S-27: LIFT-OFF OF THE PROTON K/BLOCK DM CARRYING A COMMERCIAL COMMUNICATIONS SATELLITE<br />
...................................................................................................................................................................... S-16<br />
FIGURE 2.2.1-1: THE BAIKONUR LAUNCH SITE, SHOWING AVAILABLE DIRECT INJECTION INCLINATIONS........... 2-2<br />
FIGURE 2.3.1-1: TYPICAL PROTON BOOSTER ASCENT ............................................................................................ 2-6<br />
FIGURE 2.3.1-2: TYPICAL PROTON BOOSTER ASCENT GROUND TRACK AND VACUUM IMPACT POINTS................. 2-6<br />
FIGURE 2.3.1-3: TYPICAL GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO THE<br />
SUPPORT ORBIT............................................................................................................................................... 2-7<br />
FIGURE 2.3.1-4: TYPICAL PROTON LOWER ASCENT ALTITUDE, INERTIAL VELOCITY, ACCELERATION, AND<br />
DYNAMIC PRESSURE ....................................................................................................................................... 2-7<br />
FIGURE 2.3.2-1: PROTON/BLOCK DM UPPER ASCENT.......................................................................................... 2-9<br />
FIGURE 2.3.2-2: BLOCK DM TWO-IMPULSE TRANSFER TO GEOSYNCHRONOUS TRANSFER ORBIT ..................... 2-10<br />
FIGURE 2.3.2-3: PROTON K/BLOCK DM UPPER ASCENT GROUND TRACK TO GSO ........................................... 2-11<br />
FIGURE 2.3.2-4: GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO A GSO ................................ 2-11<br />
FIGURE 2.3.3-1: TYPICAL BREEZE M FLIGHT PROFILE TO GEOSYNCHRONOUS TRANSFER ORBIT ...................... 2-13<br />
FIGURE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS<br />
TRANSFER ORBITS......................................................................................................................................... 2-17<br />
FIGURE 2.6.1-2: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS<br />
TRANSFER ORBITS......................................................................................................................................... 2-19<br />
FIGURE 2.8-1: SC SEPARATION ATTITUDE........................................................................................................... 2-25<br />
FIGURE 3.1.1.1-1A: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS<br />
TIMELINE (BLOCK DM) .................................................................................................................................. 3-4<br />
FIGURE 3.1.1.1-1B: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS<br />
TIMELINE (BREEZE M) ................................................................................................................................... 3-5<br />
FIGURE 3.1.1.2-1: SUPPLEMENTAL FAIRING AIR CONDITIONING SCHEMATIC (REPRESENTATIVE; DETAILED DESIGN<br />
CONDUCTED PER CUSTOMER REQUEST) .......................................................................................................... 3-6<br />
FIGURE 3.3.1-1: TYPICAL VENTING PROFILE DURING ASCENT ............................................................................. 3-9<br />
FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS.................................................................................. 3-11<br />
FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS (CONTINUED)........................................................... 3-12<br />
FIGURE 3.4.1.2-2: FLIGHT LIMIT ACCELERATIONS AT THE SC INTERFACE.......................................................... 3-13<br />
FIGURE 3.4.2-1: EQUIVALENT SINE LEVELS AT SPACECRAFT INTERFACE - FLIGHT ENVIRONMENT.................... 3-15<br />
FIGURE 3.4.2-2: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC IN CONTAINER<br />
AND SC ATTACHED TO ASCENT UNIT .......................................................................................................... 3-16<br />
FIGURE 3.4.2-3: TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER, TRANSPORTATION OF SC IN<br />
KHSC CONTAINER, TRANSPORTATION OF ASCENT UNIT .............................................................................. 3-17<br />
FIGURE 3.4.2-4: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC AND ASCENT<br />
UNIT ATTACHED TO L/V................................................................................................................................ 3-18<br />
FIGURE 3.4.2-4: TRANSPORTATION OF INTEGRATED PROTON LV........................................................................ 3-19<br />
FIGURE 3.4.3-1: MAX EXPECTED ACOUSTIC ENVIRONMENT (THIRD OCTAVE).................................................... 3-20<br />
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FIGURE 3.4.4-1: PYROSHOCK SPECTRUM AT ADAPTER/PAYLOAD INTERFACE......................................................3-22<br />
FIGURE 3.5.2-1A: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS<br />
(PROTON K/BLOCK DM) (TBC)................................................................................................................... 3-27<br />
FIGURE 3.5.2-1B: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS<br />
(PROTON M/BREEZE M) (TBC) ....................................................................................................................3-28<br />
FIGURE 3.5.2-2: LAUNCH VEHICLE AND LAUNCH PAD RADIATED SUSCEPTIBILITY LIMITS................................. 3-29<br />
FIGURE 4.1.1-1: LV COORDINATE SYSTEM ............................................................................................................ 4-1<br />
FIGURE 4.1.3.1-1A: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 3) .................. 4-4<br />
FIGURE 4.1.3.1-1B: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 3) .................. 4-5<br />
FIGURE 4.1.3.1-1C: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 3).................. 4-6<br />
FIGURE 4.1.3.1-2A: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 1 OF 3) ........................................................................................................................... 4-7<br />
FIGURE 4.1.3.1-2B: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 2 OF 3) ........................................................................................................................... 4-8<br />
FIGURE 4.1.3.1-2C: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 3 OF 3) ........................................................................................................................... 4-9<br />
FIGURE 4.1.3.1-3A: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT<br />
(SHEET 1 OF 3)............................................................................................................................................... 4-10<br />
FIGURE 4.1.3.1-3B: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT<br />
(SHEET 2 OF 3)............................................................................................................................................... 4-11<br />
FIGURE 4.1.3.1-3C: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT<br />
(SHEET 3 OF 3)............................................................................................................................................... 4-12<br />
FIGURE 4.1.3.1-4A: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 1 OF 3) ......................................................................................................................... 4-13<br />
FIGURE 4.1.3.1-4B: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 2 OF 3) ......................................................................................................................... 4-14<br />
FIGURE 4.1.3.1-4C: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME<br />
DIMENSIONS (SHEET 3 OF 3) ......................................................................................................................... 4-15<br />
FIGURE 4.1.3.1-5A: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 4) ....... 4-16<br />
FIGURE 4.1.3.1-5B: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 4) ....... 4-17<br />
FIGURE 4.1.3.1-5C: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 4)....... 4-18<br />
FIGURE 4.1.3.1-5D: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 4 OF 4)....... 4-19<br />
FIGURE 4.1.3.1-6A: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS<br />
(SHEET 1 OF 3)............................................................................................................................................... 4-20<br />
FIGURE 4.1.3.1-6B: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS<br />
(SHEET 2 OF 3)............................................................................................................................................... 4-21<br />
FIGURE 4.1.3.1-6C: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS<br />
(SHEET 3 OF 3)................................................................................................................................................4-22<br />
FIGURE 4.2.1.6-1: DRY LOOP FUNCTIONAL SCHEMATIC ..................................................................................... 4-25<br />
FIGURE 4.2.1.7-1A: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BLOCK DM FAIRING (TYPICAL)........ 4-32<br />
FIGURE 4.2.1.7-1B: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BREEZE M FAIRING (TYPICAL) ......... 4-33<br />
FIGURE 4.2.1.7-2: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1666 ADAPTER SYSTEM (TYPICAL)....... 4-34<br />
FIGURE 4.2.1.7-3: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1194 ADAPTER SYSTEM (TYPICAL) ...... 4-35<br />
FIGURE 4.2.1.7-4: INSTRUMENTATION DURING TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER......... 4-36<br />
FIGURE 4.2.1.7-5: INSTRUMENTATION DURING TRANSPORTATION OF SC IN KHSC CONTAINER....................... 4-36<br />
FIGURE 4.2.1.7-6: INSTRUMENTATION DURING TRANSPORTATION OF ASCENT UNIT......................................... 4-37<br />
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Issue 1, Revision 4, March 1, 1999<br />
FIGURE 4.2.1.7-7: INSTRUMENTATION DURING INTEGRATION OF ASCENT UNIT TO LV..................................... 4-37<br />
FIGURE 4.2.1.7-8: INSTRUMENTATION DURING TRANSPORTATION OF INTEGRATED PROTON LV....................... 4-38<br />
FIGURE 4.2.1.7-9: INSTRUMENTATION DURING ON-PAD OPERATIONS .............................................................. 4-39<br />
FIGURE 4.2.2-1: SPACECRAFT TO LAUNCH VEHICLE AND GROUND SYSTEMS ELECTRICAL INTERFACES............. 4-40<br />
FIGURE 4.2.3-1: SC TO BUNKER RF/ELECTRICAL INTERFACE BLOCK DIAGRAM ............................................... 4-46<br />
FIGURE 5.1.3-1A: BASELINE INTEGRATION SCHEDULE (NON-RECURRING PROGRAM)........................................ 5-2<br />
FIGURE 5.1.3-1B: BASELINE INTEGRATION SCHEDULE (RECURRING PROGRAM)................................................. 5-3<br />
FIGURE 6.1-1: BAIKONUR FACILITIES MAP............................................................................................................ 6-2<br />
FIGURE 6.2.1.1-1: BUILDING 92A-50 GENERAL ARRANGEMENT.......................................................................... 6-7<br />
FIGURE 6.2.1.4-1: BUILDING 92A-50 SPACECRAFT PROCESSING AND FUELING AREA ....................................... 6-10<br />
FIGURE 6.2.2-1: SPACECRAFT PROCESSING ZONE (AREA 31) .............................................................................. 6-15<br />
FIGURE 6.2.2.2-1: BUILDING 40, HALL 100 (COMMON HALL) LAYOUT............................................................... 6-16<br />
FIGURE 6.2.2.3-1: BUILDING 40D, FIRST FLOOR LAYOUT................................................................................... 6-17<br />
FIGURE 6.2.2.4-1: BUILDING 40D, SECOND FLOOR LAYOUT............................................................................... 6-19<br />
FIGURE 6.2.2.4-2: BUILDING 40D, THIRD FLOOR LAYOUT .................................................................................. 6-19<br />
FIGURE 6.2.3.1-1: BUILDING 44 (FILLING HALL) FLOOR PLAN .......................................................................... 6-21<br />
FIGURE 6.5.2.1-1: BUILDING 92-1....................................................................................................................... 6-24<br />
FIGURE 6.3.1-1: PROTON LAUNCH ZONE, AREA 81.............................................................................................. 6-25<br />
FIGURE 6.3.2.2-1: PROTON MOBILE SERVICE TOWER .......................................................................................... 6-27<br />
FIGURE 7.2.1-1: ORGANIZATION DURING LAUNCH CAMPAIGN ............................................................................ 7-3<br />
FIGURE 7.3-1: COUNTDOWN ORGANIZATION ........................................................................................................ 7-5<br />
FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50................................. 7-8<br />
FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50 (CONTINUED) ......... 7-9<br />
FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 .................................................. 7-10<br />
FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 (CONTINUED)........................... 7-11<br />
FIGURE 7.7-1: TYPICAL LAUNCH PAD OPERATIONS TIMELINE............................................................................ 7-17<br />
FIGURE 8-1: PROTON EVOLUTION OPTIONS........................................................................................................... 8-2<br />
FIGURE 8.1-1: PROTON LARGER PAYLOAD FAIRING CONCEPTS ............................................................................ 8-4<br />
FIGURE 8.2-1: BREEZE-M LAUNCH CONFIGURATION WITH TANDEM LAUNCH SYSTEMS (TLS).......................... 8-6<br />
FIGURE A.1-1: PROTON LAUNCH VEHICLE FAMILY...............................................................................................A-1<br />
FIGURE C3.1-1: SC PENDULUM MODEL...........................................................................................................C3-12<br />
FIGURE C3.1-2: SC SLOSH PROPERTIES DURING BALLISTIC FLIGHT AND AT SEPARATION ...............................C3-13<br />
FIGURE C3.1-3: PROPELLANT TANK GEOMETRY REQUIRED DATA ...................................................................C3-15<br />
FIGURE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D1-2<br />
FIGURE D1.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D1-4<br />
FIGURE D1.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM<br />
ADAPTER ...................................................................................................................................................... D1-6<br />
FIGURE D1.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM<br />
ADAPTER (SHEET 1 OF 4).............................................................................................................................. D1-7<br />
FIGURE D1.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM<br />
ADAPTER (SHEET 2 OF 4).............................................................................................................................. D1-8<br />
FIGURE D1.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM<br />
ADAPTER (SHEET 3 OF 4).............................................................................................................................. D1-9<br />
FIGURE D1.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM<br />
ADAPTER (SHEET 4 OF 4)............................................................................................................................ D1-10<br />
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Issue 1, Revision 4, March 1, 1999<br />
FIGURE D1.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 37 PIN ELECTRICAL<br />
CONNECTOR ............................................................................................................................................... D1-12<br />
FIGURE D1.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D1-13<br />
FIGURE D2.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D2-2<br />
FIGURE D2.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D2-4<br />
FIGURE D2.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM<br />
ADAPTER ...................................................................................................................................................... D2-6<br />
FIGURE D2.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM<br />
ADAPTER (SHEET 1 OF 4).............................................................................................................................. D2-7<br />
FIGURE D2.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM<br />
ADAPTER (SHEET 2 OF 4).............................................................................................................................. D2-8<br />
FIGURE D2.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM<br />
ADAPTER (SHEET 3 OF 4).............................................................................................................................. D2-9<br />
FIGURE D2.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM<br />
ADAPTER (SHEET 4 OF 4)............................................................................................................................ D2-10<br />
FIGURE D2.3-2A: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER ....................................................................................................................... D2-11<br />
FIGURE D2.3-2B: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 1 OF 4) (TBC) ................................................................................... D2-12<br />
FIGURE D2.3-2C: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 2 OF 4) (TBC) ................................................................................... D2-13<br />
FIGURE D2.3-2D: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 3 OF 4) (TBC) ................................................................................... D2-14<br />
FIGURE D2.3-2E: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 4 OF 4) (TBC) ................................................................................... D2-15<br />
FIGURE D2.3-3A: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER ....................................................................................................................... D2-16<br />
FIGURE D2.3-3B: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 1 OF 4) (TBC) ................................................................................... D2-17<br />
FIGURE D2.3-3C: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 2 OF 4) (TBC) ................................................................................... D2-18<br />
FIGURE D2.3-3D: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 3 OF 4) (TBC) ................................................................................... D2-19<br />
FIGURE D2.3-3E: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH<br />
1194AX X 625MM ADAPTER (SHEET 4 OF 4) (TBC) ................................................................................... D2-20<br />
FIGURE D2.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 37 PIN ELECTRICAL<br />
CONNECTOR ................................................................................................................................................D2-22<br />
FIGURE D2.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D2-23<br />
FIGURE D3.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D3-2<br />
FIGURE D3.2.2-1: CAPABILITY OF 1666V ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D3-4<br />
FIGURE D3.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM<br />
ADAPTER ...................................................................................................................................................... D3-6<br />
FIGURE D3.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM<br />
ADAPTER (SHEET 1 OF 4).............................................................................................................................. D3-7<br />
FIGURE D3.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM<br />
ADAPTER (SHEET 2 OF 4).............................................................................................................................. D3-8<br />
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Issue 1, Revision 4, March 1, 1999<br />
FIGURE D3.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM<br />
ADAPTER (SHEET 3 OF 4).............................................................................................................................. D3-9<br />
FIGURE D3.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM<br />
ADAPTER (SHEET 4 OF 4)............................................................................................................................ D3-10<br />
FIGURE D3.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 61 PIN ELECTRICAL<br />
CONNECTOR ............................................................................................................................................... D3-12<br />
FIGURE D4.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D4-2<br />
FIGURE D4.2.2-1: CAPABILITY OF 1666A ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D4-4<br />
FIGURE D4.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150<br />
ADAPTER ...................................................................................................................................................... D4-6<br />
FIGURE D4.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150<br />
ADAPTER (SHEET 1 OF 4).............................................................................................................................. D4-7<br />
FIGURE D4.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150<br />
ADAPTER (SHEET 2 OF 4).............................................................................................................................. D4-8<br />
FIGURE D4.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150<br />
ADAPTER (SHEET 3 OF 4).............................................................................................................................. D4-9<br />
FIGURE D4.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150<br />
ADAPTER (SHEET 4 OF 4)............................................................................................................................ D4-10<br />
FIGURE D4.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 61 PIN ELECTRICAL<br />
CONNECTOR ............................................................................................................................................... D4-12<br />
Page xvi
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
LIST OF TABLES<br />
TABLE 1.6-1: PROTON K AND PROTON M CHARACTERISTICS COMPARISON........................................................... 1-8<br />
TABLE 2.2-1: SUMMARY PROTON PERFORMANCE (PSM) TO REFERENCE ORBITS ................................................ 2-1<br />
TABLE 2.2.2-1: ALLOWABLE LAUNCH ENVIRONMENT CONSTRAINTS ..................................................................... 2-3<br />
TABLE 2.3.1-1: TYPICAL BOOSTER ASCENT EVENT TIMES....................................................................................... 2-5<br />
TABLE 2.3.2-1: TYPICAL BLOCK DM ATTITUDE MANEUVERS FOR GEOSYNCHRONOUS MISSION (90 O EAST<br />
LONGITUDE) ..................................................................................................................................................2-12<br />
TABLE 2.4.1-1: LAUNCH VEHICLE MISSION PECULIAR HARDWARE ..................................................................... 2-14<br />
TABLE 2.5-1: PROTON BOOSTER PERFORMANCE TO LOW EARTH ORBITS (DIRECT INJECTION, NO<br />
UPPER STAGE)................................................................................................................................................ 2-15<br />
TABLE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER<br />
ORBITS............................................................................................................................................................2-18<br />
TABLE 2.6.1-2: PROTON K/BLOCK DM THREE-BURN MISSION PERFORMANCE TO REPRESENTATIVE<br />
GEOSYNCHRONOUS TRANSFER ORBITS ..........................................................................................................2-18<br />
TABLE 2.6.1-3: PROTON K/BLOCK DM PARAMETRIC GEOSYNCHRONOUS TRANSFER PERFORMANCE DATE<br />
(2 BURN MISSION) .........................................................................................................................................2-20<br />
TABLE 2.6.1-4: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER ORBITS<br />
.......................................................................................................................................................................2-21<br />
TABLE 2.6.1-5: PROTON M/BREEZE M PARAMETRIC GEOSYNCHRONOUS TRANSFER PERFORMANCE DATA<br />
(CONFIGURATION 3).......................................................................................................................................2-22<br />
TABLE 2.7-1: BLOCK DM UPPER STAGE ORBIT INJECTION ACCURACIES.............................................................2-23<br />
TABLE 2.7-2: BREEZE M UPPER STAGE ORBIT INJECTION ACCURACIES ..............................................................2-23<br />
TABLE 2.8-1: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION I................................................................2-24<br />
TABLE 2.8-2: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION II ...............................................................2-24<br />
TABLE 2.8-3: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION III..............................................................2-24<br />
TABLE 2.9-1: FORMAT I - PRELIMINARY STATE VECTOR DATA PROVIDED FOLLOWING UPPER STAGE<br />
1 ST BURN.........................................................................................................................................................2-26<br />
TABLE 2.9-2: FORMAT II - TRANSFER ORBIT PARAMETERS FOLLOWING UPPER STAGE 1 ST BURN........................2-26<br />
TABLE 2.9-3: FORMAT III - PRELIMINARY VECTOR DATA AT SEPARATION EPOCH...............................................2-27<br />
TABLE 2.9-4: FORMAT IV - VECTOR DATA AT SEPARATION EPOCH.......................................................................2-27<br />
TABLE 2.9-5: FORMAT V - STATE VECTOR DATA AT SEPARATION EPOCH .............................................................2-28<br />
TABLE 2.9-6: SPACECRAFT SUPPLIED SEPARATION DATA .....................................................................................2-28<br />
TABLE 3.1.1-1: 3σ AMBIENT TEMPERATURES AT THE BAIKONUR COSMODROME................................................... 3-1<br />
TABLE 3.1.1-2: SPACECRAFT THERMAL ENVIRONMENT......................................................................................... 3-2<br />
TABLE 3.1.4-1: GROUND HUMIDITY ENVIRONMENT ............................................................................................. 3-7<br />
TABLE 3.2.1-1: GROUND CONTAMINATION ENVIRONMENT................................................................................... 3-8<br />
TABLE 3.4.1.1-1: GROUND LIMIT QUASISTATIC LOAD ENVIRONMENT-TRANSPORTATION AND HANDLING<br />
OPERATIONS .................................................................................................................................................. 3-10<br />
TABLE 3.4.4-1: SHOCK LOADS DURING TRANSPORTATION IN SC CONTAINER..................................................... 3-21<br />
TABLE 3.4.5.1-1: ACOUSTIC TEST REQUIREMENTS .............................................................................................. 3-23<br />
TABLE 3.4.5.2-1: SINE TEST REQUIREMENTS....................................................................................................... 3-23<br />
TABLE 3.5.2-1: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON K/BLOCK DM (TBC)............................ 3-25<br />
TABLE 3.5.2-2: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON M (TBC)............................................... 3-25<br />
TABLE 3.5.2-3: LAUNCH VEHICLE RF CHARACTERISTICS FOR BREEZE M (TBC) ............................................... 3-26<br />
TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS ........................ 4-27<br />
Page xvii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS (CONTINUED) ..4-28<br />
TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (TYPICAL).................. 4-29<br />
TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-30<br />
TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-31<br />
TABLE 4.2.3-1A: C-BAND RF LINK CHARACTERISTICS......................................................................................... 4-42<br />
TABLE 4.2.3-1B: KU-BAND RF LINK 1 CHARACTERISTICS .................................................................................. 4-43<br />
TABLE 4.2.3-1C: K-BAND RF LINK 2 CHARACTERISTICS..................................................................................... 4-44<br />
TABLE 4.2.3-1D: KU-BAND RF LINK 3 CHARACTERISTICS.................................................................................. 4-45<br />
TABLE 5.1.5.1-1A: BASELINE MEETING SCHEDULE FOR NON-RECURRING PROGRAM ......................................... 5-5<br />
TABLE 5.1.5.1-1B: BASELINE MEETING SCHEDULE FOR RECURRING PROGRAM .................................................. 5-6<br />
TABLE 5.2-1: ILS DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM................... 5-9<br />
TABLE 5.2.2-1: DESIGN REVIEW ANALYSES.......................................................................................................... 5-10<br />
TABLE 5.3-1: CUSTOMER DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM..... 5-12<br />
TABLE A.1-2: PROTON 50-LAUNCH MOVING AVERAGE.........................................................................................A-2<br />
TABLE A.2-1: PROTON LAUNCH RECORD SUMMARY (1970-1998).........................................................................A-3<br />
TABLE A.3-1: PROTON LAUNCH HISTORY..............................................................................................................A-4<br />
TABLE A.3-1A: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-5<br />
TABLE A.3-1B: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-6<br />
TABLE A.3-1C: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-7<br />
TABLE A.3-1D: PROTON LAUNCH HISTORY (CONTINUED) ....................................................................................A-8<br />
TABLE A.3-1E: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-9<br />
TABLE C3.1-1: SC MASS PROPERTIES ..................................................................................................................C3-9<br />
TABLE C3.1-1A: SC MASS PROPERTIES NEAR 0G.................................................................................................C3-9<br />
TABLE C3.1-1B: SC MASS PROPERTIES NEAR 1G.................................................................................................C3-9<br />
TABLE C3.1-1C: SC MASS PROPERTIES(DRY SPACECRAFT) ..............................................................................C3-10<br />
TABLE C3.1-2: DESCRIPTION OF LIQUID MASSES..............................................................................................C3-11<br />
TABLE C3.2-1: SC RF CHARACTERISTICS..........................................................................................................C3-16<br />
TABLE C3.2-2A: LVIJ1 UMBILICAL PIN ASSIGNMENTS......................................................................................C3-18<br />
TABLE C3.2.2B: LVIJ2 UMBILICAL PIN ASSIGNMENTS.......................................................................................C3-20<br />
TABLE C3.5-1: EGSE DESCRIPTION..................................................................................................................C3-22<br />
TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-23<br />
TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-24<br />
TABLE C3.5-2: FLUIDS/GASES REQUIREMENTS.................................................................................................C3-25<br />
TABLE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D1-1<br />
TABLE D1.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D1-11<br />
TABLE D1.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D1-14<br />
TABLE D2.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D2-1<br />
TABLE D2.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D2-21<br />
TABLE D2.5-1: STANDARD ELECTRICAL CONNECTORS..................................................................................... D2-24<br />
TABLE D3.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D3-1<br />
TABLE D3.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D3-11<br />
TABLE D3.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D3-13<br />
TABLE D4.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D4-1<br />
TABLE D4.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D4-11<br />
TABLE D4.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D4-13<br />
Page xviii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
ABBREVIATIONS AND ACRONYMS<br />
A CLA Coupled Loads Analysis<br />
A Ampere CLS Commercial Launch Services<br />
Ac Alternating Current cm Centimeter<br />
A/C Air Conditioning CRES Corrosion-Resistant Steel<br />
ADJ Attach, Disconnect, and Jettison CSO Complex Safety Officer<br />
AGE Aerospace Ground Equipment CT Command Transmitter<br />
ASCII American Standard Code for Information<br />
Interchange<br />
CV Computervision<br />
ASTR Air System, Thermal Regulation CW Continuous Wave<br />
atm Atmospheres CWA Controlled Work Area<br />
ATS Advanced Technology Satellite CX Complex<br />
AU Ascent Unit D<br />
AWG American Wire Gage DAS Data Acquisition System<br />
B DAT Digital Auto Tape<br />
Batt Battery dB Decibel(s)<br />
BOD Board of Directors DB Design Bureau<br />
BOL Beginning-of-Life dBm Decibel(s) Relative to 1 Milliwatt<br />
bpi Bit(s) per Inch dBW Decibel(s) Relative to 1 Watt<br />
BPSK Binary Phase Shift Key dc direct current<br />
Btu British Thermal Unit DEC Digital Equipment Corporation<br />
C DLF Design Load Factor<br />
O C Degree(s) Celsius DOF Degree(s) of Freedom<br />
CAB Customer Awareness Board DSN Deep Space Network<br />
CAD Computer-Aided Design Dstr Destructor<br />
CCAM Collison and Contamination Avoidance<br />
Manuever<br />
CCSCS Central Control Station of Communication<br />
System<br />
CCTV Closed-Circuit Television E<br />
DTSA Defense Technology Security Administration<br />
DUF Dynamic Uncertainty Factor<br />
CDR Critical Design Review e Eccentricity<br />
CERT Composite Electrical Readiness Test ECA Environmentally Controlled Area<br />
cg Center of Gravity ECS Environmentally Controlled System<br />
CIB Change Integration Board EED Electro-Explosive Device<br />
CIS Commonwealth of Independent States EHF Extreme High Frequency<br />
Page xix
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
EIA Electronics Industry Association ft Foot; Feet<br />
ELAN Electronic Library for Analysis of<br />
Nonconformances<br />
FTP Filt Transfer Protocol<br />
EM Electromagnetic FTS Flight Termination System<br />
EMC Electromagnetic Compatibility G<br />
EMI Electromagnetic Interference g Gravity or Gram<br />
EMK Extended Mission Kit GCS Guidance Commanded Shutdown<br />
EOL End-of-Life GEO Geosynchronous Orbit<br />
EPF Extended Payload Fairing GG Gas Generator<br />
EPT External Propellant Tank (Breeze M) GHe Gaseous Helium<br />
ERB Engineering Review Board GMM Geometric Mathematical Model<br />
ESABASE Thermal Control Geometric Math Modeling<br />
Code<br />
EUTELSAT European Telecommunications Satellite<br />
Organization<br />
G&N Guidance and Navigation<br />
GN&C Guidance, Navigation, and Control<br />
EVCF Eastern Vehicle Checkout Facility GN2 Gaseous Nitrogen<br />
EWR Eastern/Western Range Regulation GOWG Ground Operations Working Group<br />
ε Emissivity GSE Ground Support Equipment<br />
F GSO Geostationary Orbit<br />
O F Degree(s) Fahrenheit GTO Geosynchronous Transfer Orbit<br />
FAB Final Assembly Building GTR Gantry Test Rack<br />
FAST Flight Analogy Software Test GTS Ground Telemetry Station<br />
FCDC Flexible Confined Detonating Cord GTV Ground Transport Vehicle<br />
FCS Flight Control Subsystem H<br />
FDLC Final Design Loads Cycle HAIR High Accuracy Instrumentation Radar<br />
FLSC Flexible Linear-Shaped Charge HEO High-Energy (High-Eccentricity) Orbit<br />
FM Flight Model Frequency Modulation HEPA High-Efficiency Particulate Air<br />
FMAR Final Mission Analysis HiGTO High Energy Geoynchrenous Transfer Orbit<br />
FMEA Failure Modes and Effects Analysis HP Hewlett-Packard<br />
FMH Free Molecular Heating HPF Hazardous Processing Facility<br />
FOD Foreign Object Damage hr Hour(s)<br />
FOM Figure of Merit Hz Hertz<br />
FOTS Fiber-Optics Trans<strong>mission</strong> System I<br />
FPA Flight Plan Approval ICD Interface Control Document<br />
FPR Flight Performance Reserve ICWG Interface Control Working Group<br />
FSO Flight Safety Officer I/F Interface<br />
Page xx
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
IFR Inflight Retargeting LHe Liquid Helium<br />
IGES Initial Graphics Exchange Specification LKE Lockheed Khrunichev Energia<br />
IIP Instantaneous Impact Point LKEI Lockheed Khrunichev Energia International<br />
ILC Initial Launch Capability LM Lockheed Martin<br />
ILS International Launch Services LMC Lockheed Martin Corporation<br />
IMS Inertial Measurement Subsystem LMCLS Lockheed Martin Commercial Launch<br />
Services<br />
in. Inch(es) LN 2 Liquid Nitrogen<br />
INTELSAT International Telecommunications Satellite<br />
Organization<br />
LO 2<br />
Liquid Oxygen<br />
INU Inertial Navigation Unit LOB Launch Operations Building<br />
IRD Interface Requirements Document LR Load Ratio<br />
ISA Interstage Adapter LSC Linear Shaped Charge<br />
ISD Interface Scheduling Document LSTR Liquid System, Thermal Regulation<br />
I SP Specific Impulse LV Launch Vehicle<br />
ITA Intregrated Thermal Analysis LVMP Launch Vehicle Mission Peculiar<br />
ITAR International Traffic in Arms Regulations M<br />
K m Meter<br />
kg Kilogram(s) mA Milliamps<br />
KhSC Khrunichev State Research and Production<br />
Space Center<br />
MDRD Module-Level Design Requirements<br />
Document<br />
kHz Kilohertz MECO Main Engine Cutoff<br />
km Kilometer(s) MES Main Engine Start<br />
kN Kilonewton(s) MGSE Mechanical Ground Support Equipment<br />
kPa Kilopascal(s) MHz Megahertz<br />
kV Kilovolt(s) MICD Mechanical Interface Control Drawing<br />
L MIL-STD Military Standard<br />
LAE Liquid Apogee Engine min Minute(s)<br />
LAN Longitude of Acending Node MLV Medium Launch Vehicle<br />
lb Pound(s) mm Millimeter(s)<br />
lbf Pound(s)-Force MMH Monomethyl Hydrazine<br />
lb m Pound Mass MON-3 Mixed Oxides of Nitrogen<br />
LC Launch Conductor MOTR Multiple Object Tracking Radar<br />
LCC Launch Control Center MOU Memorandum of Understanding<br />
LEO Low-Earch Orbit MPDR Mission-Peculiar Design Review<br />
LH 2 Liquid Hydrogen MR Mixture Ratio<br />
Page xxi
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
MRB Material Review Board P<br />
MRR Mission Readiness Review Pa Pascal<br />
MRS Minimum Residual Shutdown PA Public Address<br />
MS&PA Mission Success and Product Assurance PCOS Power Changeover Switch<br />
MSPSP Mission System Pre<strong>launch</strong> Safety Package PDLC Preliminary Design Loads Cycle<br />
MST Mobile Service Tower PDR Preliminary Design Review<br />
MT Metric Ton PDRD Program-Level Design Requirements<br />
Document<br />
m/s Meters Per Second PFJ Payload Fairing Jettison<br />
μV Microvolt(s) PFM Protoflight Model<br />
mV Millivolt(s) PHA Preliminary Hazard Analysis<br />
MWG Management Working Group PHSF Payload Hazardous Servicing Facility<br />
N PLCP Propellant Leak Contingency Plan<br />
N Newton(s) PLCU Propellant Loading Control Unit<br />
NCAR National Center for Atmospheric Research PLF Payload Fairing<br />
N 2H 4 Hydrazine PMPCB Parts, Materials, and Processes Control Board<br />
NIOSH National Institute of Occupational Safety and<br />
Health<br />
PMR Preliminary Material Review<br />
NM Not Measured POD Program Office Directive<br />
NM Newton Meter PPF Payload Processing Facility<br />
nmi Nautical Mile(s) ppm Parts Per Million<br />
N 2O 4 Nitrogen Tetroxide psi/psf Pound(s) per Square Inch/ Pound(s) per<br />
Square Foot<br />
ns Nanosecond(s) psig Pound(s) per Square Inch, Gage<br />
NVR Nonvolatile Residue PSM Payload Systems Mass<br />
O PSS Payload Separation System<br />
O 2 Oxygen PST Product Support Team<br />
OASPL Overall Sound Pressure Level PSW Payload Systems Weight<br />
OCC/CS<br />
S<br />
Operator Control Console/Computer Subsystem PSWC Payload Systems Weight Capability<br />
Oct Octave(s) PTC Payload Test Conductor<br />
OIS Orbit Insertion Stage Payload Transport Canister<br />
ORD Operational Requirements Document PU Propellant Utilization<br />
OTM Output Transformation Matrix PVA Perigee Velocity Augmentation<br />
Ω Ohm(s) PVC Polyvinyl Chloride<br />
ω P Argument of Perigee P&W Pratt & Whitney<br />
Page xxii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Q SIU Servo-Inverter Unit<br />
q Dynamic Pressure SLC Space Launch Complex<br />
Spacecraft Launch Conductor<br />
QA Quality Assurance SOZ SOZ Unit or Auxiliary Control System<br />
QIC Quarter-Inch Cartridge SPA Spacecraft Processing Area<br />
R SPRB Space Program Reliability Board<br />
RAAN Right Ascension of Ascending Node SQEP Software Quality Evaluation Plan<br />
RCS Reaction Control System SQP Sequential Quadratic Programming<br />
R&D Research and Development SRB Solid Rocket Booster<br />
RDX Research Department Explosive SRF Strategic Rocket Forces<br />
RF Radio Frequency SRM Solid Rocket Motor<br />
RM Room SRPSC State Research and Production Space Center<br />
Roi Return on Investment SRR System Requirements Review<br />
RP-1 Rocket Propellant 1 (Kerosene) Sta Station<br />
RSA Russian Space Agency STC Satellite Test Center<br />
RSC Rocket Space Complex STDN Spaceflight Tracking and Data Network<br />
RSC Range Safety Console STM Structural Test Model<br />
RSF Russian Space Force STS Space Transportation System<br />
RTS Remote Tracking Station SVPF Space Vehicle Processing Facility<br />
S T<br />
s Second(s) tar Tape Archive (File Format)<br />
S/A Safe and Arm TBD To Be Determined<br />
SAEF Spacecraft Assembly and Encapsulation Facility TBS To Be Supplied<br />
SASU Safe/Arm and Securing Unit T&C Telemetry & Command<br />
SC Spacecraft TC Telecommand<br />
SCAPE Self-Contained Atmospheric Protective<br />
Ensemble<br />
Test Conductor<br />
SCA Spacecraft Adapter TCD Terminal Countdown Demonstration<br />
SDP Software Documentation Plan TCO Thrust Cutoff<br />
SDRC Structural Dynamics Research Corporation TDRSS Tracking and Data Rela Satellite System<br />
sec Seconds TIM Technical Interchange Meeting<br />
SEPP Systems Effectiveness Program Plan Tlm Telemetry<br />
SFC Spacecraft Facility Controller TLS Tandem Launch System<br />
SFTS Secure Flight Termination System TMM Thermal Mathematical Model<br />
SHA System Hazard Analysis TRM Tension Release Mechanism<br />
SHU Space Head Unit TSB Technical Support Building<br />
SIL Systems Integration Laboratory TVC Thrust Vector Control<br />
Page xxiii
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
TVCF Transportable Vehicle Checkout Facility<br />
TWG Technical Working Group<br />
3-D Three-Dimensional<br />
U<br />
UDMH Unsymmetrical Dimethylhydrazine<br />
UHF Ultra-High Frequency<br />
UPS Uninterruptable Power Systems<br />
U.S. United States<br />
US Upper Stage-Proton Fourth Stage<br />
UT Umbilical Tower<br />
V<br />
V Volt(s) or Velocity<br />
Vac Volt(s) Alternating Current<br />
Vdc Volt(s) Direct Current<br />
VDD Version Description Document<br />
VHF Very High Frequency<br />
VSWR Voltage Standing-Wave Ratio<br />
VTF Vertical Test Facility<br />
W<br />
W Watt(s)<br />
WB West Bay<br />
WDR Wet Dress Rehearsal<br />
W/M 2 Watts Per Square Meter<br />
WSR Weather Service Radar<br />
X<br />
Xdcr Transducer<br />
Page xxiv
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1. PROTON LAUNCH SERVICES<br />
International Launch Service’s (ILS) <strong>launch</strong> support services equal or exceed in comprehensiveness and quality those<br />
services available from any other commercial <strong>launch</strong> service organization. These services include system integration,<br />
supply of the Proton <strong>launch</strong> <strong>vehicle</strong>, custom engineering services and <strong>mission</strong> analysis, insurance brokering, ground<br />
and air transport, Government export license assistance, <strong>launch</strong> site spacecraft integration and testing, spacecraft and<br />
<strong>launch</strong> <strong>vehicle</strong> integration, <strong>launch</strong> site support and security services, <strong>launch</strong> of the spacecraft, and post <strong>launch</strong> <strong>mission</strong><br />
support (see Figure 1-1).<br />
The management approach to the provision of these services has been developed to ensure efficient task completion,<br />
with essential focus on Customer satisfaction. ILS functions as a prime contractor to manage all tasks associated with<br />
the supply of the <strong>launch</strong> <strong>vehicle</strong> and associated spacecraft <strong>launch</strong> services, including all required liaison with various<br />
United States, Russian, and other government organizations and agencies, as well as accommodation of any special<br />
Customer requirements. ILS will support all spacecraft preparation activities, oversee the integration of the spacecraft<br />
with the Proton, and conduct the spacecraft <strong>launch</strong>. The Customer need interact solely with ILS for full support in all<br />
aspects of the <strong>launch</strong>.<br />
Figure 1-1: ILS Launch Services Charter<br />
Lockheed<br />
Martin<br />
Corporation<br />
Khrunichev<br />
SRPSC<br />
RSC<br />
Energia<br />
ILS<br />
System engineering<br />
Mission planning<br />
Interface data<br />
Custom engineering<br />
Insurance brokering<br />
Export license support<br />
Air transport brokering<br />
Host services and logistics support<br />
Test planning/requirements<br />
Spacecraft Integration and Test<br />
Site GSE and consumables<br />
Launch <strong>vehicle</strong>/adapter/fairing<br />
Launch <strong>vehicle</strong> integration<br />
Payload <strong>launch</strong><br />
Post <strong>launch</strong> data analysis<br />
Page 1-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1.1 CONSTITUENT ILS COMPANIES<br />
In 1992, the Space Systems Division of Lockheed Missiles & Space Co., Inc. established the Lockheed Commercial<br />
Space Company, Inc. (U.S.). This company, along with the Russian organizations now known as the Khrunichev State<br />
Research and Production Space Center and the Rocket Space Complex Energia, signed an agreement to form a joint<br />
venture, Lockheed-Khrunichev-Energia International, Inc. (LKE). This venture, which has since become the Proton<br />
Division of ILS, which in turn reports to the Astronautics Division of Lockheed Martin Corporation (LMC), has the<br />
charter to provide commercial <strong>launch</strong> services based on the Proton <strong>launch</strong> <strong>vehicle</strong> to customers worldwide. The<br />
Corporate parentage of ILS is shown in Figure 1.1-1.<br />
Figure 1.1-1: ILS Corporate Parentage<br />
LMC Khrunichev Energia<br />
LM<br />
Astronautics<br />
ILS<br />
This joint venture has been approved by both the U.S. and Russian governments. Lockheed Martin is the United States<br />
leading company in all aspects of spacecraft/<strong>launch</strong> <strong>vehicle</strong> integration, <strong>launch</strong> <strong>vehicle</strong> design, development and<br />
manufacture, and <strong>launch</strong> site operations. Khrunichev (along with its subsidiary, the Salyut Design Bureau) is the<br />
Russian designer and manufacturer of the first stage, second stage, and third stage of the Proton space <strong>launch</strong> <strong>vehicle</strong> as<br />
well as the Breeze-M and other upper stages. Energia is the Russian designer and manufacturer of the Block DM<br />
fourth stage of the Proton, and is the largest producer of <strong>launch</strong> <strong>vehicle</strong>s in Russia. ILS has access to all resources of the<br />
constituent companies required to fulfill spacecraft <strong>launch</strong> campaign requirements.<br />
Lockheed Martin, Khrunichev, and Energia all function as subcontractors, reporting to ILS, during the execution of a<br />
Customer’s <strong>launch</strong> services contract. A memorandum of understanding (MOU) that delineates the responsibilities of<br />
each of the companies is in place. (Figure 1.1-2) Constituent senior management of each of the companies have<br />
approved the overall management approach for ILS, and are each represented on the ILS Board of Directors.<br />
The personnel, hardware resources, and facilities needed to support customer <strong>launch</strong> programs are in place and ready<br />
for immediate activation, as needed.<br />
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ILS personnel are responsible for all Customer interface activities, and for the coordination of all activities of the<br />
constituent companies so that contract objectives are met. They are also responsible for all program system<br />
engineering, custom engineering, <strong>mission</strong> analysis, program integration and program management, liaison with all<br />
government agencies, and liaison with the world’s financial and insurance markets. Khrunichev is responsible for<br />
manufacturing the Proton first, second, and third stages, manufacturing the Proton’s Breeze M fourth stage,<br />
conducting <strong>mission</strong> analysis, and providing Baikonur services. Energia is responsible for manufacturing the Proton's<br />
Block DM fourth stage, providing the Block DM timeline, dynamics, and other data for <strong>mission</strong> analysis, and<br />
providing Baikonur services.<br />
Figure 1.1-2: ILS/LKE Task Partitions<br />
Lockheed Martin<br />
Corporation<br />
Launch program integration<br />
System management<br />
Insurance liaison<br />
Customer reporting<br />
Mission analysis<br />
Third party liaison<br />
Post <strong>launch</strong> services<br />
Performance data liaison<br />
Interface data liaison<br />
Custom engineering services<br />
Launch processing liaison<br />
1.2 ILS CONSTITUENT COMPANY EXPERTISE<br />
ILS/LKE<br />
Khrunichev<br />
(KhSC)<br />
First, second, and third stage production<br />
Preparation of stages for <strong>launch</strong><br />
Proton-induced environment design criteria<br />
Interface services<br />
Launch <strong>vehicle</strong> GSE, consumables,and <strong>launch</strong><br />
services<br />
Coordination of Baikonur Cosmodrone services<br />
CIS government liaison<br />
Subcontract management<br />
Breeze M production<br />
Customerliaison/report<br />
Subcontract direction<br />
U.S. Government liaison<br />
RSC Energia<br />
Block DM production<br />
Block DM preparation services<br />
for <strong>launch</strong><br />
Proton spacecraft interface<br />
design criteria<br />
Proton-induced environment design<br />
criteria<br />
Russian subcontract management<br />
Interface services<br />
Launch <strong>vehicle</strong> GSE,<br />
consumables, and <strong>launch</strong> services<br />
Subcontract management<br />
The joint venture of Lockheed Martin Corporation, Khrunichev and Energia offers the most capable, comprehensive,<br />
and cost-effective spacecraft <strong>launch</strong> services in the world.<br />
ILS, with access to the technological expertise and resources of Lockheed Martin Corporation, Khrunichev and<br />
Energia, can provide all necessary resources required to support a spacecraft <strong>launch</strong>. ILS provides spacecraft Customers<br />
with a single point of contact for all <strong>mission</strong> engineering and <strong>launch</strong> support tasks using the Proton <strong>launch</strong> <strong>vehicle</strong>.<br />
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The constituent companies of ILS have more spacecraft and <strong>launch</strong> <strong>vehicle</strong> expertise than any other organizations<br />
providing <strong>launch</strong> services today. (Figure 1.2-1) Lockheed Martin has built more than half of all Western satellites<br />
flown, and has designed, built and <strong>launch</strong>ed the Atlas and Titan <strong>launch</strong> <strong>vehicle</strong>s, as well as the Agena and Centaur<br />
upper stages, and the Polaris, Poseidon and the Trident strategic defense missiles. Khrunichev, in addition to designing<br />
and building the Proton, is responsible for subcontract management for all of the <strong>launch</strong> <strong>vehicle</strong> subsystems.<br />
Khrunichev personnel participate in all Proton integration and <strong>launch</strong> operations at the Baikonur Cosmodrome.<br />
Khrunichev has also produced a variety of missile systems for the Russian government, as well as the Salyut, Mir, and<br />
Almaz orbital stations, and several different orbital return capsule designs. Energia has provided all of the fourth stages<br />
for the Proton from the inception of the program until 1999, and has also been responsible for significant subcontract<br />
management. Energia is the designer and builder of the R-7 series of space <strong>launch</strong> <strong>vehicle</strong>s, including the Soyuz and<br />
Molniya <strong>launch</strong>ers, of which more than 1300 have been <strong>launch</strong>ed to date. During the 1980’s they developed both the<br />
Energia heavy lift <strong>launch</strong> <strong>vehicle</strong> and the Buran space shuttle, in addition to significant portions of the Mir space<br />
station hardware. The constituent companies of ILS have the necessary expertise to successfully and efficiently support<br />
any <strong>launch</strong> campaign.<br />
Figure 1.2-1: Launch Experience<br />
1950s<br />
1950s<br />
1960s 1970s 1980s 1990s 2000s<br />
Lockheed Martin<br />
Titan I, II, III, , and IV <strong>launch</strong>es<br />
Atlas <strong>launch</strong>es<br />
Agena flights<br />
Polaris/Poseidon/Trident <strong>launch</strong>es<br />
Space Shuttle <strong>launch</strong>es<br />
Mission analysis and system engineering<br />
Khrunichev<br />
200+ Proton <strong>launch</strong>es<br />
Mission analysis and system engineering<br />
Spacecraft/<strong>launch</strong> <strong>vehicle</strong> integration<br />
RSC Energia<br />
Spunik, Vostok, Voskhod, Soyuz, Molniya <strong>launch</strong>es<br />
Proton fourth stage production<br />
Spacecraft/<strong>launch</strong> <strong>vehicle</strong> integration<br />
1960s 1970s 1980s 1990s 2000s<br />
1.3 ADVANTAGES OF USING THE PROTON LAUNCH VEHICLE<br />
The use of the Proton <strong>launch</strong> <strong>vehicle</strong> provides several significant advantages that result in operational and revenue<br />
generating benefits for the Customer (Figure 1.3-1).<br />
ILS provides full system engineering and <strong>mission</strong> analysis services and mechanical/electrical interface coordination.<br />
The proven procedures for Proton <strong>launch</strong> <strong>vehicle</strong> operations, backed by personnel with extensive experience in these<br />
procedures, provide for efficient and trouble-free <strong>launch</strong> campaigns.<br />
The considerable lift capability of the Proton, combined with the multiple restart capability of both of the fourth<br />
stages, provides the Customer with <strong>mission</strong> flexibility and maximized payload capacity to orbit. This results in unique<br />
<strong>mission</strong> design options, including delivery of spacecraft directly to geostationary orbit by the Proton. Spacecraft apogee<br />
fuel may be dedicated to extended <strong>mission</strong> life, because inclination reduction and orbit circularization are<br />
accomplished by the Proton’s 4th stage.<br />
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The ability of ILS to meet the full range of Customer requirements guarantees each <strong>launch</strong> campaign will be<br />
conducted to the Customer’s satisfaction. This approach to satisfying key requirements is shown in Figure 1.3-1. ILS’s<br />
overriding concern in providing commercial <strong>launch</strong> services is the careful coordination of our company’s resources and<br />
capabilities with the Customer’s detailed requirements, so as to allow ILS to tailor each <strong>launch</strong> campaign to meet the<br />
individual Customer’s unique needs. This ensures that each campaign will proceed in an efficient manner toward a<br />
successful, on-time <strong>launch</strong> to the precise orbit required.<br />
Figure 1.3-1: Benefits To The Spacecraft Designer And Owner<br />
Full <strong>mission</strong> analysis support by all constituent companies<br />
Access to proven <strong>launch</strong> <strong>vehicle</strong> manufacturing and support capability<br />
Use of very capable integration and <strong>launch</strong> support infrastructure<br />
Use of the massive lift capacity of the Proton <strong>launch</strong> <strong>vehicle</strong><br />
Use of the restart capability of the Proton fourth stage for final orbit insertion<br />
1.4 VEHICLE DESCRIPTION<br />
The Proton <strong>launch</strong> system is designed and manufactured by International Launch Services partners Khrunichev State<br />
Research and Production Space Center (KhSC) and Rocket Space Complex (RSC) Energia, both of Russia. Based on<br />
<strong>vehicle</strong>s in production today, the Proton is one of the most capable commercial expendable <strong>launch</strong> systems in use,<br />
delivering intermediate and heavy spacecraft to a wide range of orbits from its <strong>launch</strong> base at the Baikonur<br />
Cosmodrome in Kazakhstan.<br />
The Proton <strong>vehicle</strong> used for commercial <strong>launch</strong>es can be provided in either a 3-stage or 4-stage configuration to meet<br />
the needs of Low Earth Orbit and High-Energy <strong>launch</strong> <strong>mission</strong>s. The 3-stage versions of Proton, designated Proton-K<br />
and Proton-M, are designed to lift very heavy spacecraft systems into Low Earth Orbit. The 4-stage version of Proton,<br />
using either the Block DM or the Breeze M upper stage, is designed to meet high-energy <strong>launch</strong> <strong>mission</strong> injection<br />
requirements such as those for the Geosynchronous Transfer Mission. Multiple payload fairing and adapter systems are<br />
available in order to accommodate most commercial satellite <strong>launch</strong> <strong>mission</strong> needs.<br />
This section provides a general description of the Proton <strong>launch</strong> <strong>vehicle</strong>, summarizing the relevant capabilities of the<br />
Proton-K/Block DM <strong>launch</strong> system and the Proton M/Breeze M <strong>launch</strong> system.<br />
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Figure 1.4-1: Proton Launch Vehicles Flight Proven Hardware<br />
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1.5 GENERAL DESCRIPTION OF THE PROTON FAMILY<br />
The Proton is currently available as three-stage Proton K and Proton M models and as the four-stage Proton K/ Block<br />
DM , Proton K/Breeze M, Proton M/Block DM, and Proton M/Breeze M models. A variety of supplemental orbital<br />
propulsion units, in a range of capabilities, can be used with either the three-stage or the four-stage Proton. In<br />
addition, there are multiple fairing designs presently qualified for flight, including "Standard Commercial Payload<br />
Fairings" developed specifically to meet the needs of Western customers. Tandem <strong>launch</strong> and multi-spacecraft<br />
dispenser capabilities are in development. Later sections provide more detailed information on these hardware<br />
elements and their operation.<br />
The lower three stages of the Proton are produced by the Khrunichev State Research and Production Space Center<br />
(KhSC) plant in Moscow. Production of the Breeze-M stage is conducted by KhSC. Production of the Block DM<br />
fourth stage is carried out by Russian Space Complex (RSC) Energia, also in Moscow. Production capacity for the<br />
commercial Proton is approximately twelve <strong>vehicle</strong>s per year.<br />
General specifications for both versions of the Proton are similar. Overall height of the <strong>vehicle</strong> in either configuration<br />
is approximately 61 m (200 ft), while the diameter of the second and third stages, and of the first stage core tank, is 4.1<br />
m (13.5 ft). Maximum diameter of the first stage, including the outboard fuel tanks, is 7.4 m<br />
(24.3 ft). The Block DM fourth stage, when present, has an external diameter of 3.7 m (12.1 ft). The Breeze M, when<br />
present, has a diameter of 4.1 m (13.5 ft). Total mass of the Proton at <strong>launch</strong> is approximately 651,500 kg (1,524,000<br />
lbm).<br />
Maximum performance capability (Payload Systems Mass, or PSM) of the Proton in the configurations of principal<br />
interest to Western customers, is as follows:<br />
Proton K: Proton M:<br />
LEO (200 km circ.): 19.76 metric tons 21.0 metric tons<br />
Proton K/Block DM: Proton M/Breeze M:<br />
GTO: 4.9 metric tons 5.5 metric tons<br />
GSO: 1.9 metric tons 2.92 metric tons<br />
These figures assume the use of the Standard Commercial Payload Fairings, and optimum event sequences. Specific<br />
performance may differ from the performance cited, if:<br />
a) Use is made of adapters or fairings differing in mass or other characteristics relevant to performance<br />
b) Earlier fairing separation timing is found to be acceptable<br />
c) Special modifications are made to the Proton or its operations<br />
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1.6 PROTON THREE - STAGE BOOSTER<br />
The Proton K and M are series-staged <strong>vehicle</strong>s consisting of three stages. An isometric view of the Proton K and M,<br />
showing the relationships among the major hardware elements, is provided in Figure 1.6-1. All three stages use<br />
nitrogen tetroxide (N 2O 4) and unsymmetrical dimethylhydrazine (UDMH) as propellants. A summary comparison of<br />
the characteristics of the Proton K and Proton M <strong>launch</strong> systems are shown in Table 1.6-1.<br />
A comparison of the characteristics of the Proton M/Breeze M with those of the Proton K/Block DM is shown in<br />
Table 1.6-1.<br />
Table 1.6-1: Proton K And Proton M Characteristics Comparison<br />
Launch Vehicle/Upper Stage Proton K/Block DM Proton M/Breeze M<br />
Launch Vehicle Liftoff Mass, mt 700 700<br />
Support orbit Payload Mass, mt<br />
hcirc.= 200 km, I = 51,6 O<br />
Geostationary Orbit Payload Mass, mt<br />
hcirc.= 36000 km, I = 0 O<br />
Geostationary Transfer Orbit Payload Mass, t<br />
ha = 36000 km, I = 7…51,6 O<br />
20.7 22.0<br />
Max. 1.9-2.1 Max. 3.0<br />
3.5…6.5 5.0…7.8<br />
Payload Bay Volume, m 3 65 (Std PLF) 100 (Std PLF)<br />
Launch Vehicle Structure Mass, mt<br />
First Stage<br />
Second Stage<br />
Third Stage<br />
Upper (fourth) Stage<br />
Propulsion System Performance<br />
(Maximum Vacuum Thrust), (mt)<br />
First Stage*<br />
Second Stage<br />
Third Stage<br />
Upper (fourth) Stage<br />
31.0<br />
11.75<br />
4.15<br />
3.13<br />
Note: *Augmented First Stage Booster Engines have been used on the Proton Launch Vehicle since 1993.<br />
1070<br />
237<br />
62<br />
8.5<br />
30.6<br />
11.4<br />
3.7<br />
2.37<br />
1070<br />
237<br />
62<br />
2.0<br />
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Figure 1.6-1: Proton K and M Major Hardware Elements<br />
Payload<br />
fairing<br />
Payload<br />
adapter<br />
Third stage<br />
4.1 m ∅<br />
Second stage<br />
4.1 m ∅<br />
First stage<br />
20.2m long<br />
7.4 m dia. ∅<br />
4 4 .3 m<br />
lo n g<br />
(1) RD - 0210<br />
(4) - RD - 0210<br />
Core<br />
oxidizer<br />
tank<br />
Strap-on<br />
fuel tank<br />
1.7 m ∅<br />
(6)RD-253<br />
The third stage is equipped with one<br />
fixed single-chamber liquid propellant<br />
rocket engine developing 0.6 MN thrust<br />
and one liquid propellant control rocket<br />
engine, with four gimbaled nozzles,<br />
developing 30 kN thrust<br />
The second stage is equipped with four<br />
gimbaled single-chamber liquid<br />
propellant rocket engines developing a<br />
total thrust of 2.3 MN<br />
The first stage is equipped with six gimbaled<br />
single-chamber liquid propellant rocket<br />
engines developing a total thrust of 9 MN at<br />
lift-off<br />
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1.6.1 Proton First Stage<br />
The Proton first stage consists of a central tank containing the oxidizer, surrounded by six outboard fuel tanks.<br />
Although these fuel tanks give the appearance of being strap-on boosters, they do not separate from the core tank<br />
during first stage flight. Each fuel tank also carries one of the six RD-253 engines that provide first stage power. Total<br />
first stage sea-level thrust is approximately 9.5 MN (2.1 x 10 6 lbf). Total first stage dry mass is approximately 31,000<br />
kg (68,343 lbm); total first stage propellant load is approximately 419,410 kg (924,641 lbm).<br />
The Proton M booster first stage improves on the current booster with a small reduction in overall stage mass. Mass<br />
optimization results from modern manufacturing techniques and reduction in avionics system mass.<br />
In addition to the structual enhancement, the RD-253 first stage engines are uprated. For performance purposes, rated<br />
thrust on the engines of the first stage of Proton M is being increased by 7%. This enhancement is accomplished with<br />
only a minor modification to the propellant flow control valves. This modification first flew as a <strong>mission</strong>-unique<br />
enhancement on the Proton K that delivered the MIR space station core module in 1986. Engines incorporating this<br />
change have undergone extensive additional qualification firings since then, in order to approve them for use in<br />
standard production <strong>vehicle</strong>s. Shipsets have already been produced and incorporated onto production Proton K<br />
<strong>vehicle</strong>s. To date, 65 Proton K's have been <strong>launch</strong>ed with the modified engines operating at 102% of rated thrust.<br />
Eight Proton K <strong>vehicle</strong>s have flown with the 107% engines. There have been no flight anomalies attributed to the<br />
increased thrust engines.<br />
The propellant feed systems of the first and second stages have been simplified and redesigned in order to reduce<br />
propellant residuals in these stages by 50%, and a propellant purge system has been added to dump all residuals from<br />
the spent first and second stages before they return to the earth's surface.<br />
While a reduction in unusable propellants results in a performance gain, the primary rationale for the increased<br />
utilization of propellants is to reduce the environmental effects of the impact of the first and second stages in the<br />
downrange "land" jettison zones.<br />
1.6.2 Proton Second Stage<br />
The second stage, of conventional cylindrical design, is powered by four RD-0210 engines; it develops a vacuum thrust<br />
of 2.3 MN, or 5.24 x 10 5 1bf. Total second-stage dry mass is approximately 11,715 kg (25,827 lbm); total second stage<br />
propellant load is approximately 156,113 kg (344,170 lbm).<br />
Modifications to Proton second stage include structural reinforcement of the forward portion of the stage, in order to<br />
carry greater payload and aerodynamic loads, and minor structural weight reductions.<br />
1.6.3 Proton Third Stage<br />
The third stage is powered by one RD-0210, developing 583 kN (1.31 x 10 5 Lbf) thrust, and a four-nozzle vernier<br />
engine that produces 31 kN (6.9 x 10 3 lbf) thrust. Total third-stage dry mass is approximately 4,185 kg (9,226 lbm);<br />
total third stage propellant load is approximately 46,562 kg (102,652 lbm).<br />
Modifications to the Proton third stage, include structural reinforcement of the aft portion of the stage in order to<br />
carry greater payload and aerodynamic loads, and minor structural weight reductions.<br />
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1.6.4 Proton Flight Control System<br />
Guidance, navigation, and control of the Proton K during operation of the first three stages is carried out by a triple<br />
redundant closed-loop analog avionics system mounted principally in the Proton's third stage, using data collected<br />
from a distributed set of inertial reference platforms. This system also provides for flight termination in the event of a<br />
major malfunction during ascent.<br />
For Proton M, the principal modification to the Proton K's first three stages is the incorporation of a digital flight<br />
control system based on modern avionics technology. This digital system replaces the analog flight control hardware of<br />
the Proton K, which, although it has achieved a high demonstrated reliability, is based on obsolete 1960's era<br />
electronic designs. The new system eliminates the need to maintain a unique and limited production capability for<br />
Proton K avionics, and allows for simplified control algorithm loading and test. It also enables greater ascent program<br />
design flexibility with respect to <strong>vehicle</strong> pitch profile and other parameters. The system is being developed by the Mars<br />
and Proton Control Systems Companies.<br />
The self-contained inertial control system uses a precision three-axis gyro-stabilizer and a three-channel voting onboard<br />
digital computer. The digital control computer resides on the Proton M third stage and controls the flight of the<br />
first, second, and third stages.<br />
1.7 BLOCK DM FOURTH STAGE<br />
The Block DM fourth stage, normally uses liquid oxygen and kerosene, although the stage can be operated with several<br />
different kerosene-type fuels. The configuration of the Proton K/Block DM is depicted in Figure 1.7-1. The Block<br />
DM is shown in cutaway, within its aerodynamic shrouds ("middle and lower adapters") in Figure 1.7-2, and as it<br />
appears in flight in Figure 1.7-3. The Block DM and its predecessor, the Block D, have together flown more than 200<br />
times.<br />
The Block DM is optimized for multi-burn space transfer operations. Its main engine (model number 1lD58M)<br />
delivers a vacuum thrust of 83.5 kN (1.88 x 10 4 lbf), is gimbaled to provide three axis control during powered flight<br />
operations, and can be restarted as many as seven times during flight. The stage is 3.7 m (12.1 ft) in diameter, 6.28 m<br />
(20.6 ft) in length, with an inert mass at separation of 2440 kg (5378 lbm) and a total propellant mass of 15,050 kg<br />
(33,180 lbm). It is three-axis stabilized in unpowered flight by a storable bipropellant (N2O4/UDMH) attitude<br />
control system, comprised of two "SOZ" (or "micro") thruster units located at the base of the Block DM. The fourth<br />
stage can achieve a maximum of 1.5 rpm rotation at the moment of spacecraft separation. It should be noted, however,<br />
that the Block DM can only rotate approximately +/-180 degrees from a "reference" orientation, due to limitations of<br />
the stage's gyro platform; it cannot undergo continuous rotation. Guidance, navigation, and control of the fourth stage<br />
are provided by a triple redundant digital avionics package, which can be ground commanded in flight, if necessary.<br />
The Block DM equipment bay provides the payload adapter and electrical interfaces to the customer’s spacecraft. The<br />
interface between the stage and its payload adapter is 2000 mm in diameter, allowing the Block DM to accommodate<br />
large diameter payload adapters and a static bending moment about this interface of 11,000 kg-m. The Block DM<br />
payload fairing is a two piece composite structure developed specifically to meet the volume and environmental<br />
requirements of commercial spacecraft. Payload fairing usable volume geometry is provided in Appendix D of this<br />
Mission Planner’s Guide.<br />
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Figure 1.7-1: Proton K/Block DM Major Hardware Elements<br />
Adapter<br />
Fourth Stage<br />
Block - DM<br />
Forward fourth<br />
stage shroud<br />
Aft fourth<br />
stage shroud<br />
Third stage<br />
4.1 m ∅<br />
Second stage<br />
4.1 m ∅<br />
First stage<br />
20.2m long<br />
7.4 dia m ∅<br />
Payload<br />
Fairing<br />
Payload<br />
(1) 11D58M<br />
4 4 .3 m<br />
lon g<br />
(1) RD - 0210<br />
(4) RD - 0210<br />
Core<br />
oxidizer<br />
tank<br />
Strap-on<br />
fuel tank<br />
1.7 m∅<br />
(6) RD-253<br />
The fourth stage is equipped with<br />
one liquid propellant rocket engine<br />
developing 86 kN thrust, and two<br />
"micro" engine clusters for attitude<br />
control and ullage maneuvers<br />
The third stage is equipped with one<br />
fixed single-chamber liquid propellant<br />
rocket engine developing 0.6 MN<br />
thrust and one liquid propellant control<br />
rocket engine, with four gimbaled<br />
nozzles, developing 30 kN thrust<br />
The second stage is equipped with four<br />
gimbaled single-chamber liquid<br />
propellant rocket engines developing a<br />
total thrust of 2.3 MN<br />
The first stage is equipped with six gimbaled<br />
single-chamber liquid propellant rocket<br />
engines developing a total thrust of 9 MN<br />
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Figure 1.7-2: Proton DM Upper Stage Within Its Aerodynamic Shrouds<br />
1. Equipment bay<br />
2. Oxidizer tank<br />
3. Fuel tank<br />
4. Auxilary propulsion for attitude control<br />
("SOZ " units)<br />
5. Main engine<br />
6. Middle and lower adapters (shrouds)<br />
6280 mm<br />
4900 mm<br />
1<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
φ2000 mm<br />
φ3700 mm<br />
φ4100 mm<br />
Plane of spacecraft adapter/<br />
Block DM joint<br />
1<br />
6<br />
2<br />
3<br />
4<br />
5<br />
6<br />
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Figure 1.7-3: Block DM As It Appears In Flight<br />
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1.8 BREEZE M FOURTH STAGE<br />
The Breeze M upper stage, which is derived from the Breeze K stage flown on the Rokot, offers substantially improved<br />
payload performance and operational capabilities over the Block DM flown on the current Proton K. The Breeze M<br />
program was initiated in 1994 by the Khrunichev Space Center and has the full support of the Russian government. An<br />
isometric view of the Proton M/Breeze M is shown in Figure 1.8-1.<br />
The Breeze M upper stage is 2.61 meters in height and 4.0 meters in diameter, with an inert mass of 2,370 kg and a<br />
total propellant mass of 19,800 kg. It consists of two main elements:<br />
1) a core section (derived from the original Breeze K stage) that accommodates a set of propellant tanks, the<br />
propulsion system, and the avionics equipment bay, and<br />
2) a toroidal auxiliary propellant tank that surrounds the core section, and which is jettisoned in flight following<br />
depletion. Use of an external drop tank substantially improves the performance of the Breeze M stage.<br />
Figures 1.8.-2 and 1.8-3 illustrates the layout and dimensions of the Breeze M.<br />
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Figure 1.8-1: Proton M/Breeze M Major Hardware Elements<br />
Adapter<br />
Fourth Stage<br />
Block - DM<br />
Forward fourth<br />
stage shroud<br />
Aft fourth<br />
stage shroud<br />
Third stage<br />
4.1 m ∅<br />
Second stage<br />
4.1 m ∅<br />
First stage<br />
20.2m long<br />
7.4 dia m ∅<br />
Payload<br />
Fairing<br />
Payload<br />
(1) 11D58M<br />
4 4 .3 m<br />
lon g<br />
(1) RD - 0210<br />
(4) RD - 0210<br />
Core<br />
oxidizer<br />
tank<br />
Strap-on<br />
fuel tank<br />
1.7 m∅<br />
(6) RD-253<br />
The fourth stage is equipped with<br />
one liquid propellant rocket engine<br />
developing 86 kN thrust, and two<br />
"micro" engine clusters for attitude<br />
control and ullage maneuvers<br />
The third stage is equipped with one<br />
fixed single-chamber liquid propellant<br />
rocket engine developing 0.6 MN<br />
thrust and one liquid propellant control<br />
rocket engine, with four gimbaled<br />
nozzles, developing 30 kN thrust<br />
The second stage is equipped with four<br />
gimbaled single-chamber liquid<br />
propellant rocket engines developing a<br />
total thrust of 2.3 MN<br />
The first stage is equipped with six gimbaled<br />
single-chamber liquid propellant rocket<br />
engines developing a total thrust of 9 MN<br />
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Figure 1.8-2: Breeze M (With Auxiliary Propellant Tank)<br />
1. Equipment Bay<br />
2. Core stage<br />
3. Auxiliary propellant tank<br />
4. Oxidizer tank<br />
5. Fuel tank<br />
6. Auxiliary propulsion for attitude control<br />
7. Main engine<br />
2610 mm<br />
Figure 1.8-3: The Breeze M And Breeze M Core As They Appear In Flight<br />
? 2490 mm<br />
? 4000 mm<br />
The Breeze M uses nitrogen tetroxide (N 2O 4) and unsymmetrical-dimethylhydrazine (UDMH) as propellants.<br />
Propulsion for the Breeze M consists of one pump-fed, gimbaled main engine developing 19.62 kN of thrust, four<br />
"impulse adjustment thrusters" of 396 N thrust each for making fine corrections to the main engine impulse, and 12<br />
attitude control thrusters of 13.3 N thrust each. The main engine can relight up to eight times per <strong>mission</strong>, and is<br />
equipped with a backup restart system that can fire the engine in the event of a primary ignition sequence failure.<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
7<br />
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During two flights of the Phobos space probes in 1988 and three flights of the Breeze K on the Rokot in 1990, 1991,<br />
and 1994 the main engine demonstrated up to five restarts in flight. Following minor modifications to adapt the engine<br />
for the Breeze M, eleven main engines have been ground tested, some up to 6,000 seconds total burn duration. The<br />
Breeze M attitude control thrusters were previously used on the Kvant, Kristall, Spektr, and Priroda modules of the<br />
MIR space station, and are used on the Russian FGB and Service Module components of the International Space<br />
Station.<br />
The control system of the Breeze M includes an on-board digital computer and three axis inertial measurement unit.<br />
It also incorporates GLONASS and NAVSTAR GPS navigation systems. Breeze M can perform preprogrammed<br />
maneuvers about all axes during parking orbit and transfer orbit coast. The stage is normally three axis stabilized<br />
during coast, but can be rotated at up to 2 degrees per second for thermal control. During powered flight the upper<br />
stage attitude is determined by <strong>mission</strong> specific pitch, yaw and roll programs. Breeze M can perform separation of a<br />
customer's spacecraft in either one of two modes:<br />
1) attitude hold mode, during which the angular rates in relation to any of the coordinate system axes will not exceed<br />
0.5 degrees per second, and the spatial attitude error in relation to the inertial coordinate system will not exceed 1<br />
degree, or<br />
2) spin-up mode, during which the stage can achieve a maximum angular rate with respect to the upper stage<br />
longitudinal axis of 30 degrees per second, and the spin axis deflection from the upper stage longitudinal axis will<br />
not exceed 0.05 degrees.<br />
The typical Proton M/Breeze M <strong>mission</strong> profile is discussed further in Section 3 of this Mission Planner’s Guide.<br />
The Breeze core structure provides the payload adapter and electrical interfaces to the customer's spacecraft. The<br />
interface between the stage and its payload adapter is 2490-mm in diameter, allowing the Breeze M to accommodate<br />
large diameter payload adapters and a static bending moment about this interface of 18,000 kg-m. The Breeze M stage<br />
is encapsulated within the payload fairing, along with the customer's spacecraft, allowing loads from the payload<br />
fairing to be borne by a short (600-mm) spacer ring attached to the Proton third stage equipment section, rather than<br />
by the Breeze M. The Breeze M payload fairing is a derivative of the payload fairing currently in use with commercial<br />
spacecraft on the Proton K/Block DM. Modifications consist of a redefined attachment geometry at the aft end of the<br />
fairing; the attachment and separation hardware, however, is essentially unchanged from the current design. Payload<br />
fairing usable volume geometry is provided in Appendix D of this Mission Planner’s Guide.<br />
1.9 PAYLOAD FAIRINGS<br />
Figure 1.9-1 illustrates the available payload fairings for the three stage Proton K and Proton M. These fairings<br />
typically enclose both the payload of the Proton and any supplemental orbital propulsion units employed by the<br />
payload. The payload fairing is jettisoned after second stage separation, with the exact time determined by spacecraft<br />
free-molecular heating and impact zone constraints. Normally payload fairing separation occurs at approximately 344<br />
seconds into flight for commercial <strong>mission</strong>s.<br />
Figure 1.9-2 illustrates the standard commercial payload fairing for use with the Block DM upper stage, and either the<br />
Proton K or the Proton M.<br />
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Figure 1.9-3 shows the standard commercial payload fairing available for use with the Breeze M upper stage. The<br />
usable volume of this fairing is comparable to that of the payload fairing for the Block DM. Figure 1.9-4 gives<br />
dimensions for the long version of the Breeze M payload fairing. Use of this fairing results in a decrease in<br />
performance to GTO of approximately 100kg.<br />
Figure 1.9-1: Proton K and M Payload Fairings<br />
250 mm<br />
150 mm<br />
250 mm<br />
250 mm<br />
100 mm<br />
30°<br />
9047.5 mm<br />
L = 2245 mm<br />
11970 mm<br />
12650 mm<br />
300 mm<br />
300 mm<br />
150 mm<br />
1600 mm<br />
2840 mm<br />
100 mm<br />
φ4100 mm φ4100 mm<br />
2150 mm<br />
15882 mm<br />
1000 mm<br />
3000 mm<br />
3752 mm<br />
14562 mm<br />
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Figure 1.9-2: Block DM Payload Fairing for Single Spacecraft<br />
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Figure 1.9-3: Breeze-M Payload Fairing (Standard) For Single Spacecraft<br />
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Figure 1.9-4: Breeze M Payload Fairing (Long)<br />
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1.10 BAIKONUR INFRASTRUCTURE AND SERVICES<br />
All Proton <strong>launch</strong>es occur from the Baikonur Cosmodrome, which is located in the Republic of Kazakhstan in central<br />
Asia. Baikonur is approximately 2,000 km (1,300 miles) southeast of Moscow. Located in a semiarid zone, annual<br />
temperature and climate vary considerably. The Proton <strong>launch</strong> system is designed to reliably operate from just such a<br />
climatic range.<br />
1.10.1 Baikonur Infrastructure<br />
All primary <strong>launch</strong> systems infrastructure at Baikonur are being maintained in accordance with standard Khrunichev<br />
Space Center (KhSC) and Russian Space Agency (formerly Russian Space Forces) policies concerning operational<br />
efficiency and system safety. A number of facilities in use to support the Proton and satellite <strong>mission</strong> <strong>launch</strong> campaign<br />
process have been upgraded to support “Western-style” <strong>launch</strong> campaign processes. Of primary importance are<br />
upgrades to Launch Complex 24 at Baikonur and the addition of new, integrated payload processing facilities in close<br />
proximity to the <strong>launch</strong> complex. Figure 1.10.1-1 shows the locations of the major facilities at the Baikonur<br />
Cosmodrome.<br />
Figure 1.10.1-1: Overall Layout Of The Baikonur Cosmodrome<br />
SC Processing Area<br />
(Facility 92A-50)<br />
Area 92<br />
Proton LV<br />
Processing Area<br />
6<br />
7<br />
Area 95<br />
Living Quarters<br />
Kranyi Airfield<br />
Proton LV<br />
Launch Complex 81<br />
(Pads 23 and 24)<br />
9<br />
4<br />
5<br />
90 km<br />
DM Upper Stage<br />
Processing Area<br />
2<br />
1<br />
Yubileini Airfield<br />
Sir-Darya River<br />
Baikonur Town<br />
(Leninsk)<br />
Area 31<br />
3<br />
DM Upper Stage<br />
Fueling Area<br />
Legend:<br />
Railroad<br />
Road<br />
1-2 24 km<br />
2-3 18 km<br />
2-4 18 km<br />
4-5 16 km<br />
4-6 23.5 km<br />
6-7 1.25 km<br />
7-8 9 km<br />
8-9 7 km<br />
N<br />
75 km<br />
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1.10.1.1 Yubileini Airport<br />
All <strong>launch</strong> campaign personnel, ground support equipment, and the satellite arrive at Baikonur by air via the Yubileini<br />
airport. Yubileini operates as a Category 3 airport facility capable of supporting all large transport aircraft. Airport<br />
instrumentation and support services have been brought inline with western requirements with regard to aircraft<br />
operations and aircraft onload/offload. The facility is capable of supporting operations in most weather conditions.<br />
Upon offload of the spacecraft and support equipment, ground handling equipment is available to place palettes onto a<br />
properly configured series of rail cars for transportation to the payload processing facility. The aircraft offload/rail car<br />
onload process can take up to 8 hours. ILS offers an environmentally controlled rail car for thermal conditioning of the<br />
satellite transportation container.<br />
1.10.1.2 Area 92 Launch Vehicle and Satellite Processing<br />
Area 92 contains the major processing facilities for both the spacecraft and the Proton <strong>launch</strong> system. Building 92-1 is<br />
the facility in which Proton <strong>vehicle</strong>s are integrated and tested before mating with the fourth-stage/payload “ascent<br />
unit.” Building 92A-50 houses the new SVPF (Figure 1.10.1.2-1). The Space Vehicle Processing Facility is covered in<br />
greater detail in Section 6 of this Mission Planner’s Guide. Area 92A-50 payload processing facilities are<br />
approximately 20 km from airport facilities.<br />
Figure 1.10.1.2-1: Building 92A-50 Space Vehicle Processing Facility<br />
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1.10.1.3 Area 31 Payload Processing<br />
Area 31 was configured to enable nonhazardous and hazardous processing of Western communications satellites to<br />
support the first commercial Proton <strong>launch</strong>es. All ILS <strong>launch</strong>es through the end of 1997 were processed in Area 31<br />
facilities, specifically Building 40D for nonhazardous processing and Building 44 for hazardous processing.<br />
Following transition to the SVPF, Area 31 has been retained in a backup mode to support contingency and/or<br />
processing “surge” commercial requirements. ICD’s for commercial spacecraft should continue to be written to<br />
support this facility. Area 31 will continue to support Russian government payloads such as the SOYUZ manned<br />
system.<br />
1.10.1.4 Area 81 Launch Complex<br />
Area 81 contains both active commercial Proton <strong>launch</strong> pads. Launch Pad 23 is presently <strong>launch</strong>ing commercial<br />
Proton K/Block DM <strong>vehicle</strong>s in support of ILS. Pad 24 has been refurbished to support <strong>launch</strong> operations for the<br />
Proton K/Block DM system and the Proton M/Breeze M <strong>vehicle</strong>. Launch operations on Complex 24 will commence<br />
in mid-1999 after refurbishment and upgrades are complete. Complex 23 will support the ILS Proton business as a<br />
backup pad, as business requirements dictate.<br />
1.10.1.5 Area 95 Hotel Facilities<br />
ILS and our KhSC partners have spent significant effort and resources upgrading personnel and logistics-related<br />
facilities. Area 95 houses the three primary hotel facilities used by ILS and western <strong>launch</strong> campaign teams. Hotels<br />
Fili, Cometa and Polyot are the primary accommodations for satellite contractor personnel.<br />
In addition to the physical accommodations, enhancements have been made to the cafeteria services and utilities<br />
infrastructure to better accommodate western <strong>launch</strong> crews. A water treatment plant is in operations 24-hours per day,<br />
offering water purified to U.S. standards. Additional support services, such as recreation equipment, have also been<br />
implemented.<br />
ILS typically arranges for the accommodation of up to 30 customer and satellite contractor personnel per spacecraft<br />
during the Baikonur <strong>launch</strong> campaign process, with each spacecraft campaign expected to last approximately 30 days.<br />
1.10.2 Proton Launch Campaign<br />
For typical commercial Proton <strong>launch</strong>es, ILS has developed a standard 30-day <strong>launch</strong> campaign schedule for purposes<br />
of initial discussions with potential Proton <strong>launch</strong> system users. As shown in Figure 1.10.2-1, a 30-day campaign is<br />
based on the general philosophy that the satellite is shipped from the satellite manufacturer’s facility in a near “shipand-shoot”condition.<br />
For most <strong>mission</strong>s in which ILS is under contract, the satellite is shipped to Baikonur in a<br />
configuration that requires minimal nonhazardous checkout. ILS <strong>launch</strong> campaign process goals reflect the desire to<br />
minimize <strong>launch</strong> campaign duration, without jeopardizing <strong>mission</strong> success. The 30-day cycle is consistent with this<br />
philosophy.<br />
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Figure 1.10.2-1: Proton Launch Campaign for 92A-50<br />
L-30 days<br />
ILS & SC Technical Support<br />
Team Onsite<br />
Spacecraft L-30 days L-28 days L-27 days L-26 days<br />
SC Loaded<br />
into SC Container<br />
at Factory<br />
Ground Support<br />
Equipment L-34 days<br />
GSE Loaded in Cargo<br />
Containers & on<br />
Pallets at Factory<br />
L-3 months L-2.5 months L-2 months L-18 days<br />
Propellant Shipment to<br />
Russia Port-of-Entry<br />
L-11 » L-10 days L-16 » L-14 days<br />
• Transfer Payload to<br />
Launch Vehicle 4th<br />
Stage Refueling Area<br />
• Mate Spacecraft<br />
Vertically To Launch<br />
Vehicle 4th Stage<br />
• Tilt to Horizontal &<br />
Mate Fairing<br />
• Test Transit Cable<br />
• Load Spacehead Unit<br />
on Transporter<br />
Move SC Container to Airport<br />
via 5-Wheel Tow; Load SC<br />
Container into Transport Aircraft<br />
Move GSE to Embarkation<br />
Facility for Shipment to Russia<br />
L-9 » L-7 days<br />
Rail Shipment to<br />
Baikonur<br />
• Final Operations on<br />
Spacecraft<br />
• Fuel Launch Vehicle<br />
Fourth Stage<br />
Transport Payload<br />
Spacehead<br />
To Launch Vehicle<br />
Assembly Area<br />
L-13 » L-12 days<br />
Checkout<br />
Spacecraft<br />
Pyrotechnics<br />
SC & GSE<br />
Arrives at Airport<br />
Near Baikonur<br />
Launch Site<br />
Temporary Storage<br />
Move Fuel Charts<br />
to Fuel Hall<br />
L-30 » L-19 days<br />
L-17 days L-18 days<br />
Fuel SC at Fueling Hall with<br />
SC Fuel Carts & Return to<br />
Technical Assembly Hall<br />
L-6 days L-5 » L-4 days L-3 days<br />
• Prepare & Mate Payload<br />
Section with Launch<br />
Vehicle<br />
• Checkout Electrical<br />
Connections<br />
• Transfer Launch Vehicle<br />
with Payload Section<br />
onto Erector Transporter<br />
Transport SC<br />
to Technical<br />
Assembly Hall<br />
• Proton Rollout from<br />
Horizontal Building<br />
• Install on Launch Pad<br />
• Do General SC<br />
Checkout on<br />
Launch Pad<br />
• General Testing &<br />
Preparation of SC<br />
& Launch Vehicle<br />
• Air Conditioning &<br />
Batteries Charge<br />
Hook-Up<br />
• Electrical Tests of SC<br />
Install SC<br />
in Cleanroom<br />
Connect Electrical<br />
GSE to SC<br />
• Test TC&C Systems<br />
• Test Solar Arrays<br />
• Pneumatic &<br />
Pressure Tests<br />
• Test Deployment<br />
Mechanisms<br />
L-2 days<br />
Prepare SC<br />
for Fueling &<br />
Transfer to<br />
Fueling Hall<br />
Integrated SC<br />
& Launch<br />
Vehicle Test<br />
Flight<br />
Readiness<br />
Review<br />
L-1 » L-0 days<br />
Countdown<br />
& Launch<br />
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1.11 PLANNED ENHANCEMENTS<br />
Beyond the Proton M/Breeze M, our Khrunichev partners have planned a number of possible upgrades to the Proton<br />
<strong>launch</strong> system. Payload fairings that are longer in length and/or wider in diameter have been assessed. A tandem<br />
<strong>launch</strong> system concept has been studied.<br />
Additionally, a new liquid oxygen/liquid hydrogen high energy upper stage, to be used in place of the current<br />
LOX/hydrocarbon Block DM or storable propellant Breeze M for some <strong>mission</strong>s, is also under development.<br />
Maximum performance capability of the improved Proton is expected to be as follows:<br />
LEO (200 km circ.) 22.0 metric tons<br />
GSO (LOX/LH2) 4.5 metric tons<br />
The above enhancements are currently projected to begin appearing in flight <strong>vehicle</strong>s starting in the year 2001 and on.<br />
Additional detail can be found in Section 8 of this Guide.<br />
1.12 PROTON PRODUCTION AND OPERATIONS-SLIDES<br />
Figures S-1 to S-10 show manufacturing operations for first three stages of Proton at the Khrunichev plant outside of<br />
Moscow. Figures S-11 to S-13 show the Block DM assembly area at Energia. Figures S-14 and S-15 show external<br />
and internal views of the Proton Standard Commercial Payload Fairing. Figures S-16 to S-19 show the Breeze M<br />
being manufactured at Khrunichev. Figure S-20 shows the arrival of a commercial spacecraft, in its shipping<br />
container, at the Yubileni airport at Baikonur. Figures S-21 and S-22 show spacecraft processing operations in Area 92<br />
and Area 31, respectively. Figure S-23 shows the complete Proton <strong>launch</strong> <strong>vehicle</strong>, after integration of the Ascent Unit<br />
onto the first three stages, being lifted onto its transporter/erector. Figure S-24 illustrates the <strong>launch</strong> <strong>vehicle</strong> interfaces<br />
at the base of a Proton <strong>launch</strong> pad. Figure S-25 shows the Proton being erected onto the pad. Figure S-26 shows the<br />
Proton’s Mobile Service Tower. Figure S-27 shows the Proton at the moment of lift-off.<br />
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PROTON<br />
PRODUCTION AND OPERATIONS<br />
SLIDES<br />
S-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
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Figure S-1: Tank Component Fabrication At Khrunichev Uses Automated Machining Centers<br />
Figure S-2: First Stage Propellant Tank Automated Weld-Up<br />
S-2
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Figure S-3: First, Second, And Third Stage Sub-Assemblies Awaiting Integration<br />
Figure S-4: Proton Interstage Component Fabrication<br />
S-3
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Figure S-5: First Stage Build Up On Proton Fixture<br />
Figure S-6: Interstage Joining Second and Third Stages<br />
S-4
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Figure S-7: RD-253 High-Pressure Engine On The First Stage External Fuel Tank<br />
Figure S-8: Proton Final Assembly Hall At Khrunichev<br />
S-5
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Figure S-9: Assembled First Stage Showing Hold Down Points and Aft End Services Connectors<br />
Figure S-10: End-To End Test Of Assembled Proton At Khrunichev Factory<br />
S-6
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Issue 1, Revision 4, March 1, 1999<br />
Figure S-11: The Block DM Undergoing Final Assembly And Testing At The Energia Plant In Korolev, Near Moscow<br />
Figure S-12: The Completed Block DM Stage, Before Attachment Of The External Shroud<br />
S-7
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Issue 1, Revision 4, March 1, 1999<br />
Figure S-13: Finished Block DM Stages In Their Aerodynamic Shrouds, Awaiting Shipment To Baikonur<br />
Figure S-14: Proton Standard Commercial Payload Fairing External View<br />
S-8
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Figure S-15: Proton Standard Commercial Payload Fairing Internal View<br />
Figure S-16: Breeze M Stage Structural Components In Manufacture<br />
S-9
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Figure S-17: Breeze M Core Stage Manufacturing<br />
Figure S-18: Breeze M Avionics Bay Assembly<br />
S-10
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Figure S-19: Breeze M Stage Final Assembly<br />
Figure S-20: Spacecraft Arrival at Baikonur<br />
S-11
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Figure S-21: Spacecraft Processing In Area 92 (SVPF)<br />
S-12
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Issue 1, Revision 4, March 1, 1999<br />
Figure S-22: Spacecraft Processing and Encapsulation In Area 31<br />
Figure S-23: The Complete Proton Launch Vehicle Being Lifted Onto Its Transporter/Erector<br />
S-13
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Figure S-24: Launch Vehicle Interfaces at Proton Launch Pad, Baikonur<br />
Figure S-25: Proton Erection At The Launch Pad<br />
S-14
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Figure S-26: Mobile Service Tower (MST)<br />
S-15
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Figure S-27: Lift-Off Of The Proton K/Block DM Carrying A Commercial Communications Satellite<br />
S-16
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2. VEHICLE PERFORMANCE<br />
2.1 OVERVIEW<br />
This section provides the information needed to make preliminary performance estimates for several variants of the<br />
Proton family, into a variety of <strong>mission</strong> orbits. It is organized so as to provide the user with essential background<br />
<strong>mission</strong> planning information; detailed performance tables and charts follow the text material.<br />
Trajectory profile and operational <strong>mission</strong> characteristics of the Proton K and Proton M <strong>launch</strong> systems using the<br />
Block DM and Breeze M upper stages are provided in the first sections of the chapter. Mission performance data,<br />
guidance accuracy data, and upper stage attitude control capabilities are found in the last half of this section.<br />
2.2 PROTON LAUNCH SYSTEM CAPABILITIES<br />
The Proton has been operational since 1970 (Pre-1970 <strong>launch</strong>es were considered developmental), and has carried out<br />
more than 246 <strong>launch</strong>es as of December 1998. Over the last 50 <strong>launch</strong>es, the Proton has achieved a success rate in<br />
excess of 92%. The Proton <strong>launch</strong> <strong>vehicle</strong> is presently available in three variants: the three-stage Proton booster the K<br />
and M variants offer similar performance in this configuration and the four-stage Proton K/Block DM and Proton<br />
M/Breeze M variants. Proton K/Breeze M and Proton M/Block DM variants may also be implemented, should this<br />
be desireable.<br />
The production Proton K 4-stage <strong>vehicle</strong> is equipped with a large upper stage known as the Block DM. The Block DM<br />
can place spacecraft into a variety of orbits including low, intermediate, and high Earth circular, geotransfer,<br />
geosynchronous, sun synchronous, and inter-planetary trajectories.<br />
The Proton M is presently completing development. Although identical in outward appearance to the existing Proton<br />
K, it incorporates improvements to the avionics and structures of the first three stages. It also incorporates improved<br />
engines that have been flying since 1996. The Proton M is capable of delivering 21.0 metric tons into a 200 km<br />
circular, 51.6° inclination orbit. The Proton M/Breeze M will become available for commercial use in 2000. The<br />
Breeze M storable propellant upper stage offers enhanced performance and operational capability and is described in<br />
Section 1. Table 2.2-1 summarizes performance for Proton K, Proton M, Proton K/Block DM, and Proton<br />
M/Breeze M to a range of <strong>mission</strong> orbits.<br />
Table 2.2-1: Summary Proton Performance (PSM) to Reference Orbits<br />
Mission Proton K (Standard) Proton K (Enhanced) Proton M<br />
LEO (51.6 degrees) 19,760 kg (43,560 lb) 19,760 kg (43,560 lb) 21,000 kg (46,300 lb)<br />
GSO 1,880 kg (4,145 lb) 1,880 kg (4,145 lb) 2,920 kg (6,435 lb)<br />
GTO (1800 m/s from GSO) 4,350 kg (9,590 lb) N/A 6,220 kg (13,710 lb)<br />
GTO (1500 m/s from GSO) 4,350 kg (9,590 lb) 4,910 kg (10,825 lb) 5,500 kg (12,125 lb)<br />
Lunar 4,530 kg (9,987 lb) 4,530 kg (9,987 lb) 5,600 kg (12,345 lb)<br />
Mars Transfer 2,940 kg (6,482 lb) 2,940 kg (6,482 lb) 4,800 kg (10,580 lb)<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
2.2.1 The Baikonur Launch Site<br />
The Proton <strong>launch</strong> complex, consisting of spacecraft and <strong>launch</strong> <strong>vehicle</strong> processing and integration facilities and four<br />
<strong>launch</strong> pads (two of which are available for commercial use), is located at the Baikonur Cosmodrome. Baikonur,<br />
shown in Figure 2.2.1-1, is located approximately 2,000 km (1,300 miles) southeast of Moscow in the Republic of<br />
Kazakhstan. The Baikonur Cosmodrome measures approximately 90 km east-to-west, and 75 km north-to-south, and<br />
supports many other <strong>launch</strong> <strong>vehicle</strong>s, including the Soyuz, Vostok, Molniya, Zenit and Energia. Temperatures range<br />
from -40°C to 45°C during the year.<br />
Figure 2.2.1-1: The Baikonur Launch Site, Showing Available Direct Injection Inclinations<br />
80 o<br />
60 o<br />
40 o<br />
20 o<br />
0o 20o 0 o<br />
20 o<br />
BLACK SEA<br />
MEDI TERRANEAN SEA<br />
2.2.2 Launch Availability<br />
MOSCOW<br />
RED SEA<br />
40 o<br />
CA SPIAN SEA<br />
60 o<br />
FLIGHT AZIMUTH: 22.5 o<br />
INCLINATION: 72.7 o<br />
BAIKONUR<br />
COSMODROME<br />
(45.6 o N)<br />
(TYURATAM)<br />
80 o<br />
100 o<br />
120 o<br />
FLIGHT AZIMUTH: 31.0 o<br />
INCLINATION: 64.8 o<br />
FLIGHT AZIMUTH: 61.3 o<br />
INCLINATION: 51.6 o<br />
The Proton <strong>launch</strong> system is designed to operate under the severe environmental conditions encountered at Baikonur<br />
(Table 2.2.2-1). The Proton can be <strong>launch</strong>ed year around, and the time between <strong>launch</strong>es from an individual pad can<br />
be as short as 25 days. Proton has demonstrated a <strong>launch</strong> rate of four per month from multiple <strong>launch</strong> pads, and a long<br />
term average <strong>launch</strong> rate of approximately twelve per year. The capability of the Proton system to <strong>launch</strong> in severe<br />
environmental conditions decreases <strong>launch</strong> delays and ensures that payloads reach orbit as scheduled to begin revenue<br />
generating activities. The short turnaround time between <strong>launch</strong>es can ensure that spacecraft constellations are<br />
deployed quickly, minimizing the time required to enter service.<br />
140 o<br />
160 o<br />
180 o<br />
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Issue 1, Revision 4, March 1, 1999<br />
Table 2.2.2-1: Allowable Launch Environment Constraints<br />
Temperature -40 O C to 45 O C<br />
Maximum Launch Ground Winds (Commercial PLF) 16.5 m/s<br />
Times Launches available year round<br />
Turn Around Times 25 days per pad<br />
Number of Pads 4 (2 Commercial)<br />
2.2.3 Payload Fairings and Adapters<br />
2.2.3.1 Payload Fairings<br />
Multiple payload fairings types have flown on the Proton; these fairings are discussed in detail in Sections 1 and 4 of<br />
this document. ILS currently offers one standard commercial payload fairing for the four stage Proton K/Block DM.<br />
Two additional fairings are available for the Proton M/Breeze M.<br />
Payload fairing jettison times are constrained to occur so that fairing hardware will impact in designated areas. For<br />
Proton K/Block DM and initial Proton M/Breeze M flights, fairing jettison occurs at approximately 342 to 344<br />
seconds (121-125 kilometers altitude) into the flight. Free molecular heating is below 1,135 W/m 2 for these cases.<br />
2.2.3.2 Payload Adapters<br />
Available Proton payload adapters are shown in Appendix D. These adapters provide mechanical and electrical<br />
interfaces compatible with established commercial <strong>launch</strong> industry standards and can accommodate a wide range of<br />
payload requirements. ILS can assist in the development of specialized adapters as required.<br />
2.2.4 Upper Stage Capabilities<br />
Two upper stages are currently available for commercial Proton <strong>mission</strong>s, the Block DM and the Breeze M.<br />
The Block DM has many unique capabilities that can accommodate unique <strong>mission</strong> requirements. The heavy lift<br />
capability of the Proton’s first three stages, coupled with the Block DM’s high performance, allows the placement of<br />
spacecraft directly into their final orbits, reducing the size and complexity of spacecraft propulsion systems. The Block<br />
DM’s multiple start capability, designed for up to 7 starts in flight (with 5 demonstrated to date), and a minimum 24hr<br />
orbital lifetime, can increase <strong>mission</strong> utility through unique <strong>mission</strong> design. Block DM orbital lifetime can be<br />
extended through battery upgrades and other hardware modifications.<br />
The Breeze M upper stage expands upon the capabilities enabled by the Block DM. The main engine of the<br />
Breeze M can be restarted up to 8 times in flight and allows the stage to offer high precision placement of spacecraft<br />
into orbit. With storable propellant, Breeze M orbital lifetime is limited only by available on board battery power and<br />
is currently in excess of 24 hours. The jettisonable, toroidal propellant tanks offer significant <strong>mission</strong> design flexibility<br />
and enable <strong>launch</strong> services to be offered for low and high energy delivery requirements.<br />
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Issue 1, Revision 4, March 1, 1999<br />
2.3 PROTON ASCENT PROFILE<br />
2.3.1 Proton Booster Ascent<br />
The first three stages of the Proton <strong>launch</strong> <strong>vehicle</strong> use a standard ascent trajectory to place the orbital unit (upper stage<br />
and/or payload) into a 200 km (108 nmi) circular parking orbit inclined at either 51.6°, 64.8°, or 72.7°. A standard<br />
ascent trajectory is required to meet jettisoned stage and payload fairing impact point constraints. The use of a standard<br />
ascent trajectory also simplifies lower ascent <strong>mission</strong> design and related analysis, thereby increasing system reliability.<br />
Once a payload is in the standard parking orbit at one of the three available inclinations, it can be transferred to its<br />
final orbit by either the Block DM or Breeze M.<br />
Figure 2.3.1-1 pictorially illustrates a typical Proton ascent into the standard parking orbit. Table 2.3.1-1 lists the time<br />
of occurrence for major ascent events for a typical <strong>launch</strong>. The six Stage 1 RD-253 engines are ignited at<br />
approximately T-1.6 sec. and are commanded to 40% of nominal thrust. Thrust is increased to 100% at T-0 sec.<br />
Liftoff confirmation is signaled at T+0.5 sec. The staged ignition sequence allows verification that all engines are<br />
functioning nominally before being committed to <strong>launch</strong>. The <strong>launch</strong> <strong>vehicle</strong> executes a roll maneuver beginning at<br />
T+10 sec. to align the flight azimuth to the desired direction. The <strong>vehicle</strong> incurs its maximum dynamic pressure of<br />
3,890 N/m 2 (800 psf) approximately 65 sec. into flight. Stage 2’s four RD-0210 engines begin their ignition sequence<br />
at 122 sec. and are commanded to full thrust when Stage 1 is jettisoned at 126 sec. Payload fairing jettison typically<br />
occurs at 344.2 sec into flight, depending on spacecraft heating constraints. Stage 3’s vernier engines are ignited at 332<br />
sec. followed by Stage 2 shutdown at 334 sec. Stage 2 separation occurs after six small, solid retro-fire motors are<br />
ignited at 335 sec. into flight. Stage 3’s single RD-0210 engine is ignited at 339 sec. and burns until shutdown at 567.1<br />
sec. The four vernier engines burn for an additional 10 sec. and are shutdown at 585 sec. After a 10 second coast, the<br />
Stage 3 retro-fire motors are ignited and Stage 3 is separated from the upper stage or spacecraft. Figure 2.3.1-2 shows<br />
the ascent vacuum instantaneous impact points and ground track. Figure 2.3.1-3 shows the ascent telemetry coverage<br />
provided by the CIS ground stations. Figure 2.3.1-4 shows the times and values for the <strong>vehicle</strong>’s inertial velocity,<br />
altitude, and dynamic pressure.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Table 2.3.1-1: Typical Booster Ascent Event Times<br />
Event Description Event Time (sec)<br />
Stage 1 ignition - 10% thrust -1.60<br />
Begin stage 1 thrust to 100% -0.00<br />
Lift-off 0.57<br />
Stage 1 thrust to 100% 1.00<br />
Stage 2 ignition 116.91<br />
Stage 1 / 2 separation 121.11<br />
Stage 3 vernier engine ignition 330.00<br />
Stage 2 engine shutdown 332.70<br />
Stage 2 / 3 separation 333.40<br />
Stage 3 main engine ignition 335.80<br />
PLF jettison 344.20<br />
Stage 3 main engine shutdown 567.11<br />
Stage 3 vernier engine shutdown 577.11<br />
Stage 3 upper stage separation 582.01<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 2.3.1-1: Typical Proton Booster Ascent<br />
Separation<br />
of 1st stage<br />
T=126.73<br />
V=1649.4<br />
H=43.5<br />
n x =3.6/0.4<br />
q=3350<br />
1750 x 6 = 10500 kN<br />
q max<br />
T=65<br />
V=125.1<br />
H=10.5<br />
n x=1.97<br />
q=35265<br />
583 x 4 = 2332 kN<br />
L = 310<br />
Separation<br />
of 2nd stage<br />
T=335.12<br />
V=4331.9<br />
H=148.11<br />
n x =2.8/0.04<br />
q=0<br />
Retro<br />
Separation of payload fairing<br />
T=344.2<br />
V=4340.5<br />
H=157.7<br />
n x=0.9<br />
q=0<br />
L = 1985<br />
583 + 31 = 614 kN<br />
Separation<br />
of 3rd stage<br />
T=589.0<br />
V=7477.4<br />
H=228.0<br />
n x =0.13/0<br />
q=0<br />
Ignition No. 1<br />
of 4th stage<br />
n x=0.44<br />
Burnout<br />
n x=0.9<br />
Ignition No. 2<br />
of 4th stage<br />
n x=0.9<br />
Burnout and<br />
separation<br />
n x=1.7/0<br />
T = time (seconds)<br />
V = velocity (meters per second)<br />
H = altitude (kilometers)<br />
L = horizontal distance (kilometers)<br />
q = dynamic pressure (Pa)<br />
n x = axial acceleration at end of prior<br />
Retro<br />
Figure 2.3.1-2: Typical Proton Booster Ascent Ground Track and Vacuum Impact Points<br />
L a ti t u d e , d e g r e e s<br />
90<br />
75<br />
60<br />
45<br />
30<br />
15<br />
0<br />
-15<br />
-30<br />
-45<br />
-60<br />
-75<br />
Orbit Injection<br />
Vacuum Impact Points<br />
Moscow<br />
Stage 1 Jettison PLF Jett Option 1<br />
stage/beginning of next stage<br />
83.5 kN 83.5 kN<br />
Stage 2 Jettison PLF Jett Option 2<br />
-90<br />
-180 -150 -120 -90 -75 -60 -45 -30 -15 0 15 30 45 60 75 90 105 135 165<br />
-165 -135 -105 120 150<br />
Longitude, degrees<br />
180<br />
Page 2-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 2.3.1-3: Typical Ground Tracker Acquisition Times for Proton Ascent to the Support Orbit<br />
Ussurick (USK)<br />
Kolpashevo (KLP)<br />
Dshusali (DZhS)<br />
0 100 200 300 400 500 600<br />
Time from Liftoff (seconds)<br />
Figure 2.3.1-4: Typical Proton Lower Ascent Altitude, Inertial Velocity, Acceleration, and Dynamic Pressure<br />
Velo cit y ( m/s ec)<br />
Accel erat ion ( g)<br />
Alti tud e (k m)<br />
Dynami c pres sure (N/m 2 )<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
2.3.2 Block DM Trajectory and Sequence<br />
The Block DM transfers payloads to a variety of <strong>mission</strong> orbits. This section describes the two options enabling the<br />
Block DM to transfer a payload to a geosynchronous transfer orbit.<br />
Figures 2.3.2-1 and 2.3.2-2 illustrate a typical Block DM ascent into a geosynchronous transfer orbit. The Block DM<br />
is delivered to the 200 km circular parking orbit with a 51.6-deg inclination. Once on orbit and separated from the<br />
Proton third stage, the Block DM executes a 15-min duration maneuver to properly align its longitudinal axis for the<br />
first burn. After the alignment maneuver, the Block DM enters into a stabilized flight mode. Twenty-five min after the<br />
longitudinal alignment maneuver, the Block DM executes a 180-deg turn about the roll axis to compensate for<br />
possible gyroscopic drift. This maneuver can also help with spacecraft thermal management. Forty min after the roll<br />
maneuver, the Block DM reaches the first ascending node, and the two SOZ unit’s four 2.5-kg axial loading engines<br />
begin a 300-sec burn to settle the stage’s propellants. After this settling burn, the main engine ignites, raising the<br />
transfer orbit apogee to geosynchronous altitude. The first main engine burn lasts approximately 450 sec. After engine<br />
cutoff, the Block DM executes another alignment maneuver to place the longitudinal axis in the correct orientation for<br />
the second burn. The Block DM then enters stabilized flight for the approximately 5 hr and 15 min transfer required<br />
to reach transfer orbit apogee. At approximately 2.5 hr into this transfer ellipse, another 180-deg rotation about the<br />
roll axis is executed. After reaching the transfer orbit apogee, the Block DM initiates another 300-sec propellant<br />
settling burn followed by a main engine burn of approximately 230 sec. duration to circularize the orbit and reduce the<br />
inclination to 0 deg. The Block DM then commands spacecraft separation 14.8 ±0.05 sec after the end of the final<br />
burn. Figure 2.3.2-3 depicts the upper ascent ground track and Figure 2.3.2-4 shows available tracker coverage<br />
provided by the CIS ground stations. Telemetry (including state vector) data can also be collected from the Block DM<br />
during transfer orbit flight by Russian geostationary spacecraft. Table 2.3.2-1 gives the times and values of the Block<br />
DM attitude maneuvers. These maneuvers may be modified to assist in spacecraft thermal management; the Block<br />
DM can perform a maximum of 11 such maneuvers between the 1 st and 2 nd burn during a normal <strong>mission</strong>. with a<br />
<strong>mission</strong> maximum of 14. The standard <strong>mission</strong> scenario for commercial spacecraft involves retention of the SOZ units<br />
to provide attitude control following main engine shutdown. This allows spacecraft separation to be delayed until after<br />
the Block DM completes reorientation to a customer specified separation attitude.<br />
To more optimally deliver western satellites to orbit, the Block DM is also capable of performing a suborbital burn in<br />
order to enable larger masses to be delivered to low earth orbit. The larger mass is made up of the heavier payload and<br />
more fuel on the Block DM stage for orbital maneuvers. The “enhanced” Block DM performance level is enabled<br />
with the suborbital burn.<br />
Normally, injection into final orbit occurs at 90 O however, longitude placement for geosynchronous orbits can be<br />
controlled with an increment of 12.25 deg by allowing the Block DM and payload to coast in the standard parking<br />
orbit for the required number of revolutions, and initiating the first transfer orbit burn on either the ascending or<br />
descending node. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parking<br />
orbit revolutions are within the lifetime constraints of the Block DM, giving complete 360-deg coverage.<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 2.3.2-1: Proton/Block DM Upper Ascent<br />
180° rotation<br />
about roll axis<br />
Begin attitude<br />
maneuver for<br />
block DM 1st burn<br />
Begin propellant<br />
settling burn<br />
(MES1 - 300sec)<br />
Block DM<br />
1st burn<br />
(MES1)<br />
Event description<br />
(block DM upper ascent)<br />
MECO1<br />
Block 4 upper adapter Jettison<br />
SOZ unit first setting burn (SOZ1)<br />
Block 4 DM first burn<br />
SOZ unit engine shutdown<br />
SOZ unit second burn (SOZ2)<br />
Block 4 DM second burn<br />
SOZ unit engine shutdown<br />
Block 4 DM separation<br />
Begin altitude<br />
maneuver for<br />
block DM<br />
2nd burn<br />
Event time<br />
(sec)<br />
637<br />
5437<br />
5732<br />
5737<br />
24,800<br />
25,105<br />
25,105<br />
25, 345<br />
180° rotation<br />
about roll<br />
axis<br />
Begin propellant<br />
settling burn<br />
(MES1 - 300sec)<br />
MECO2<br />
Block DM<br />
2nd burn<br />
(MES 2)<br />
Block DM/<br />
payload<br />
separation<br />
MECO2<br />
+14.8sec<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 2.3.2-2: Block DM Two-Impulse Transfer to Geosynchronous Transfer Orbit<br />
SC Separation<br />
Target Orbit<br />
H a = 36,000 km<br />
H p = variable km<br />
7<br />
DV 3<br />
Parking Orbit<br />
1 - Launch from Baikonur Cosmodrome and booster ascent<br />
6<br />
5<br />
Intermediate<br />
Transfer Orbit<br />
2 - (DV 1 ) For sub-orbital burn cases, first Block DM main engine fairing<br />
3 - Coast to ascending node of parking orbit<br />
4 - (DV 2 ) Block DM main engine firing to raise orbit from LEO to GTO<br />
5 - Coast to apogee of transfer orbit (5.25 hours)<br />
6 - (DV 3 ) Block DM main engine firing to optimally raise perigee and reduce inclination<br />
7 - Spacecraft separation from Block DM<br />
3<br />
DV 1<br />
2<br />
1<br />
4<br />
DV 2<br />
i ref = 51.6 o<br />
GSO<br />
i t= variable<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 2.3.2-3: Proton K/Block DM Upper Ascent Ground Track to GSO<br />
Booster<br />
Burn 3<br />
K O<br />
AY1<br />
Figure 2.3.2-4: Ground Tracker Acquisition Times for Proton Ascent to a GSO<br />
Shelkovo (SHLK)<br />
Petropavlovsk-Kamchatski<br />
(PPK)<br />
Ussurick<br />
Ulan-Uden (ULD)<br />
Kolpaashevo (KLP)<br />
Dshusali (DZhS)<br />
600 5,60<br />
0<br />
10,600 15,60<br />
Time from liftoff<br />
(seconds)<br />
0<br />
Burn 2<br />
20,600 25,600<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 2.3.2-1: Typical Block DM Attitude Maneuvers for Geosynchronous Mission (90 O East Longitude)<br />
Start Time of Attitude<br />
Maneuver (hr.min.sec)<br />
Change in<br />
Pitch Angle<br />
(deg)<br />
Change in<br />
Yaw Angle<br />
(deg)<br />
Change in<br />
Roll Angle<br />
(deg)<br />
Duration of<br />
Maneuver<br />
(sec)<br />
00.10.44 43.00 - - 86<br />
Comments<br />
00.10.59.9 33.02 7.54 -13.25 269 Attitude alignment for First<br />
Block DM burn<br />
00.17.51 - - 180.00 900 IMU Compensation<br />
00.35.45 - - -180.00 -900 IMU Compensation<br />
01.01.19 12.07 - - 60 First Block DM burn<br />
-01.13.28 -20.90 - - 420 First Block DM burn<br />
01.21.36 9.00 - - 48 First Block DM burn<br />
02.26.45 - 180.00 - 900 IMU Compensation<br />
03.38.37 180.00 - - 900 IMU Compensation<br />
04.50.10 - 180.00 - 900 IMU Compensation<br />
06.18.33.5 7.78 68.37 2.6 273 Second Block DM burn<br />
2.3.3 Breeze M Trajectory Sequence<br />
The Breeze/payload unit is placed into a high-energy suborbital state by the third stage of Proton. After jettison of the<br />
third stage, the Breeze upper stage performs a small propulsive maneuver to deliver itself and the attached satellite to a<br />
standard low earth parking orbit. After a coast of approximately 45 minutes, the Breeze stage performs the second of<br />
four propulsive maneuvers. This second main burn is used to begin the process of raising the apogee of the transfer orbit<br />
to geosynchronous altitude.<br />
Due to the thrust of the Breeze M stage (19.6 kN), the optimum <strong>mission</strong> profile results by splitting the "apogee raising"<br />
propulsive maneuver into two smaller burns. The first burn raises apogee altitude to approximately 5,000 km - 7,000<br />
km; the actual value is <strong>mission</strong> and satellite mass dependent. After one full revolution in this intermediate transfer<br />
orbit (approximately 2.5 hours), the third main burn of the Breeze M main engine occurs raising apogee altitude to<br />
geosynchronous. Perigee altitude and inclination are adjusted somewhat based on the trajectory optimization process<br />
that occurs during the <strong>mission</strong> integration process.<br />
During all nonthrusting periods, the fourth stage of Proton is able to align the satellite to an attitude that is compatible<br />
with the thermal and solar incidence requirements for the satellite. Roll maneuvers can be programmed to ensure even<br />
heating/cooling of the satellite surfaces.<br />
After a coast of approximately 5.2 hours, the Breeze M provides its fourth and typically final propulsive maneuver,<br />
raising perigee altitude and lowering orbit inclination to the optimum extent. After completion of the Breeze M fourth<br />
burn, the satellite is oriented and separated from the upper stage. Total elapsed time from <strong>launch</strong> for a typical Breeze<br />
M <strong>mission</strong> is approxiamately 10 hours.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Normally, injection into final orbit occurs at 90 O however, longitude placement for geosynchronous orbits can be<br />
controlled with an increment of 12.25 deg by allowing the Breeze M and payload to coast in the standard parking orbit<br />
for the required number of revolutions, and initiating the first transfer orbit burn on either the ascending or descending<br />
node. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parking orbit<br />
revolutions are within the lifetime constraints of the Breeze M, giving complete 360-deg coverage.<br />
Figure 2.3.3-1 illustrates the main characteristics of the trajectory for a Proton M/Breeze M <strong>launch</strong> to geosychronous<br />
transfer orbit.<br />
Figure 2.3.3-1: Typical Breeze M Flight Profile to Geosynchronous Transfer Orbit<br />
2.3.4 Collision and Contamination Avoidance Maneuver<br />
The Block DM and Breeze M can perform a variety of maneuvers to minimize the possibility of recontact with or<br />
contamination of a customer’s spacecraft. The separation event provides a typical relative velocity between the<br />
spacecraft and the upper stage of 0.3 meter/sec.<br />
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2.3.4.1 Block DM Upper Stage<br />
Approximately 2 ½ hours after spacecraft separation, a 300-sec burn of the SOZ units is performed to increase the<br />
relative separation velocity between the upper stage and the spacecraft. This is followed by a burn of opposing pairs of<br />
SOZ thrusters to SOZ propellant depletion (with zero net delta-velocity). Finally the LOZ tank residuals are vented<br />
through a pair of forward facing nozzles located on either side of the stage near the bottom of the LOZ tank. These<br />
nozzles are canted at 15 degrees from the vertical in such a way as to spin up the stage whole simultaneously adding a<br />
relative velocity increment between spacecraft and upper stage of up to 5 m/sec.<br />
2.3.4.2 Breeze M Upper Stage<br />
Approximately 1 ½ hours after spacecraft separation, the Breeze M stage performs an attitude change maneuver to reorient<br />
the stage. A small propulsive maneuver is made to increase relative velocity between the upper stage and<br />
spacecraft. After completion of the maneuver, the upper stage propellant tanks are depressurized and the stage is made<br />
inert. Relative velocity increments between the spacecraft and upper stage are similar to those for Block DM are<br />
achieved.<br />
2.4 PERFORMANCE GROUNDRULES<br />
A number of standard <strong>mission</strong> groundrules have been used to develop the reference Proton K and Proton M<br />
performance capabilities identified in this document. They are identified in this section.<br />
2.4.1 Payload Systems Mass Definition<br />
Performance capabilities quoted throughout this document are presented in terms of payload systems mass. Payload<br />
Systems Mass (PSM) is defined as the total mass delivered to the target orbit, including the separated spacecraft, the<br />
spacecraft-to-<strong>launch</strong> <strong>vehicle</strong> adapter, and all other hardware required on the <strong>launch</strong> <strong>vehicle</strong> to support the payload<br />
(e.g., harnessing). Table 2.4.1-1 provides masses for the standard Proton adapter systems. Data is also provided for<br />
estimating the performance effects of various <strong>mission</strong>-peculiar hardware requirements. As a note, the performance<br />
effects shown are approximate.<br />
Table 2.4.1-1: Launch Vehicle Mission Peculiar Hardware<br />
∅1666 mm<br />
two piece adapter<br />
∅1194 mm adapter<br />
500 mm height<br />
∅1194 mm adapter<br />
625 mm height<br />
Block DM Breeze M<br />
175 kg (386 lb) ∅1666 mm adapter 175 kg (386 lb)<br />
120 kg (265 lb) ∅1194 mm adapter 150 kg (331 lb)<br />
130 kg (287 lb) TBD adapter 200 kg (441 lb)<br />
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2.4.2 Payload Fairings<br />
All performance quotes in this document are based on use of standard payload fairing options. For Proton K/Block<br />
DM the standard commercial payload fairing is used. For Proton M/Breeze M the baseline “short” payload fairing<br />
option that enables a usable volume similar to Block DM is used. For Proton M/Breeze M, a performance<br />
degradation of approximately 100 kg results for high-energy geosynchronous transfer <strong>mission</strong>s when the “long”<br />
payload fairing option is used.<br />
In all cases, the payload fairing is jettisoned at a location that ensures a free molecular heating rate after jettison of<br />
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
2.6 GEOSYNCHRONOUS TRANSFER MISSIONS<br />
2.6.1 Launch to Geosynchronous Transfer Orbit<br />
The high-energy geosynchronous transfer <strong>mission</strong> is the standard <strong>mission</strong> profile for most commercial Proton<br />
<strong>launch</strong>es. Variable mass satellites are delivered to a geosynchronous transfer orbit with a 35,786 km apogee altitude, a 0<br />
degree argument of perigee, and a variable orbit perigee and inclination consistent with the liftoff mass of the satellite<br />
and the delivered <strong>launch</strong> <strong>vehicle</strong> performance. From this point, the satellite will perform the remaining perigee raising<br />
and inclination reduction to reach geosynchronous orbit.<br />
For reference purposes, ILS has established a reference goesynchronous transfer <strong>mission</strong> performance quotation based<br />
on an orbit that is 1,500 m/sec delta-velocity from geosynchronous orbit. This reference <strong>mission</strong> is indicative of the<br />
geosynchronous transfer <strong>mission</strong>s used by <strong>vehicle</strong>s <strong>launch</strong>ed from low inclination <strong>launch</strong> sites. The reference orbit<br />
assumes a 5,500 km perigee altitude, a 35,786 km apogee altitude, a 25.0 degree orbit inclination, and a 0.0 degree<br />
argument of perigee.<br />
Proton K/Block DM Performance - Performance to elliptical transfer orbits with GSO apogees is shown in Figure<br />
2.6.1-1 for the Proton K/Block DM. Data is shown that represents <strong>launch</strong> <strong>vehicle</strong> performance as a function of<br />
payload systems and residual delta-velocity from targeted transfer orbit to geosynchronous orbit. Analyses have been<br />
conducted to determine the near optimum orbit that can be achieved with Proton given a spacecraft mass and the<br />
various performance variants of the Proton <strong>launch</strong> system. Given a payload mass and <strong>launch</strong> from the Baikonur<br />
Cosmodrome, the upper stage of Proton delivers the satellite to a high-energy GTO that results in the minimum deltavelocity<br />
remaining to reach GSO. The derived perigee altitudes and orbit inclinations are provided in Table 2.6.1-1 for<br />
the standard Block DM 2-Burn <strong>mission</strong> profile. Table 2.6.1-3 provides parametric GTO performance data for a full<br />
range of perigee altitude and orbit inclination combinations.<br />
For specific payload mass ranges, the Proton booster ascent profile can be tailored to enable a suborbital burn of the<br />
Block DM upper stage enabling a higher upper stage/payload mass to be delivered to low earth orbit. The 3-Burn<br />
<strong>mission</strong> is executed by the Proton K booster delivering a 23,000 kg (50,700 lb) orbit unit “stack” to a high-energy<br />
suborbital state. The Block DM, after separation from the third stage of Proton, performs a main engine firing to<br />
achieve a low earth parking orbit. The Block DM <strong>mission</strong> then progresses similarly to the two burn profile. In this<br />
instance, 3 firings of the Block DM main engine result in a higher performance capability results as the Block DM<br />
reaches LEO with a higher propellant load than if it were offloaded and delivered directly to LEO by the booster<br />
statges. Table 2.6.1-2 identifies the optimum geosynchronous transfer orbit versus payload relationship for the<br />
“Enhanced” 3-Burn <strong>mission</strong> capability. The 3-burn profile is enabled with payloads that are greater than or equal to<br />
4,600 kg (10,141 lb). Additionally, the Block DM structural load limitations, as specified elsewhere in this document,<br />
must be adhered to in order for this capability to be enabled.<br />
Proton M/Breeze M Performance – For Proton M/Breeze M, three performance configurations are identified. Figure<br />
2.6.1-2 plots optimum GTO payload versus Delta-Velocity to GEO for these three performance variants of Proton<br />
M/Breeze M. The first configuration (Configuration 1) represents the performance capability for Proton K/Breeze M<br />
for initial flights which are to commence in early 1999. In this configuration, the Proton K booster delivers the Breeze<br />
M to orbit. A payload systems mass of 4800 kg can be delivered to a reference GTO in this case. With introduction of<br />
the Proton M booster, <strong>vehicle</strong> performance to GTO increases to 5200 kg PSM. Flights 5, 6 and 7 of the Breeze M, to<br />
be <strong>launch</strong>ed during the second half of 2000, will deliver this performance capability to GTO and are represented by the<br />
Configuration 2 data. Table 2.6.1-4 provides tabular data corresponding to the data identified in the performance plot.<br />
Additionally, Table 2.6.1-5 provides parametric performance data for a full range of geosynchronous transfer <strong>mission</strong>s<br />
for the mature flight configuration, Configuration 3, of Proton M/Breeze M.<br />
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Figure 2.6.1-1: Proton K/Block DM Performance To Representative Geosynchronous Transfer Orbits<br />
Payload System Mass (kg)<br />
5500<br />
5000<br />
4500<br />
4000<br />
3500<br />
3000<br />
2500<br />
Enhanced<br />
Standard<br />
500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900<br />
Delta-V to GSO (m/s)<br />
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Table 2.6.1-1: Proton K/Block DM Performance To Representative Geosynchronous Transfer Orbits<br />
Delta-V to GSO Orbit Inclination Perigee Altitude Payload Systems Mass (PSM)<br />
(m/s) (deg) Hp (km) (2-Burn)<br />
600 7.0 16600 2732<br />
700 8.2 14400 2887<br />
800 9.7 12600 3046<br />
900 11.3 11000 3210<br />
1000 13.1 9600 3382<br />
1100 15.0 8300 3560<br />
1200 17.0 7200 3740<br />
1300 19.0 6100 3925<br />
1400 21.1 5100 4143<br />
1500 23.3 4200 4350<br />
1600 25.7 3400 4350<br />
1800 31.0 2100 4350<br />
Transfer Orbit Parameters as Specified with Apogee Altitude of 35,786 km;<br />
Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m 2 ;<br />
Launch Mission Duration Approximately 6.5 hours.<br />
Table 2.6.1-2: Proton K/Block DM Three-Burn Mission Performance to Representative Geosynchronous Transfer<br />
Orbits<br />
Delta-V to GSO Orbit Inclination Perigee Altitude Payload Systems Mass (PSM)<br />
(m/s) (deg) Hp (km)<br />
1343.6 19.83 5600 4600<br />
1375.0 20.52 5310 4666<br />
1400.0 21.08 5082 4719<br />
1425.0 21.64 4859 4772<br />
1450.0 22.21 4640 4825<br />
1475.0 22.79 4425 4879<br />
1500.0 23.37 4217 4932<br />
1525.0 23.96 4016 4985<br />
1500.0 25.00 5500 4910<br />
Transfer Orbit Parameters with Apogee Altitude of 35,786 km;<br />
Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m 2 ;<br />
Launch Mission Duration Approximately 6.5 hours.<br />
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Figure 2.6.1-2: Proton M/Breeze M Performance To Representative Geosynchronous Transfer Orbits<br />
Payload Systems Mass (kg)<br />
6500<br />
6000<br />
5500<br />
5000<br />
4500<br />
4000<br />
3500<br />
3000<br />
Config 1<br />
Config 2<br />
Config 3<br />
500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900<br />
Delta-V to GSO (m/s)<br />
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Table 2.6.1-3: Proton K/Block DM Parametric Geosynchronous Transfer Performance Date (2 Burn Mission)<br />
Inclination, deg. 0.0 5.0 10.0 15.0 20.0 25.0<br />
Perigee altitude, km<br />
200 3450 3680 3920 4170 4350 4350<br />
500 3430 3655 3895 4150 4350 4350<br />
1000 3390 3615 3860 4115 4350 4350<br />
2000 3310 3540 3780 4040 4310 4350<br />
4000 3150 3380 3625 3880 4150 4350<br />
6000 3010 3240 3475 3725 3985 4250<br />
8000 2880 3100 3335 3580 3830 4085<br />
10000 2760 2980 3210 3445 3690 3930<br />
12000 2650 2870 3090 3320 3555 3790<br />
15000 2505 2715 2935 3150 3380 3600<br />
19000 2340 2545 2750 2965 3175 3385<br />
23000 2200 2400 2600 2805 3000 3200<br />
27000 2080 2275 2470 2670 2860 3050<br />
35786 1880 2060 2245 2430 2610 2780<br />
Transfer Orbit Parameters as specified with Apogee Altitude of 35,786 km;<br />
Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m 2 ;<br />
Launch Mission Duration Approximately 6.5 hours.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Table 2.6.1-4: Proton M/Breeze M Performance to Representative Geosynchronous Transfer Orbits<br />
Delta-V to GSO<br />
(m/s) Inclination (deg) Perigee Alt (km) Config 1 Config 2 Config 3<br />
600 7.0 16600 3181 3581 3825<br />
700 8.2 14400 3339 3739 3990<br />
800 9.7 12600 3504 3904 4160<br />
900 11.3 11000 3673 4073 4335<br />
1000 13.1 9600 3850 4250 4518<br />
1100 15.0 8300 4036 4436 4710<br />
1200 17.0 7200 4221 4621 4903<br />
1300 19.0 6100 4416 4816 5104<br />
1400 21.1 5100 4616 5016 5311<br />
1500 23.3 4200 4821 5221 5523<br />
1600 25.7 3400 5034 5434 5743<br />
1700 28.3 2700 5276 5676 5993<br />
1800 31.0 2100 5496 5896 6220<br />
1500 25.0 5500 4800 5200 5500<br />
Transfer Orbit Parameters as Specified with Apogee Altitude of 35,786 km;<br />
Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m 2 ;<br />
Launch Mission Duration Approximately 10.0 hours.<br />
Configuration 1: Performance for Proton K/Breeze M Flights (Flights 1, 2, and 3) and contrained performance<br />
for Proton M/Breeze M (Proton/Block DM backup)<br />
Configuration 2: Performance for Proton M/Breeze M Flights 5, 6, and 7<br />
Configuration 3: Performance for Proton M/Breeze M Flights 8 and on<br />
Proton Vehicle Variant Performance (kg)<br />
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Table 2.6.1-5: Proton M/Breeze M Parametric Geosynchronous Transfer Performance Data (Configuration 3)<br />
Inclination, deg. 0.0 5.0 10.0 15.0 20.0 25.0 30.0<br />
Perigee altitude, km<br />
200<br />
500<br />
1000<br />
2000<br />
4000<br />
6000<br />
8000<br />
10000<br />
12000<br />
15000<br />
19000<br />
23000<br />
27000<br />
35786<br />
Transfer Orbit Parameters as specified with Apogee Altitude of 35,786 km;<br />
Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m 2 ;<br />
Launch Mission Duration Approximately 10 hours.<br />
TO BE SUPPLIED<br />
AT A LATER TIME<br />
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Issue 1, Revision 4, March 1, 1999<br />
2.7 ORBIT INJECTION ACCURACY<br />
Table 2.7-1 shows Block DM 3-sigma accuracy predictions for various <strong>mission</strong>s. The accuracy predictions are<br />
enveloping values and <strong>mission</strong>-unique analysis will be performed to verify that payload accuracy requirements are<br />
satisfied. Table 2.7-2 provides similar orbit injection accuracy values for the Breeze M upper stage.<br />
Table 2.7-1: Block DM Upper Stage Orbit Injection Accuracies<br />
200 km Circular<br />
Support Orbit<br />
10000 km Circular<br />
Orbit<br />
5500 km X 36000 km<br />
@ 25.0 deg GTO<br />
Perigee Apogee Inclination Arg of Perigee Longitude of<br />
Ascending Node<br />
Period<br />
± 6 km ± 15 km ± 0.5 O ± 0.25 O ± 0.025 O ±8 sec<br />
± 45 km ± 30 km ± 0.5 O ± 0.25 O ± 0.1 O ± 550 sec<br />
± 400 km ± 150 km ± 0.5 O ± 0.25 O ± 0.5 O ± 100 sec<br />
Eccentricity Longitude Inclination Period<br />
Geostationary 0.009 ± 1 O 0.75 O ± 20 min<br />
Table 2.7-2: Breeze M Upper Stage Orbit Injection Accuracies<br />
200 km Circular<br />
Support Orbit<br />
Perigee Apogee Inclination Arg of Perigee Longitude of<br />
Ascending Node<br />
Period<br />
± 6 km ± 15 km ± 0.025 O ± 0.3 O ± 0.15 O ±8 sec<br />
500 km Circular Orbit ± 5 km ± 5 km ± 0.05 O ± 0.3 O ± 0.15 O ± 100 sec<br />
1000 km Circular<br />
Orbit<br />
5500 km x 36000 km<br />
@ 25.0 deg GTO<br />
± 10 km ± 10 km ± 0.05 O ± 0.3 O ± 0.15 O ± 100 sec<br />
± 400 km ± 150 km ± 0.5 O - - ± 550 sec<br />
Eccentricity Longitude Inclination Period<br />
Geostationary 0.009 ± 1 O 0.75 O ± 20 min<br />
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Issue 1, Revision 4, March 1, 1999<br />
2.8 SPACECRAFT ORIENTATION AND SEPARATION<br />
The Proton is upper stages are capable of aligning the spacecraft and separating in a 3-axis stabilized mode; with a<br />
longitudinal spinup, or with a transverse spinup enabled with either asymmetric separation springs or via upper stage<br />
maneuvering. Tables 2.8-1 through 2.8-3 show the separation and pointing accuracies for the Block DM and Breeze M<br />
upper stages, for all three separation options. The selected payload separation mechanism will affect separation rates.<br />
The orientation and separation conditions are typical values, and <strong>mission</strong> unique analysis will be performed to verify<br />
that payload requirements are satisfied. The Proton upper stagges can accommodate Customer requirements for spin<br />
rates at separation of up to 9.0 deg/sec. Figure 2.8-1 illustrates separation attitude capabilities.<br />
Table 2.8-1: Upper Stage Orbit Injection Accuracies, Option I<br />
Reference Value<br />
SC spin rate about Z sc axis -6 deg/s ± 1%<br />
SC tip-off rate about Y sc and Z sc axes ± 0.8 deg/ s<br />
Relative separation velocity ≥ 0.35 m/s<br />
SC separation pointing vector error ± 5 deg<br />
Table 2.8-2: Upper Stage Orbit Injection Accuracies, Option II<br />
Reference Value<br />
SC spin rate about X sc axis -2 deg/s ± 1%<br />
SC tip-off rate about Y sc and Z sc axes ± 0.7 deg/ s<br />
Relative separation velocity ≥ 0.5 m/s<br />
SC separation pointing vector error ± 5 deg<br />
Table 2.8-3: Upper Stage Orbit Injection Accuracies, Option III<br />
Reference Value<br />
Relative separation velocity ≥ 0.3 m/s<br />
SC tip-off rate about any of SC axes ± 1.8 deg/ s<br />
SC separation pointing vector error ± 5 deg<br />
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Issue 1, Revision 4, March 1, 1999<br />
Figure 2.8-1: SC Separation Attitude<br />
Up p e r S t ag e<br />
L o n g it u d i na l a x i s<br />
Sp a c e c ra ft P o s i t i o n<br />
V e c t o r<br />
2.9 LAUNCH VEHICLE TELEMETRY DATA<br />
T ra ns v e r s e<br />
Sp i n<br />
L o ng it u d i n a l S p i n<br />
X<br />
Y<br />
Su n L i ne<br />
ILS will provide (in Formats I-V , Tables 2.9-1 through 2.9-5 respectively) Launch Vehicle state vector data following<br />
stage 4 insertion into transfer orbit and Spacecraft separation.<br />
Such data is then submitted to the customer by ILS at Baikonur via fax or voice to the Spacecraft Mission Control<br />
Center.<br />
ILS has adopted standard formats regarding orbital state vector data that are provided to the <strong>launch</strong> services customer<br />
during and after the <strong>launch</strong> <strong>mission</strong>. These standard formats enable the satellite operator to properly determine orbital<br />
conditions at various times during the <strong>mission</strong>. The standard data is transmitted to the spacecraft Mission Control<br />
Center at relevant times<br />
The data formats are:<br />
a) Format I - preliminary within 30 minutes after main engine first cut-off, final within 60 minutes. (Table 2.9-1)<br />
b) Format II - within 120 minutes after main engine first cut-off. (Table 2.9-2)<br />
c) Format III - within 30 minutes after Spacecraft separation. (Table 2.9-3)<br />
d) Format IV - within 120 minutes after Spacecraft separation. (Table 2.9-4)<br />
e) Format V - within 30 minutes after deorbit maneuver. (Table 2.9-5)<br />
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Issue 1, Revision 4, March 1, 1999<br />
The data in Formats I and III are preliminary and subject to clarification in Formats II and IV.<br />
Table 2.9-1: Format I - Preliminary State Vector Data Provided Following Upper Stage 1 st Burn<br />
Item Units<br />
Magnitude of �V for main engine 1 st firing m/sec<br />
Upper stage 1 st burn cutoff (actual) Date and Time (hr, min, sec (GMT))<br />
Roll angle deg<br />
Yaw angle deg<br />
Pitch angle deg<br />
Roll angular velocity deg/sec<br />
Yaw angular velocity deg/sec<br />
Pitch angular velocity deg/sec<br />
Table 2.9-2: Format II - Transfer Orbit Parameters Following Upper Stage 1 st Burn<br />
Item Units<br />
Launch time Date and time(hr, min, sec (GMT))<br />
Upper stage 1 st burn cutoff (actual)<br />
Semi-major axis km<br />
Eccentricity ---<br />
Inclination deg<br />
Right ascension of the ascending node deg<br />
Argument of perigee deg<br />
Argument of latitude deg<br />
Perigee altitude km<br />
Apogee altitude km<br />
Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation is<br />
the estimated time of the Block DM 1 st burn cutoff.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Table 2.9-3: Format III - Preliminary Vector Data At Separation Epoch<br />
Item Units<br />
Magnitude of �V for main engine 2 nd firing m/sec<br />
Upper stage 2 nd burn cutoff (actual) Date and Time (GMT)<br />
Actual time of Spacecraft separation from 4 th stage Date and Time (GMT)<br />
Roll angle deg<br />
Yaw angle deg<br />
Pitch angle deg<br />
Roll angular velocity deg/sec<br />
Yaw angular velocity deg/sec<br />
Pitch angular velocity deg/sec<br />
Table 2.9-4: Format IV - Vector Data At Separation Epoch<br />
Item Units<br />
Launch time Date and time(hr, min, sec (GMT))<br />
Separation time Date and time(hr, min, sec (GMT))<br />
Estimated spacecraft separation time Date and time(hr, min, sec (GMT))<br />
Semi-major axis km<br />
Eccentricity ---<br />
Inclination deg<br />
Right ascension of the ascending node deg<br />
Argument of perigee deg<br />
Argument of latitude deg<br />
Perigee altitude km<br />
Apogee altitude km<br />
Spacecraft +X axis right ascension deg<br />
Spacecraft +X axis declination deg<br />
Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation is<br />
the estimated moment of spacecraft separation.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Table 2.9-5: Format V - State Vector Data At Separation Epoch<br />
Item Units<br />
Fourth Stage third ignition time(UTC) for deorbit<br />
maneuver<br />
Fourth Stage third burn cutoff time (UTC) for deorbit<br />
maneuver<br />
De-orbit �V m/s<br />
Date and time(hr, min, sec (GMT))<br />
Date and time(hr, min, sec (GMT))<br />
Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation is<br />
the estimated moment of spacecraft separation.<br />
Within two days following separation, the SC contractor will provide spacecraft derived state vector data to ILS as<br />
shown in Table 2.9-6.<br />
The SC contractor will provide SC rotation data about the SC X,Y and Z axes from the point of SC acquisition until<br />
15 minutes after acquisition. This data will be provided at [Launch + 15 Days].<br />
Table 2.9-6: Spacecraft Supplied Separation Data<br />
Spacecraft Supplied Separation Data<br />
Parameter Units<br />
Separation date (GMT) MDY<br />
Separation time (GMT) HMS<br />
Semi-major axis km<br />
Eccentricity --<br />
Inclination deg<br />
Right ascension of the ascending node deg<br />
True anomaly deg<br />
Perigee radius km<br />
Apogee radius km<br />
Longitude of ascending node deg<br />
Spacecraft separation spin axis relative right ascension deg<br />
Spacecraft separation spin axis declination deg<br />
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Issue 1, Revision 4, March 1, 1999<br />
2.10 MISSION OPTIMIZATION/ PERFORMANCE ENHANCEMENTS<br />
2.10.1 Nonstandard Mission Designs<br />
The <strong>mission</strong>s presented in the previous sections represent standard <strong>mission</strong>s for the Proton K and M Boosters and<br />
Block DM and Breeze M Upper Stages. ILS can take advantage of the unique capabilities of the Proton <strong>launch</strong> system<br />
to design nonstandard <strong>mission</strong> profiles to meet unique <strong>mission</strong> and payload requirements.<br />
2.10.2 Subsynchronous Transfer<br />
Perigee velocity augmentation by the spacecraft can be used to increase payload weight to high Earth orbits. This<br />
maneuver consists of separating the spacecraft in a transfer orbit whose apogee is less than its final desired value. The<br />
spacecraft then performs one or more burns to raise apogee, as well as a final burn to raise perigee. This maneuver is<br />
generally not used with the Proton <strong>launch</strong> system due to the Proton’s high throw weight capability, but it can be<br />
investigated if desired.<br />
2.10.3 Super Synchronous Transfer<br />
Use of a super synchronous transfer trajectory can increase performance into 24-hr orbits. The super synchronous<br />
transfer trajectory takes advantage of the increased efficiency with which the inclination change can be performed at a<br />
higher altitude. A number of <strong>vehicle</strong> hardware and satellite operational constraints interact with <strong>mission</strong>s utilizing<br />
supersynchronous transfer. ILS is able to asses supersynchronous <strong>mission</strong>s by specific request and with detailed satellite<br />
configuration data.<br />
The long coast life and multiple restart capabilities of the Block DM and Breeze M can also assist in constellation<br />
phasing, thereby reducing spacecraft propellant usage.<br />
ILS also has the resources available to develop special hardware items, such as dispensers for multiple spacecraft, to<br />
meet unique <strong>mission</strong> requirements.<br />
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Issue 1, Revision 4, March 1, 1999<br />
3. SPACECRAFT ENVIRONMENTS<br />
This section provides the ground and flight environments applicable for a Proton <strong>launch</strong> campaign and flight.<br />
3.1 THERMAL/HUMIDITY<br />
The thermal and humidity environment for the spacecraft is defined in this section, from transportation from the<br />
Yubeleini Airport through <strong>launch</strong> base processing, <strong>launch</strong> and separation. SC component temperatures to be used for<br />
assessing thermal compatibility will be calculated by analysis using a SC thermal model provided by the Customer.<br />
3.1.1 Ground Thermal Environment<br />
Ambient temperatures at the Baikonur Cosmodrome are provided in Table 3.1.1-1. Facility and transportation<br />
temperatures are provided in Table 3.1.1-2. During transport, the spacecraft is air-conditioned either by Customer<br />
provided air-conditioning equipment or by a railcar mounted thermal control unit. While on the pad, thermal control<br />
is provided by a pad air conditioning system and/or a liquid thermal control system (refer to Section 3.1.1.1 for a<br />
description of the on-pad systems).<br />
Table 3.1.1-1: 3σ Ambient Temperatures at the Baikonur Cosmodrome<br />
Month Max ( O C) Min ( O C)<br />
January 8 -40<br />
February 12 -38<br />
March 24 -28<br />
April 35 -12<br />
May 42 0<br />
June 46 8<br />
July 46 9<br />
August 42 8<br />
September 38 0<br />
October 30 -12<br />
November 22 -30<br />
December 13 -40<br />
Page 3-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 3.1.1-2: Spacecraft Thermal Environment<br />
Location/event Item Temperature<br />
(deg. C)<br />
Temperature Control<br />
Min. Max.<br />
Spacecraft Container Inlet air to 15 30 External air-conditioning unit provided by ILS/Khrunichev.<br />
Fueling Hall<br />
container<br />
Flow rate of 2000 to 8000 cubic m/hr. Monitoring of the<br />
temperature environment in S/C container is the SC<br />
contractor’s responsibility. *<br />
92A-50<br />
Spacecraft 15 25 Building air-conditioning*<br />
Bldg 44<br />
Payload Processing Area<br />
ambient air 17 27<br />
92A-50<br />
Spacecraft 15 25 Building air-conditioning*<br />
Area 31 rm 119<br />
Fourth Stage Integration Area<br />
ambient air 17 27<br />
92A-50<br />
Spacecraft 15 25 Building air-conditioning*<br />
Area 31 rm 100A<br />
ambient air 17 27<br />
Upper Stage fueling station/ Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev.<br />
(Breeze M only)<br />
ambient air<br />
Maximum flow rate to 4000 m3 /hr. *<br />
Bldg 92-1 Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev.<br />
ambient air<br />
Maximum flow rate to 4000 m3 /hr. *<br />
Encapsulated in fairing during Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev.<br />
transportation<br />
ambient air<br />
Preconditioning will be established by Khrunichev (taking<br />
into account SC contractor recommendations) prior to<br />
transport to pad to guarantee temperature range during<br />
erection. Maximum flow rate to 6000 m3 /hr. *<br />
Erection<br />
On-pad with air-conditioning<br />
through umbilical<br />
Spacecraft<br />
ambient air<br />
10 30 No active temperature control is provided.<br />
Block DM<br />
Spacecraft 13 27 Air-conditioning through umbilical, flow rates adjustable<br />
Breeze M<br />
ambient air 13 23 from 5000 to 13000 M 3<br />
/hr.*<br />
On-pad following removal of Spacecraft 10 30 Temperature control panels in fairing.<br />
air-conditioning umbilical ambient air<br />
Temperature could decrease to -5 deg. in Boattail area.<br />
*Temperature is to be preset within the indicated range per SC contractor request and maintained accurate to + 2 O C.<br />
Page 3-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.1.1.1 On-Pad Thermal Control<br />
An external thermal insulation shroud is placed around the fairing prior to pad rollout to provide additional insulation<br />
during the erection of the Launch Vehicle on the pad when there is no active air conditioning. During transportation to<br />
the pad, conditioned air is provided to the spacecraft from the Thermal Control Railcar. At the pad, the air<br />
conditioning is disconnected and the Launch Vehicle is erected. Following erection, the Mobile Service Tower is<br />
brought up to the Launch Vehicle and the pad air conditioning is connected. This pad air conditioning is known as the<br />
“ASTR” for Air System, Thermal Regulation. Total time between disconnection of the Thermal Control Railcar air<br />
conditioning to the connection of on pad air conditioning is 4 hours maximum. A thermal analysis is performed to<br />
verify that under worst case ambient conditions, the spacecraft temperature will not exceed allowable temperature<br />
limits during the erection process.<br />
The on-pad air conditioning system remains active 24 hours a day until approximately 1.5 hours prior to <strong>launch</strong> when<br />
preparations are begun for Mobile Service Tower rollback. To provide thermal conditioning of the fairing after Mobile<br />
Service Tower rollback, a liquid thermal control system is provided in the fairing. This system is known as the<br />
“LSTR” for Liquid System, Thermal Regulation. It consists of radiators on the fairing inside wall connected to<br />
ethylene glycol filled pipes which run to a thermal control system in the <strong>launch</strong> pad complex. This system is activated 3<br />
hours prior to <strong>launch</strong> and purged with dry nitrogen 5 minutes prior to <strong>launch</strong> to insure that the lines are free of liquid<br />
prior to liftoff. Should the <strong>launch</strong> be aborted, the liquid system can be quickly reactivated and the Mobile Service<br />
Tower will be brought up to renew air-conditioning within 2 hours. A schematic of both the liquid and air thermal<br />
control systems is shown in Figures 3.1.1.1-1a and 3.1.1.1-1b along with an approximate operational timeline, for<br />
both the Block DM and the Breeze M upper stages.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Figure 3.1.1.1-1a: Fairing Air and Liquid Thermal Control System Schematic and Operations Timeline (Block DM)<br />
Liquid flow rate = 0.250 - 0.9 m3/h<br />
T = -10 to +80oC deg. C<br />
27<br />
22<br />
17<br />
12<br />
Thermal<br />
cover<br />
mounting<br />
Launch Vehicle installed<br />
on <strong>launch</strong> pad<br />
Winter<br />
Summer<br />
Preliminary<br />
preparation<br />
(12-28 hr)<br />
Transportation<br />
92-1 to<br />
<strong>launch</strong> pad<br />
Ascent<br />
Unit<br />
H<br />
~H/2<br />
LSTR<br />
Liquid system,<br />
thermal regulation<br />
SC<br />
Air outlet<br />
emissivity =
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.1.1.1-1b: Fairing Air and Liquid Thermal Control System Schematic and Operations Timeline (Breeze M)<br />
Liquid flow rate = 0.250 - 0.9 m3/h<br />
de g. C<br />
27<br />
22<br />
17<br />
12<br />
Thermal<br />
cover<br />
mounting<br />
T = -10 to +80 o C<br />
Launch Vehicle installed<br />
on <strong>launch</strong> pad<br />
Winter<br />
Summer<br />
Preliminary<br />
preparation<br />
Air out le t<br />
emissivity =
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.1.1.2 Supplemental Air Conditioning for Spacecraft Battery Charging<br />
Supplemental SC air-conditioning can be provided on the pad during SC battery charging operations if required. Up<br />
to 480 m 3 /hr can be provided through 2 access doors in the fairing at a temperature selectable between 10 and 16 deg.<br />
C. up until the time of Mobile Service tower rollback 1.5 hours prior to <strong>launch</strong>. The air inlets can be positioned in<br />
existing fairing access doors in the aft section of the fairing. See Figure 3.1.1.2-1 below.<br />
Figure 3.1.1.2-1: Supplemental Fairing Air Conditioning Schematic (Representative; detailed design conducted per<br />
customer request)<br />
3.1.2 Ascent<br />
35 o<br />
30<br />
Fairing/Block DM Interface Plane<br />
1200<br />
2100<br />
7.5 o<br />
Battery Radiator Cooling System Nozzle Location<br />
840<br />
160<br />
260<br />
30 o<br />
300<br />
Fairing Access Door<br />
During Ascent, the <strong>launch</strong> <strong>vehicle</strong> will be exposed to aerodynamic heating. Following fairing jettison, the space craft<br />
will be exposed to solar radiation and free molecular heat flux. A thermal analysis will be performed using the<br />
Customer supplied SC thermal model to predict spacecraft temperatures during this phase of the <strong>mission</strong>. The heat<br />
flux density radiated upon the spacecraft by the internal surfaces of the NF should not exceed 500 W/M 2 from the time<br />
of <strong>launch</strong> until NF jettison. For commercial <strong>mission</strong>s, the fairing is jettisoned at 342 - 344 seconds (121 – 125 km<br />
alt.) into flight and the free molecular heat flux shall not exceed 1135 W/M 2 at any time following fairing jettison.<br />
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Issue 1, Revision 4, March 1, 1999<br />
3.1.3 Orbit<br />
Following injection into parking orbit, the spacecraft thermal environment is determined mainly by solar radiation,<br />
albedo and infrared earth radiation flux. Reorientation maneuvers of the Fourth Stage can be programmed to provide<br />
desired sun angles for maintaining thermal control. An integrated thermal analysis is performed to determine<br />
spacecraft temperatures as a function of time throughout the flight up to spacecraft separation.<br />
3.1.4 Humidity<br />
The ground humidity environment is shown in Table 3.1.4-1 below.<br />
Table 3.1.4-1: Ground Humidity Environment<br />
Location/event Item Relaative Humidity<br />
(RH) %<br />
Inside Spacecraft<br />
Container<br />
Min. Max.<br />
Humidity Control<br />
Inlet air to container 35 60 ILS/Khrunichev provided air-conditioning<br />
Fueling Hall Building air 35 60 Building air-conditioning<br />
Payload Processing Area Building air 35 60 Building air-conditioning<br />
Fourth Stage Integration Building air 35 60 Building air-conditioning<br />
Encapsulated in fairing<br />
during transportation<br />
Fairing conditioning in<br />
Bldg 92-1<br />
Air inside fairing 35 60 ILS/Khrunichev provided air conditioning<br />
Air inside fairing 35 60 ILS/Khrunichev provided air conditioning<br />
Erection Air inside fairing 35 60 None<br />
On-pad with airconditioning<br />
through<br />
umbilical<br />
On-pad following removal<br />
of air-conditioning<br />
umbilical<br />
Air inside fairing 0.5 60 Air-conditioning through umbilical, flow rates<br />
Air inside fairing 0.5 60 None<br />
adjustable from 5000 to 13000 M 3 /hr.<br />
Max dewpoint approx 0 deg. C.<br />
Note: Should the relative humidity drop below 30%, SC contractor will be consulted prior to any work beginning in the vicinity of<br />
the spacecraft (inside the payload fairing after encapsulation).<br />
3.1.5 Air Impingement Velocity<br />
The air impingement velocity on the spacecraft surfaces shall not exceed 3m/sec during ground operations following<br />
encapsulation through <strong>launch</strong>.<br />
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Issue 1, Revision 4, March 1, 1999<br />
3.2 CONTAMINATION ENVIRONMENT<br />
3.2.1 Ground Contamination Control<br />
The contamination environment around the spacecraft is controlled by use of Class 100 000 clean room facilities and<br />
strict control of material cleanliness of flight hardware in proximity to the spacecraft. During transportation using ILS<br />
provided air-conditioning systems and while on the pad, air is filtered to provide better than Class 100 000 particle<br />
content. Air cleanliness is monitored regularly in all areas where the SC is present to ensure particle count levels are<br />
maintained within specification. In addition, witness plates can be mounted inside the fairing following encapsulation<br />
to monitor particle fallout inside the fairing up until the day prior to <strong>launch</strong>. The ground contamination environment<br />
around the spacecraft meets the cleanliness levels specified in Table 3.2.1-1.<br />
Table 3.2.1-1: Ground Contamination Environment<br />
Location/event Cleanliness Level<br />
Required*<br />
Comments<br />
Spacecraft container 100 000 ILS/Khrunichev supplied conditioned air.<br />
SC Processing Facilities 100 000 Facility air conditioning<br />
Bldg 92-1 100 000 Payload encapsulated, filtered air provided<br />
Encapsulated in fairing during<br />
transportation and battery charging<br />
100 000 Payload encapsulated, filtered air provided<br />
Erection 100 000 Payload compartment sealed<br />
On-pad with air-conditioning through<br />
umbilical<br />
On-pad following removal of airconditioning<br />
umbilical<br />
* Per FED Std 209E<br />
3.2.2 In Flight Contamination Control<br />
100 000 Filtered air provided<br />
100 000 Payload compartment sealed<br />
The Launch Vehicle systems are designed to preclude in-flight contamination of the SC. The Launch Vehicle<br />
pyrotechnic devices near the spacecraft used for fairing jettison and SC separation have sealed gas chambers and do not<br />
release significant contamination to the outside environment. The fairing liquid thermal control system pipes are<br />
sealed by automatic valves which close at fairing jettison. The third stage retro rocket plume does not result in any<br />
significant particle contact with the spacecraft due to their position on the aft end of the third stage and the orientation<br />
of the retro rocket plume 15 degrees away from the Launch Vehicle longitudinal axis. The thrusters are located on the<br />
aft section of the Fourth Stage and oriented perpendicular to the attitude control longitudinal axis (worst case) such<br />
that the plume does not contact the spacecraft while the spacecraft is attached to the Fourth Stage. Following<br />
separation, the Fourth Stage is attitude controlled to prevent reorientation relative to the spacecraft until the spacecraft<br />
is a safe distance away from the Fourth Stage. Finally, an evaporative cooling system in the Fourth Stage ejects a 20<br />
percent alcohol/water vapor periodically following third/fourth stage separation at the aft end of the Fourth Stage.<br />
This gas is diverted away from the spacecraft and therefor has no contaminating influence.<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.3 PRESSURE<br />
3.3.1 Payload Compartment Venting<br />
During ascent, the payload compartment is vented through 4 venting orifices distributed equally about the cylindrical<br />
portion of the fairing. Maximum rate of pressure drop in the fairing will not exceed 3.5 kPa/s. A representative<br />
pressure drop profile inside the fairing during flight is given in Figure 3.3.1-1.<br />
At the moment of fairing jettison, the pressure across the fairing halves shall not exceed 700 Pa.<br />
The archimedes volume of the spacecraft to be taken into account for the venting analysis will be provided by the<br />
Customer.<br />
Figure 3.3.1-1: Typical Venting Profile During Ascent<br />
Pascals<br />
1.20E+05<br />
1.00E+05<br />
8.00E+04<br />
6.00E+04<br />
4.00E+04<br />
2.00E+04<br />
0.00E+00<br />
0 10 20 30 40 50 60 70 80 90 100 110 120 130<br />
Time, seconds<br />
Time, sec Pressure (Pascals)<br />
Atmosphere Inside Fairing<br />
0 100000 100000<br />
10 98000 99000<br />
20 91000 93000<br />
30 80000 83000<br />
40 65000 70000<br />
50 49000 55000<br />
60 32500 38500<br />
70 19300 26500<br />
80 10000 17000<br />
90 5000 11000<br />
100 2500 7500<br />
110 1000 5000<br />
120 0 3500<br />
130 0 2000<br />
Atmosphere<br />
Inside Fairing<br />
Page 3-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.4 MECHANICAL LOADS<br />
The spacecraft is subject to various types of mechanical loads due to transportation and handling during the <strong>launch</strong><br />
campaign as well as due to the various flight events following liftoff. This section breaks down these events by type of<br />
excitation: quasi-static, sine/random, acoustic and shock.<br />
3.4.1 Quasi-Static Loads<br />
Design Load Factors are provided in Tables 3.4.1.1-1 and Figure 3.4.1.2-1 for use in preliminary design of primary<br />
structure and/or evaluation of compatibility for existing spacecraft with the Proton <strong>launch</strong> <strong>vehicle</strong>. The load factors<br />
were derived for application at the center of gravity of a rigid spacecraft and generate a conservative estimate of flight<br />
maximum interface axial, shear, and bending moments. The actual interface responses experienced during flight are a<br />
function of the static and dynamic characteristics of the spacecraft, but these load factors have generally proven<br />
conservative for spacecraft in the weight range of 2000 - 4800 kg with c.g. height (above the interface) between 1.0 and<br />
m − z<br />
2.0 meters and where > 0.<br />
85 ; where: m=mass, z=c.o.g. location, I=maximum lateral moment of inertia. The<br />
I<br />
spacecraft cantilevered fundamental frequencies are assumed to be a minimum of 10 Hz lateral and 25 Hz axial to<br />
insure applicability of both the ground Transport and the Flight load factors. Spacecraft that do not meet these criteria<br />
will require preliminary analyses to generate loads environments for assessing compatibility with Proton. The Proton<br />
load factors are conservative, but do not include uncertainty factors. It is recommended that the spacecraft analyst use<br />
uncertainty factors in preliminary sizing of primary structure if the new design is not within the family of spacecraft<br />
integrated on Proton.<br />
3.4.1.1 Ground Loads<br />
During ground transportation and handling operations, the spacecraft will be subjected to low frequency loads. Table<br />
3.4.1.1-1 provides the bounding quasistatic loads at the S/C center of mass in longitudinal and lateral axes.<br />
Table 3.4.1.1-1: Ground Limit Quasistatic Load Environment-Transportation and Handling Operations<br />
Operations Accelerations, (g) Safety Factor<br />
X Y Z<br />
SC in container transported as a separate item + 1.0 1.0 + 1.0 + 0.4 1.5<br />
SC transported as part of the Ascent Unit +0.5 1 + 0.5 + 0.4 1.5<br />
SC transported as part of the Proton LV +0.4 1.0 + 0.3 + 0.15 1.5<br />
Handling 0.15 1 + 0.5 1.5<br />
Notes:<br />
In the transportation case, the axes are those of the transporter, namely:<br />
a) X axis runs along the direction of movement.<br />
b) Y axis is downward in direction of gravity.<br />
c) Z is lateral axis in right hand frame.<br />
For handling:<br />
a) Y axis runs vertically along the lifting and lowering direction, respectfully.<br />
b) X axis runs along any lateral directions.<br />
c) Accelerations exist simultaneously along X,Y,Z.<br />
d) All accelerations are specified for a wind velocity W< 20 m/s.<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.4.1.2 Flight Loads<br />
In flight, the spacecraft will be subjected to low frequency input at the base of the spacecraft. The events which produce<br />
the worst case loading are liftoff and 1st/2nd stage separation. Figure 3.4.1.2-1 provides the bounding quasistatic loads<br />
in longitudinal and lateral axes at the SC center of mass. Loads along perpendicular axes are independent.<br />
Static and dynamic accelerations at the SC interface are measured by flight instrumentation and their maximum 3σ<br />
values are as provided in Figure 3.4.1.2-2. Flight loads shall be evaluated by Coupled Loads Analysis using a dynamic<br />
model incorporating the spacecraft and <strong>launch</strong> <strong>vehicle</strong>.<br />
Figure 3.4.1.2-1: Quasi-Static Design Load Factors<br />
(Axial)<br />
Londitudinal, g<br />
5<br />
4<br />
3<br />
2<br />
1<br />
0<br />
-3 -2 -1 0 1 2 3<br />
-1<br />
-2<br />
-3<br />
-4<br />
Lateral, g<br />
Block DM<br />
Breeze M<br />
Page 3-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.1.2-1: Quasi-Static Design Load Factors (Continued)<br />
Block DM Quasi-Static Design Load Factors<br />
Event Longitudinal, g Lateral, g<br />
Lift-off 2.3 2.3 -2.3<br />
0.3 2.3 -2.3<br />
Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.2<br />
1st/2nd stages before separation 4.3 1.4 -1.4<br />
1st/2nd stages after separation (max<br />
compression) 3 1.4 -1.4<br />
1st/2nd stages after separation (max<br />
tension) -2.8 1.4 -1.4<br />
2nd/3rd stages separation 3 0.3 -0.3<br />
3rd/4th stages separation 2.8 0.3 -0.3<br />
Breeze M Quasi-Static Design Load Factors<br />
Event Longitudinal, g Lateral, g<br />
Lift-off 2.3 1.35 -1.35<br />
0.3 1.35 -1.35<br />
Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.2<br />
1st/2nd stages before separation 4.3 0.9 -0.9<br />
1st/2nd stages after separation (max 3 0.9 -0.9<br />
1st/2nd stages after separation (max<br />
tension) -3 0.9 -0.9<br />
2nd/3rd stages separation 3 0.3 -0.3<br />
3rd/4th stages separation 2.8 0.3 -0.3<br />
Page 3-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.1.2-2: Flight Limit Accelerations at the SC Interface<br />
(Axial) Longitudinal,g<br />
5<br />
4<br />
3<br />
2<br />
1<br />
0<br />
-3 -2 -1 0 1 2 3<br />
-1<br />
-2<br />
-3<br />
Transverse,g<br />
Event Longitudinal Longitudinal Transverse<br />
Static g Dynamic g Dynamic g<br />
Liftoff<br />
Wind and Blast<br />
1.5 1.5 -1.5 1.1 -1.1<br />
(Qmax) 2.2 0.5 -0.5<br />
Separation 1st/2nd<br />
Before 1st stage<br />
booster separation 3.6 0.9 -0.9 0.9 -0.9<br />
stages<br />
After 1st stage<br />
booster separation 1 2 -2.8 0.9 -0.9<br />
2nd stage engine cutoff 2.7 0.3 -0.3 0.3 -0.3<br />
3rd stage engine cutoff 2.5 0.3 -0.3 0.5 -0.5<br />
Note: Used only to evaluate measured flight loads.<br />
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Issue 1, Revision 4, March 1, 1999<br />
3.4.2 Sine and Random Vibration Loads<br />
At <strong>launch</strong>, when the propellant valves of the first stage engines are opened, reactive forces act on the liquid propellant<br />
in the tanks (for approximately 0.1 sec) causing <strong>launch</strong> <strong>vehicle</strong> lateral oscillations on the elastic pad supports.<br />
Prevailing oscillation frequencies are approximately 4 Hz with amplitudes of 0.3 g.<br />
The engines operate at a preliminary thrust level that remains constant for approximately 1.6 sec. During this period,<br />
the Launch Vehicle experiences flexible bending oscillations brought about by uneven thrust among the six engines and<br />
unequal offloading of the pad supports. The prevailing frequencies are 5 to 7 Hz.<br />
Longitudinal flexible body oscillations appear simultaneously with frequencies ranging from 5 to 15 Hz. They are<br />
magnified as the engines are throttled up to full thrust within 0.5 sec as the Launch Vehicle leaves the pad.<br />
During first stage flight, lateral dynamic loads are generated by wind gusts superimposed on steady state wind loads<br />
generated by the jetstream. Launch Vehicle longitudinal flexible oscillations are produced at 10-12 Hz by the natural<br />
random pulsation of the engine thrust. There is no pogo phenomenon. The maximum value of these oscillations based<br />
on telemetry measurements is +/- 0.35 g’s.<br />
From 0.5 to 0.6 sec before first stage cutoff, the four second stage engines start up and gain preliminary thrust. Because<br />
of the uneven thrust of the four engines, lateral reaction forces are generated, causing lateral flexible oscillations of the<br />
Launch Vehicle body. These oscillations are influenced additionally by the first stage engines reacting to control system<br />
commands. The first stage cutoff is characterized by an abrupt decay from 90% to 20% within 0.03 sec which causes<br />
significant flexible longitudinal oscillations of the Launch Vehicle second stage, driven by the preliminary thrust of its<br />
own engines. The oscillations are additionally magnified due to the increase in thrust to 100%. These oscillations damp<br />
out within about 3 seconds.<br />
Dynamic loads occuring during the propulsive events following first/second stage separation are enveloped by the<br />
preceding events.<br />
The above dynamic load environment can be represented by the quasi-sinusoidal vibration environment applied at the<br />
SC/LV interface plane shown in Figure 3.4.2-1. This can be considered a flight limit load environment.<br />
Page 3-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.2-1: Equivalent Sine Levels at Spacecraft Interface - Flight Environment<br />
Lateral Axes<br />
Level (g's)<br />
1.2<br />
1<br />
0.8<br />
0.6<br />
0.4<br />
0.2<br />
0<br />
Lateral Axes<br />
0 20 40 60 80 100 120<br />
Longitudinal Axes<br />
Level (g’s)<br />
1.2<br />
1<br />
0.8<br />
0.6<br />
0.4<br />
0.2<br />
0<br />
Frequency (hz)<br />
Longitudinal Axes<br />
0 20 40 60 80 100 120<br />
Frequency (hz)<br />
Frequency (hz) Level (g's)<br />
5 10 0.3<br />
10 20 0.4<br />
20 40 0.6<br />
40 100 0.7<br />
Frequency (hz) Level (g's)<br />
5 8 0.3<br />
8 20 0.8<br />
20 40 0.6<br />
40 100 0.9<br />
Page 3-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
During ground transportation, random excitation is produced by the rail system. The random vibration levels for<br />
different transport configurations are as shown in the following Figures 3.4.2-2, 3.4.2-3, 3.4.2-4 and 3.4.2-5.<br />
Figure 3.4.2-2: Random Vibration Levels-Ground Transportation By Rail, SC In Container And SC Attached To<br />
Ascent Unit<br />
Power Spectral Density (G2/hz)<br />
3.50E-03<br />
3.00E-03<br />
2.50E-03<br />
2.00E-03<br />
1.50E-03<br />
1.00E-03<br />
5.00E-04<br />
0.00E+00<br />
1 10<br />
Frequency (Hz)<br />
100<br />
Frequency (hz) PSD (G Notes:<br />
X-X Y-Y Z-Z X axis is in the direction of movement<br />
2 0.000075 0.000150 0.000150 Y axis is in the vertical direction parallel to gravity field<br />
4 0.000575 0.003300 0.000330 Z axis is in the direction that provide right a handed set<br />
2 /hz)<br />
8 0.002000 0.003200 0.000660<br />
10 0.000600 0.003200 0.000800<br />
14 0.000280 0.000833 0.000330<br />
20 0.000275 0.000150 0.000320<br />
25 0.000275 0.000150 0.000310<br />
30 0.000275 0.000150 0.000300 *Durations are as follows:<br />
- Transportation velocity less than or equal to 15 km/hr<br />
- Levels are at Container Base for container transportation and at<br />
SC/adapter interface for transportation as part of Ascent Unit<br />
35 0.000500 0.000150 0.000185 Transport SC Container Yubeleini to Area 31 120 minutes<br />
40 0.000180 0.000150 0.000037 Transport SC Container Yubeleini to 92A-50 60 minutes<br />
45 0.000125 0.000150 0.000037 Transport Ascent Unit Area 31 to 92-1 120 minutes<br />
50 0.000125 0.000150 0.000037 Transport Ascent Unit 92A-50 to 92-1 5 minutes<br />
Duration<br />
(minutes) * * *<br />
X-X<br />
Y-Y<br />
Z-Z<br />
Page 3-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.2-3: Transportation of SC in Contractor’s Container, Transportation of SC in KhSC Container,<br />
Transportation of Ascent Unit<br />
y<br />
y<br />
y<br />
x<br />
x<br />
x<br />
Page 3-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.2-4: Random Vibration Levels-Ground Transportation by Rail, SC and Ascent Unit Attached to L/V<br />
Power Spectral Density (G2 /hz)<br />
2.00E-05<br />
1.80E-05<br />
1.60E-05<br />
1.40E-05<br />
1.20E-05<br />
1.00E-05<br />
8.00E-06<br />
6.00E-06<br />
4.00E-06<br />
2.00E-06<br />
0.00E+00<br />
1 10 100<br />
Frequency (Hz)<br />
Frequency (hz)<br />
PSD (G<br />
X-X Y-Y Z-Z<br />
2 0.000002 0.000002 0.000002<br />
4 0.000002 0.000002 0.000002<br />
8 0.000002 0.000002 0.000002<br />
10 0.000002 0.000002 0.000002<br />
14 0.000002 0.000020 0.000002<br />
20 0.000010 0.000001 0.000010<br />
25 0.000001 0.000001 0.000001 Notes:<br />
30 0.000001 0.000001 0.000001 X axis is in direction of movement<br />
35 0.000003 0.000001 0.000001 Y axis is in vertical direction parallel to gravity field<br />
40 0.000001 0.000001 0.000001 Z axis is in direction to provide right handed set<br />
45 0.000001 0.000001 0.000001 Transportation velocity less than or equal to 10 km/hr<br />
50<br />
Duration<br />
0.000001 0.000001 0.000001 Levels are at SC to L/V interface<br />
(minutes) 10 10 10<br />
2 /hz)<br />
X-X<br />
Y-Y<br />
Z-Z<br />
Page 3-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.2-4: Transportation of Integrated Proton LV<br />
3.4.3 Acoustic Loads<br />
The <strong>launch</strong> acoustic loads arise from acoustic sound waves generated by the supersonic jets from the first-stage engine<br />
nozzles being diverted by the <strong>launch</strong> pad and flame deflectors. At transonic velocity and maximum aerodynamic drag,<br />
acoustic loads are caused by aerodynamic pressure pulsation effects on the payload fairing surface. The peak acoustic<br />
loads do not act longer than 3 sec at liftoff and 50 sec while passing through the zone of maximum aerodynamic drag.<br />
Acoustic load characteristics normalized to the threshold pressure of 20 µPa are shown in Figure 3.4.3-1.<br />
y<br />
x<br />
Page 3-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.3-1: Max Expected Acoustic Environment (Third Octave)<br />
Level (dB)<br />
140<br />
130<br />
120<br />
110<br />
100<br />
10 100 1000 10000<br />
Frequncy (Hz)<br />
1/3 Octave Band Center Acoustic Levels on<br />
Frequency (hz)<br />
Spacecraft (dB)<br />
25 117<br />
31.5 123<br />
40 127<br />
50 126.5<br />
63 127.8<br />
80 131.6<br />
100 132.4<br />
125 131.3<br />
160 132.1<br />
200 132.1<br />
250 132.1<br />
315 129.9<br />
400 129<br />
500 127<br />
630 124<br />
800 121<br />
1000 119<br />
1250 117<br />
1600 114.5<br />
2000 112.5<br />
2500 111<br />
3150 109<br />
4000 108<br />
5000 107<br />
6300 105.5<br />
8000 104<br />
10000 103.5<br />
OASPL 141.4<br />
Page 3-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.4.4 Shock Loads<br />
Worst case shock levels are introduced into the spacecraft during the firing of the SC/adapter separation system. The<br />
level is dependent on the type of clampband and the clampband tension. For the existing standard adapter<br />
configurations, three specific levels may be encountered as indicated in Figure 3.4.4-1. The 1194 separation systems<br />
use either a 26.6 kN or a 40 kN preload and the shock levels differ accordingly. The 1666 separation systems use a 30<br />
kN preload and the corresponding shock level is as indicated.<br />
Shock loads during transportation in the SC container at the <strong>launch</strong> site shall not exceed the levels provided in Table<br />
3.4.4-1.<br />
Table 3.4.4-1: Shock Loads During Transportation in SC Container<br />
Event Acceleration amplitude, g<br />
X axis runs along the direction of motion + 1.0<br />
Y axis upward vertical axis 1.0 + 1.0<br />
Z lateral axis + 0.4<br />
One Semisinusoidal impulse duration, ms 30<br />
Note: Assumes 5% damping (Q=10)<br />
Page 3-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.4.4-1: Pyroshock Spectrum at Adapter/Payload Interface<br />
g’s<br />
10000<br />
1000<br />
100<br />
100 1000 10000<br />
Frequency (Hz)<br />
Frequency (Hz) g's<br />
40 kN 1194 26.6 kN 1194 30 kN 1666<br />
Sep Syst. Sep Syst. Sep Syst.<br />
100 150 100 150<br />
600 - 1800 -<br />
800 3000 - 3000<br />
1200 5500 - -<br />
1500 6500 - -<br />
2000 5000 5000 -<br />
3000 - - 3000<br />
4000 5000 - -<br />
6000 - - 3500<br />
10000 8000 5000 4000<br />
Q= 10<br />
40 kN 1194 Sep Syst.<br />
26.6 kN 1194 Sep Syst.<br />
30 kN 1666 Sep Syst.<br />
Page 3-22
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
3.4.5 Environmental Test Requirements<br />
3.4.5.1 Acoustic Test Requirements<br />
The spacecraft test requirements are as follow:<br />
Table 3.4.5.1-1: Acoustic Test Requirements<br />
Type of Test Levels Test Duration (seconds)<br />
Acceptance Minimum of Figure 3.4.3-1 in each band 60<br />
Qualification Acceptance levels + 3 db 120<br />
Protoqualification Acceptance levels +3 db 60<br />
3.4.5.2 Static and Sine Test Requirements<br />
Static testing of primary structure is required as a qualification of the structure for flight. This static test must<br />
demonstrate the capability of the structure to withstand the worst case combination of quasi-static loads shown in<br />
Table 3.4.1.1-1 and Figure 3.4.1.2-1 or obtained from coupled loads analysis. Qualification margins for structure<br />
static testing of 1.1 to yield and 1.25 to ultimate must be assumed. For ground handling lift points, qualification<br />
margin to ultimate shall be 1.5 minimum.<br />
Demonstration of the spacecraft secondary structures ability to withstand dynamic loads induced by flight events and<br />
ground transportation is required for qualification and acceptance of the spacecraft design.<br />
The following tests are required:<br />
Table 3.4.5.2-1: Sine Test Requirements<br />
Type of Test Levels Test Frequency Sweep Rate<br />
(octaves/min)<br />
Sine sweep qualification Sine levels in Figure 3.4.2-1 x 1.25 2<br />
Sine sweep<br />
protoqualification<br />
Levels in Figure 3.4.2-1 x 1.25 4<br />
Sine sweep acceptance Levels in Figure 3.4.2-1 4<br />
Page 3-23
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Issue 1, Revision 4, March 1, 1999<br />
Notching should be minimized, and as a test goal should not allow base inputs to decrease below the base equivalent<br />
level produced by the Final Coupled Loads Analysis multiplied by the appropriate Factor of Safety and should not<br />
allow spacecraft response to go below the worst case response predicted by the Final Coupled Loads Analysis<br />
multiplied by the appropriate Factor of Safety.<br />
For Spacecraft secondary structure which does not attain the worst case level predicted from the Final Coupled Loads<br />
Analysis multiplied by the appropriate Factor of Safety, an analysis must show the capability to sustain the CLA result<br />
multiplied by the appropriate Factor of Safety.<br />
3.4.5.3 Shock Test Requirements<br />
A shock test using the adapter clamp system with the flight adapter and spacecraft is required at the SC manufacturer’s<br />
facility in conjunction with a mechanical/electrical fitcheck for first of a kind spacecraft and the first follow-on<br />
spacecraft in a series. For this test the flight band will be tensioned to the same level as will be used for the flight<br />
installation per the manufacturers test procedure. Shock levels will be measured at a location on the adapter 120 mm<br />
above the SC separation plane.<br />
3.5 ELECTROMAGNETIC COMPATIBILITY<br />
Electromagnetic Interference (EMI) e<strong>mission</strong>s and susceptibility of the SC and the LV shall be individually controlled<br />
to the extent necessary to ensure EMC of the fully integrated system.<br />
3.5.1 EMI Safety Margin (EMISM)<br />
The integrated SC/LV system shall be designed to provide EMC with a minimum of 20 dB EMISM (vs. DC no-fire<br />
Threshold) for ordnance circuits and 6dB EMISM overall.<br />
3.5.2 Radiated E<strong>mission</strong>s<br />
The <strong>launch</strong> <strong>vehicle</strong> intentional emissons are described in Tables 3.5.2-1 and 3.5.2-2a and 3.5.2-2b. The SC needs to be<br />
compatible with these e<strong>mission</strong>s.<br />
The <strong>launch</strong> <strong>vehicle</strong> generated and <strong>launch</strong> base spurious EMI sources shall not exceed the levels of Figures 3.5.2-1a and<br />
3.5.2-1b in a plane 1 meter below and parallel to the SC/LV interface plane.<br />
The SC generated and spurious EMI sources shall not exceed the levels of Figure 3.5.2-2 in a plane 1 meter below and<br />
parallel to the SC/LV interface plane.<br />
Page 3-24
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 3.5.2-1: Launch Vehicle RF Characteristics for Proton K/Block DM (TBC)<br />
Parameters First Stage<br />
Transmitter 1<br />
First Stage<br />
Transmitter 2<br />
Second Stage<br />
Transmitter<br />
Third Stage<br />
Transmitter 1<br />
Third Stage<br />
Transmitter 2<br />
Block DM<br />
Transmitter<br />
Carrier Frequency (Mhz) 192 137 240 232 132 923 769<br />
3db Bandwidth (Mhz) .256 3.0 .256 .256 3.0 0.5 0.5<br />
Block DM<br />
Receiver<br />
Modulation type and<br />
characteristics PM APM/FM PM PM APM/FM PM PM<br />
Transmitter output power<br />
(dbW) Max (at carrier freq.) 20.8 11.8 20.8 20.8 11.8 9.0<br />
Receiver sensitivity at the<br />
carrier freq (dbW), Nom. -137<br />
Antenna gain, db -10 -10 -10 -10 -10 -10 9<br />
Antenna description and<br />
polarization<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Circular<br />
OMNI EP<br />
Operating on pad? Yes Yes Yes Yes Yes Yes Yes<br />
Operating in flight? Yes Yes Yes Yes Yes Yes Yes<br />
Table 3.5.2-2: Launch Vehicle RF Characteristics for Proton M (TBC)<br />
Parameters First Stage<br />
Transmitter 1<br />
First Stage<br />
Transmitter 2<br />
Second Stage<br />
Transmitter<br />
Third Stage<br />
Transmitter 1<br />
Third Stage<br />
Transmitter 2<br />
Carrier Frequency (MHz) 192 137 240 232 132 3410<br />
3db Bandwidth (MHz) 0.256 3.0 .256 .256 3.0 40<br />
Modulation type and<br />
characteristics PM APM/FM PM PM APM/FM PM<br />
Transmitter output power<br />
(dbW) Max (at carrier freq.) 20.8 11.8 20.8 20.8 11.8 9.0<br />
Receiver sensitivity at the<br />
carrier freq (dbW), Nom.<br />
Tracking<br />
Antenna gain, dB -10 - 10 - 10 - 10 - 10 -9 within angles of<br />
+70 degrees<br />
-10 within angles of<br />
+85 degrees<br />
Antenna description and<br />
polarization<br />
Omni-directional, circular limited view omni<br />
antenna<br />
Operating on pad? Yes Yes Yes Yes Yes No<br />
Operating in flight? Yes Yes Yes Yes Yes Yes<br />
Page 3-25
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 3.5.2-3: Launch Vehicle RF Characteristics for Breeze M (TBC)<br />
Parameters Breeze M<br />
Telemetry<br />
Breeze M<br />
Telemetry<br />
Breeze M<br />
Telemetry<br />
Breeze M<br />
Receiver<br />
Glonass/ GPS<br />
Carrier Frequency (MHz) 15150 1018.5 1020.5 5750-5760 1570-1640<br />
3db Bandwidth (MHz) 0.5 0.128 0.128 50 55<br />
Modulation type and<br />
characteristics PM FM FM PM PM<br />
Transmitter output power<br />
(dbW) Max (at carrier freq.) 9.0 11.8 11.8<br />
Receiver sensitivity at the<br />
carrier freq (dbW), Nom. - 137 - 137<br />
Antenna gain , dB 21 3 to 10<br />
(Parallel to<br />
Xsc axis)<br />
Antenna description and<br />
polarization<br />
right circular<br />
horn<br />
limited view<br />
omni antenna<br />
3 to 10<br />
(Parallel to<br />
Xsc axis)<br />
limited view<br />
omni antenna<br />
-7 3<br />
limited view<br />
omni antenna<br />
limited view<br />
omni antenna<br />
Operating on pad? No Yes Yes No Yes<br />
Operating in flight? Yes Yes Yes Yes Yes<br />
Page 3-26
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.5.2-1a: Launch Vehicle and Launch Base Pad Narrowband Radiated E<strong>mission</strong>s (Proton K/Block DM) (TBC)<br />
dBm V/m<br />
140<br />
120<br />
100<br />
80<br />
60<br />
40<br />
20<br />
0<br />
100 1000 10000 100000<br />
Frequency (MHz)<br />
Frequency (MHz) dBmV/m 3dB Band width (MHz)<br />
132 98 3<br />
137 74 3<br />
192 84 0.256<br />
232 100 0.256<br />
240 87 0.256<br />
769 25 0.5<br />
923 120 0.5<br />
2000 70 N/A<br />
100000 15 N/A<br />
Page 3-27
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.5.2-1b: Launch Vehicle and Launch Base Pad Narrowband Radiated E<strong>mission</strong>s (Proton M/Breeze M) (TBC)<br />
dBm V/m<br />
140<br />
120<br />
100<br />
80<br />
60<br />
40<br />
20<br />
0<br />
100 1000 10000 100000<br />
Frequency (MHz)<br />
Frequency (MHz) dBmV/m 3dB Band width (MHz)<br />
132 98 3<br />
137 74 3<br />
192 84 0.256<br />
232 100 0.256<br />
240 87 0.256<br />
1000 127 0.5<br />
2000 70 N/A<br />
3400 123 0.5<br />
15150 135 0.5<br />
30000 15 N/A<br />
Page 3-28
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 3.5.2-2: Launch Vehicle and Launch Pad Radiated Susceptibility Limits<br />
dBmV/m<br />
130<br />
110<br />
90<br />
70<br />
50<br />
30<br />
10<br />
1.00E-01 1.00E+00 1.00E+01 1.00E+02<br />
Frequency (GHz)<br />
Frequency (GHz) dBmV/m<br />
0.01 120<br />
0.1 120<br />
0.75 60<br />
1 129<br />
20 80<br />
3.5.3 RF Transmitter/Receiver Systems EMC<br />
Standard RF system compatibility analyses will be performed which shall insure integrated system EMC of<br />
simultaneously operated SC and LV transmitters and receivers during time frames where such operations are necessary.<br />
Page 3-29
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4. SPACECRAFT INTERFACES<br />
4.1 MECHANICAL INTERFACES<br />
4.1.1 Structural Interfaces<br />
The SC to LV structural/mechanical interfaces include a payload adapter, a separation system, umbilical connectors,<br />
separation switches and bonding straps. The structural/mechanical interfaces are defined for each adapter system in<br />
Appendix D of this Mission Planner’s Guide.<br />
The LV coordinate system is shown in Figure 4.1.1-1 with a representative SC.<br />
Figure 4.1.1-1: LV Coordinate System<br />
4.1.2 General SC Structural and Load Requirements<br />
4.1.2.1 Design Criteria<br />
The SC and LV interface structure shall support the SC during the maximum load condition without yielding. The<br />
clearance between the flanges of the SC and the adapter prior to clampband tensioning shall not exceed 0.6 mm. The<br />
geometry of the spacecraft flange is provided in Appendix D for a temperature of 21° C. The surface flatness of the SC<br />
interface ring shall be less than 0.3 mm.<br />
4.1.2.2 SC Stiffness<br />
The spacecraft primary structural stiffness shall be such that the minimum fundamental lateral and axial mode<br />
frequencies shall be greater than 10 Hz and 25 Hz respectively as cantilevered from a rigid interface. The SC/LV<br />
interface is assumed to behave linearly under all loading conditions.<br />
Page 4-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.1.2.3 SC Interface Loads<br />
The SC lifting device and structure shall be capable of lifting the SC plus the payload adapter and the separation<br />
system. Maximum adapter, separation system and other mass to be lifted by SC = 220kg.<br />
Loads affecting the SC at the SC/LV interface include the adapter springs and the SC/LV electrical umbilical<br />
connectors. The adapter spring forces and the SC/LV electrical umbilical connector forces are provided in the adapter<br />
Appendix D of this Mission Planner’s Guide.<br />
4.1.2.4 SC Center of Gravity Offset Requirements<br />
TBS<br />
4.1.3 Fairing Interfaces<br />
This section provides a description of fairing interfaces including generic fairing useable volume, allowable access door<br />
locations and RF window locations.<br />
4.1.3.1 Fairing General Description<br />
A general layout of the Proton M/Block DM Commercial Fairing is provided in Figure 4.1.3.1-1. A general layout of<br />
the Breeze M Standard and Long Commercial Fairings are provided in Figures 4.1.3.1-3 and 4.1.3.1-5.<br />
Figure 4.1.3.1-2 provides the layout for the Proton/Block DM generic useable volume. Figures 4.1.3.1-4 and<br />
4.1.3.1-6 provide the layouts for the Proton/Breeze M Standard and Long Commercial fairing generic useable<br />
volumes. These generic useable volumes do not take into account any specific adapter configuration. Specific adapters<br />
will alter the bottom portion of the useable volume in order to take into account required adapter clearances for<br />
installation and required flight clearances with the adapter structure. Specific useable volumes tailored to individual<br />
adapter systems are provided in the adapter Appendix D of this Mission Planner’s Guide.<br />
The definition of useable volume used throughout this Mission Planner’s Guide is as follows:<br />
Useable Volume: The spacecraft static envelope (maximum dimensions of unloaded spacecraft, including<br />
manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the useable volume, except<br />
where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements due to ground or flight<br />
loads and deviations caused by an imperfect installation of the spacecraft on the Fourth Stage may protrude beyond the<br />
boundaries of this useable volume. An assumption made is that spacecraft dynamic displacements will not exceed 50<br />
mm. This must be verified by coupled loads analysis.<br />
Page 4-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.1.3.2 Fairing Access Locations<br />
Four fairing access doors are located on the lower boattail of the fairing structure of the Block DM <strong>vehicle</strong> and are<br />
dimensioned as shown in Figure 4.1.3.1-1. These doors are nominally used for access to the Block DM. The Customer<br />
may use these doors for access to spacecraft related interface equipment. These access requirements need to be<br />
coordinated and agreed upon with ILS in the <strong>mission</strong> specific ICD. Up to 2 access doors may be provided in the<br />
fairing in the locations shown in Figures 4.1.3.1-1, 4.1.3.1-3 and 4.1.3.1-5 for the Block DM and Breeze M <strong>vehicle</strong><br />
versions. These access locations may be accessed at times coordinated with ILS from the time of fairing encapsulation<br />
up to the beginning of Launch Vehicle fueling on the <strong>launch</strong> pad.<br />
Page 4-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-1a: Proton/Block DM Commercial Fairing General Layout (Sheet 1 of 3)<br />
Page 4-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-1b: Proton/Block DM Commercial Fairing General Layout (Sheet 2 of 3)<br />
Page 4-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-1c: Proton/Block DM Commercial Fairing General Layout (Sheet 3 of 3)<br />
Page 4-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-2a: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)<br />
Page 4-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-2b: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)<br />
Page 4-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-2c: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)<br />
Page 4-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-3a: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 1 of 3)<br />
Page 4-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-3b: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 2 of 3)<br />
Page 4-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-3c: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 3 of 3)<br />
Page 4-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-4a: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)<br />
Page 4-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-4b: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)<br />
Page 4-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-4c: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)<br />
Page 4-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-5a: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 1 of 4)<br />
Page 4-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-5b: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 2 of 4)<br />
Page 4-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-5c: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 3 of 4)<br />
Page 4-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-5d: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 4 of 4)<br />
Page 4-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-6a: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)<br />
Page 4-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-6b: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)<br />
Page 4-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.1.3.1-6c: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)<br />
Page 4-22
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.1.4 GN2/Dry Air Purge Option<br />
For an additional fee the Customer can obtain a GN2/dry air purge at the adapter interface via special pneumatic<br />
fittings. GN2 can be provided via Customer provided gas bottles during operations in the PPF and on the Launch Pad<br />
up to Mobile Service Tower rollback. At this time, the line can be connected to a dry air source running through the<br />
LV to provide a dry air purge up to liftoff. Characteristics of this purge system are as follows:<br />
Item Characteristic<br />
Number of fittings 1<br />
Type fitting pneumatic 0.172 inches internal diameter, 0.281 inches external diameter, 303 CRES material<br />
(provided by customer)<br />
Period of operation a)accessible by Customer during payload operations in PPF up to on pad prior to Mobile<br />
Service Tower rollback (including during transportation operations as mutually agreed upon<br />
between ILS and Customer)<br />
Operational gas Gaseous nitrogen or air<br />
Gas Content<br />
Particulate Size
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2 ELECTRICAL INTERFACES<br />
Electrical interfaces include the SC/LV airborne interfaces, EGSE interfaces, and telemetry/command links.<br />
4.2.1 Airborne Interfaces<br />
Electrical umbilical interfaces are used primarily for providing power to the spacecraft from Customer ground power<br />
supplies located in the Vault under the <strong>launch</strong> pad. They are also used for hardline telemetry and command links<br />
between the spacecraft and the Customer ground support equipment located in the Bunker. Additionally, the Customer<br />
has an option to provide spacecraft measurements to the Launch Vehicle telemetry system on the Fourth Stage via<br />
these umbilicals.<br />
4.2.1.1 Electrical Connectors<br />
Two 37 or 61 umbilical connectors are provided at the Spacecraft interface with the Launch Vehicle adapter. The<br />
connectors are spring-loaded and at separation will disconnect from the adapter.<br />
Appendix D of this Mission Planner’s Guide describes the standard adapters also provides the type, location and<br />
mechanical configuration for these connectors.<br />
Umbilical connector brackets provide +/- 2 mm adjustment in longitudinal direction and +/- 4 mm adjustment in<br />
the lateral axes.<br />
4.2.1.2 Separation Verification<br />
Two diametrically opposed separation microswitches are provided on the top adapter interface flange. Refer to<br />
Appendix D for specific locations and mounting configuration for each specific adapter. At separation, the<br />
microswitches will open a circuit and the Launch Vehicle telemetry will detect this as the separation event.<br />
In addition, continuity loops are provided in each umbilical connector on the spacecraft side. At separation, the<br />
umbilical connectors will disengage, thereby opening these circuits and providing a redundant indication of separation<br />
to the Launch Vehicle telemetry system.<br />
4.2.1.3 Interface Electrical Constraints<br />
All SC and LV electrical interface circuits shall be constrained at least 20 seconds prior to SC separation such that there<br />
is no current flow greater than 100 mA per wire during the Separation event.<br />
4.2.1.4 Accelerometer Measurements<br />
Five accelerometers are mounted near the top of the adapter interface flange to record acceleration from liftoff until<br />
stage three/four separation. Their positions are illustrated in Section 4.2.1.7. Three accelerometers measure<br />
longitudinal loads and two measure lateral loads. Refer to Section 4.2.1.7 for characteristics of these telemetry<br />
channels and Appendix D for the installation of the accelerometers on the adapter.<br />
4.2.1.5 Separation Microswitches<br />
Two diametrically opposed separation microswitches are provided, located on the top adapter interface flange. Refer to<br />
Appendix D for angular locations and for cross-section views.<br />
Page 4-24
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2.1.6 Pre-Separation “Dry Loop” Commands<br />
The Customer may choose as an optional service up to 2 primary and 2 redundant in flight commands in the form of<br />
relay closures for initiating spacecraft commands during flight. The command for closure will be issued after <strong>launch</strong><br />
and before SC/LV separation. Timing and signal characteristic requirements need to be provided by the Customer no<br />
later than at L-12 months. Characteristics of this command are as follows:<br />
Item Characteristic<br />
Type of relay Single pull single throw<br />
Actuation time Any time from liftoff up to separation<br />
Pulse duration up to 210 ms<br />
Timing accuracy +/- 0.03 ms<br />
Allowable max voltage through relay contact at relay closure 36 Volts<br />
Allowable max steady state current through SC/LV interface<br />
contact<br />
1 Ampere<br />
A schematic of this dry loop command is provided in Figure 4.2.1.6-1.<br />
Figure 4.2.1.6-1: Dry Loop Functional Schematic<br />
PO2A<br />
K7 K13 K8 K14<br />
K11<br />
K1<br />
3<br />
K7<br />
K11<br />
K14 K8<br />
K9<br />
K12<br />
K1<br />
5<br />
2 6 1 5 11 15 12 16<br />
J1 J2<br />
K15 K10<br />
K9<br />
K12<br />
K1<br />
6<br />
K1<br />
6<br />
K10<br />
SC/LV Interface<br />
Page 4-25
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2.1.7 Launch Vehicle Telemetry, Command and Power<br />
The <strong>launch</strong> <strong>vehicle</strong> provides the spacecraft separation command and the power for initiating the separation system.<br />
There are no <strong>launch</strong> <strong>vehicle</strong> power or command lines which pass across the spacecraft separation plane.<br />
Table 4.2.1.7-1 provides a description of the telemetry which is used during ground handling. Table 4.2.1.7-2 provides<br />
the characteristics and location of each flight telemetry sensor. Finally, Figures 4.2.1.7-1 to 4.2.1.7-9 show the<br />
locations of each sensor on the LV or ground transportation device.<br />
4.2.1.7.1 Optional SC Telemetry through LV Telemetry System<br />
There is a Customer option for monitoring up to 2 spacecraft parameters during flight by routing up to 2 spacecraft<br />
telemetry points through the umbilicals to the Third and Fourth Stage telemetry system. Characteristics are described<br />
below:<br />
Item Characteristic<br />
Data channel voltage range TBD<br />
Data channel sample frequency 8000 Hz<br />
Data channel measurement window From lift-off command to third stage separation<br />
Allowable round trip impedance from umbilical connector to<br />
spacecraft TM point (ohms) to maintain measurement error<br />
within +/- 7%<br />
Minimum isolation required between signal line and spacecraft<br />
structural ground (Megohms)<br />
2 kohm (any combination of resistance, impedance and<br />
inductance)<br />
Data recorded for these TM points will be provided to the Customer as part of the Post-Launch Report sub<strong>mission</strong><br />
and in electronic format.<br />
TBD<br />
Page 4-26
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.1.7-1: Instrumentation Quantities and Locations for Ground Operations<br />
Accelerations<br />
Accelerometer Location and<br />
Measurement Directions<br />
Amplitude<br />
Measurement<br />
Dynamic<br />
Range, G’s<br />
Frequency<br />
Measurement<br />
Range, Hz<br />
Transport by Motor Vehicle or Location: In the area of the attachment of the container<br />
Rail, Spacecraft Mounted in on shock pallet to the Transport Vehicle<br />
Shipping Container on Shock Vertical<br />
1.0 (TBX1) Up to 50 Hz<br />
Pallet<br />
Lateral<br />
1.0 (TBY1)<br />
Longitudinal<br />
1.0 (TBZ1)<br />
Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface Ring<br />
Mounted on LV Fourth Stage Vertical<br />
1.0 (TBX3) Up to 50 Hz<br />
(Fourth Stage and Fairing Lateral<br />
1.0 (TBY3)<br />
Only)<br />
Longitudinal<br />
Support Point of Fairing Assembly at Cylinder-Nose<br />
Cone Transition<br />
1.0 (TBZ3)<br />
Vertical<br />
1.0 (TBX4) Up to 50 Hz<br />
Lateral<br />
1.0 (TBY4)<br />
Longitudinal<br />
Spacecraft to SCA Separation Joint Interface<br />
1.0 (TBZ4)<br />
Vertical<br />
1.0 (TBX) Up to 50 Hz<br />
Lateral<br />
1.0 (TBY)<br />
Longitudinal<br />
1.0 (TBZ)<br />
Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface Ring<br />
Mounted on Proton Launch Vertical<br />
1.0 (TBX6) Up to 50 Hz<br />
Vehicle Assembly<br />
Lateral<br />
1.0 (TBY6)<br />
Longitudinal<br />
Support Point of Proton First Stage at Aft Ring<br />
1.0 (TBZ6)<br />
Vertical<br />
1.0 (TBX5) Up to 50 Hz<br />
Lateral<br />
1.0 (TBY5)<br />
Longitudinal<br />
Spacecraft to SCA Separation Joint Interface<br />
1.0 (TBZ5)<br />
Vertical<br />
1.0 (TBX) Up to 50 Hz<br />
Lateral<br />
1.0 (TBY)<br />
Longitudinal<br />
1.0 (TBZ)<br />
Temperature Temperature Sensors Location Measurement Range, °C<br />
All ground events 2 Thermocouples mounted on spacecraft support ring -10°C to +40°C (TA5, TA6)<br />
All ground events 2 Thermocouples measuring bulk air temperature inside<br />
Khrunichev container<br />
-10°C to +40°C (TBK1, TBK2)<br />
All ground events Temperature of inlet/exit air from thermal conditioning<br />
car<br />
0-50 °C<br />
On Pad Temperature on Adapter<br />
-50 to 80 °C (TA5, TA6)<br />
Temperature of Air Supplied to Fairing<br />
-50 to 80 °C (T20, T21)<br />
Page 4-27
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.1.7-1: Instrumentation Quantities and Locations for Ground Operations (Continued)<br />
Humidity Humidity Sensors Location Measurement Range,<br />
Transportation in Container Relative humidity inside shipping container 0-90%<br />
All transportation events Relative humidity of inlet, exit air from KhSC thermal<br />
conditioning car<br />
0-80%<br />
Contamination Contamination Sensors Location Measurement Range,<br />
All ground events Particulate size at inlet/exit from air conditioning car (see 0.5 microns/5 microns and<br />
facility drwg)<br />
higher<br />
All ground events Witness Plates (2) located inside fairing on boat-tail section<br />
On Pad Access to fairing air supply for manual reading of 0.5 microns/5 microns and<br />
contamination levels<br />
higher (SENSOR LABEL**)<br />
Page 4-28
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Typical)<br />
Purpose Quantity, measurement<br />
range<br />
Temperature sensors<br />
Environment<br />
under fairing<br />
Ext. surface<br />
thermal<br />
insulation<br />
External<br />
surfaces of the<br />
cellular<br />
construction<br />
skin and<br />
thermal<br />
insulation<br />
1 0...150C<br />
2 -40...100C<br />
2 -40...100C<br />
1 -40...600C<br />
1 -40...500C<br />
1 -40...200C<br />
1 -40...200C<br />
1 -40...200C<br />
2 -40...200C<br />
1 -40...200C<br />
1 -40...200C<br />
1 -40...200C<br />
1 -40...200C<br />
2 -40...200C<br />
1 -40...200C<br />
1 -40...200C<br />
1 -40...100C<br />
1 -40...100C<br />
2 -40...100C<br />
1 -40...100C<br />
1 -40...100C<br />
Sensor designation and location Time of operation Sensor<br />
sampling<br />
frequency<br />
T1 Nose internal surface<br />
T22 (T20) Cylinder, IV pl.<br />
T23 (T21) Cylinder, II pl.<br />
T26 30 degree cone<br />
T25 20 degree cone<br />
T27 Cylinder<br />
External skin<br />
T2 30 degree cone<br />
T3 20 degree cone<br />
T5,T4 Cylinder I pl.<br />
T6 Cylinder III pl.<br />
T7 Inverse cone I pl.<br />
Internal Skin<br />
T8 30 deg. cone I pl.<br />
T9 20 deg. cone I pl.<br />
T11,T10 Cylinder I pl.<br />
T12 20 deg. cone I pl.<br />
T13 Inverse cone I pl.<br />
Thermal insulation<br />
T16 30 deg. cone I pl.<br />
T28 20 deg. cone I pl.<br />
T17,T29 Cylinder I pl.<br />
T18 Cylinder III pl.<br />
T19 Inverse cone I pl.<br />
Panel liquid thermoregulating system<br />
heater (LTS)<br />
T14 20 deg. cone I pl.<br />
T30,T15 Cylinder I pl.<br />
1 -40...100C<br />
1 -40...100C<br />
Construction 4 -10...77C TA1<br />
of adapters<br />
TA2 adapter<br />
TA3<br />
TA4<br />
upper frame<br />
2 -90...100C TA5<br />
TA6 adapter middle<br />
4 -90...100C TA7<br />
TA8 adapter bottom<br />
TA9<br />
TA10<br />
frame<br />
Spacecraft separation sensors<br />
Adapter 2<br />
Separation<br />
plane<br />
J1, J2<br />
2<br />
From liftoff command<br />
to nose fairing separation<br />
(Temperature sensors in<br />
parenthesis are used on<br />
the ground only).<br />
From liftoff command<br />
to spacecraft separation<br />
To liftoff command<br />
(ground sensors)<br />
From liftoff command<br />
to spacecraft separation<br />
To spacecraft<br />
separation<br />
0.3 Hz<br />
1.6 Hz<br />
continuous<br />
1.6 Hz<br />
50-100 Hz<br />
Page 4-29
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Continued)<br />
Purpose Quantity, measurement<br />
range<br />
Nose fairing separation sensors<br />
Nose fairing 2<br />
1-2<br />
4<br />
5-6<br />
7-8<br />
Pressure sensors<br />
4<br />
1-2<br />
3-4<br />
Nose fairing 4 0...780 1 20 degree cone,<br />
mm Hg 2<br />
3<br />
4<br />
external surface<br />
1: 0...400 mm Hg 5 Inverse cone,<br />
1: 0...780 mm Hg 6 external surface<br />
1: 0...780 mm Hg 1 20 degree cone,<br />
1: 0...250 mm Hg 2 internal surface<br />
1: 0...780 mm Hg 3 Inverse cone,<br />
1: 0...50 mm Hg 4 internal surface<br />
4: 0...780 mm Hg 1<br />
6<br />
8<br />
10<br />
Above vents<br />
7: 0...780 mm Hg 2<br />
3<br />
4<br />
5<br />
7<br />
9<br />
11<br />
Below vents<br />
Acoustic loads sensors<br />
Nose fairing<br />
internal and<br />
external<br />
acoustic<br />
pressure<br />
measurement<br />
2 places<br />
2 places<br />
30-2000 Hz<br />
120-155 dB<br />
30-2000 Hz<br />
125-165 dB<br />
Sensor designation and location Time of operation Sensor<br />
sampling<br />
frequency<br />
AB-1 In the internal<br />
AB-2 volume between<br />
SC and nose<br />
fairing. AB-1 on KhSC<br />
adapter, AB-2 on fairing<br />
AH-1 On the nose<br />
AH-2 fairing external<br />
To nose fairing<br />
separation<br />
From liftoff command to<br />
nose fairing separation<br />
From liftoff command to<br />
first stage separation<br />
50-100 Hz<br />
50 Hz<br />
8000 Hz<br />
Page 4-30
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Continued)<br />
Purpose Quantity, measurement<br />
range<br />
Vibration loads sensors (low frequency)<br />
Low frequency<br />
vibration<br />
measurement<br />
along X, Y, and<br />
Z axes<br />
3 places along X<br />
up to 32 Hz<br />
from -2...+4 g<br />
2 places along Y & Z<br />
up to 32 Hz<br />
±0.6 g<br />
Vibration loads sensors (high frequency)<br />
High<br />
frequency<br />
vibration<br />
measurement<br />
along X and Y<br />
axes<br />
2 places<br />
15...2000 Hz<br />
up to 8 g<br />
up to 8 g<br />
Sensor designation and location Time of operation Sensor<br />
sampling<br />
frequency<br />
KX-1 On<br />
KX-2 adapter near<br />
KX-3 SC interface<br />
KY-4<br />
KZ-5<br />
KX-1<br />
KX-2<br />
KX-3<br />
KY-4<br />
KZ-5<br />
BX-CT<br />
BY-CT<br />
On adapter<br />
From liftoff command to<br />
third stage separation<br />
At time of each stage<br />
separation<br />
From liftoff command to<br />
third stage separation<br />
200 Hz<br />
200 Hz<br />
200 Hz<br />
200 Hz<br />
200 Hz<br />
400 Hz<br />
200 Hz<br />
200 Hz<br />
100 Hz<br />
100 Hz<br />
8000 Hz<br />
Page 4-31
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-1a: Locations of Measurement System Sensors on Block DM Fairing (Typical)<br />
Page 4-32
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-1b: Locations of Measurement System Sensors on Breeze M Fairing (Typical)<br />
TBS<br />
Page 4-33
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-2: Locations of Measurement System Sensors on 1666 Adapter System (Typical)<br />
Page 4-34
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-3: Locations of Measurement System Sensors on 1194 Adapter System (Typical)<br />
TBS<br />
Page 4-35
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-4: Instrumentation During Transportation of SC in Contractor’s Container<br />
Rail Transportation of SC in SC contractor’s container from Yubeleini to SC Processing Facility (70 km at<br />
≤ 15 km/hr)<br />
Thermal Rail Car<br />
Air Temperature<br />
Relative Humidity<br />
Particle Count<br />
Conditioned Air Inlet<br />
TBX1, TBY1, TBZ1<br />
Figure 4.2.1.7-5: Instrumentation During Transportation of SC In KhSC Container<br />
Conditioned Air Exhaust<br />
Rail Transportation of SC in KhSC container to/from Fueling Hall (0.5 km at ≤ 5 km/hr). (For fueling operations in<br />
Area 31 only)<br />
Thermal Rail Car<br />
Air Temperature<br />
Relative Humidity<br />
Particle Count<br />
Conditioned Air Inlet<br />
TBX1, TBY1, TBZ1<br />
Exhaust to Atmosphere (2)<br />
T21<br />
(internal to container<br />
opposite to T20)<br />
T20<br />
(internal to container<br />
opposite to T21)<br />
y<br />
y<br />
x<br />
Page 4-36<br />
x
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-6: Instrumentation During Transportation of Ascent Unit<br />
Rail Transportation of SC with upper stage from SC processing facility to area 95: 70 km at ≤ 15 km/hr.<br />
Thermal Rail Car<br />
Air Temperature (inlet, exit)<br />
Relative Humidity (inlet, exit)<br />
Particle Count (inlet, exit)<br />
Conditioned Air Inlet<br />
TBX, TBY, TBZ<br />
TBX1, TBY1, TBZ1<br />
TA5, TA6<br />
Figure 4.2.1.7-7: Instrumentation During Integration of Ascent Unit To LV<br />
T20, T21<br />
TBX3, TBY3, TBZ3<br />
Conditioned Air Exhaust<br />
Temp., Humidity, Particle Count and Witness Plate measurements during the Ascent Unit mate to the LV.<br />
Exhaust to Atmosphere<br />
TBX, TBY, TBZ Conditioned Air Inlet (2)<br />
TA5, TA6<br />
T20, T21<br />
Witness Plates<br />
y<br />
x<br />
Air Temperature<br />
Relative Humidity<br />
Particle Count<br />
y<br />
Page 4-37<br />
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Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-8: Instrumentation During Transportation of Integrated Proton LV<br />
Temp., Humidity, Particle Count and Witness Plate measurements during transport from Area 95 to the Launch Pad.<br />
Conditioned Air Exhaust<br />
TBX, TBY,TBZ<br />
Filter Block<br />
T20, T21<br />
TBX3, TBY3, TBZ3<br />
TBX1, TBY1, TBZ1<br />
Air Temperature (inlet, exit)<br />
Relative Humidity (inlet, exit)<br />
Particle Count (inlet,exit)<br />
Witness Plates<br />
Conditioned Air Inlet<br />
TA5, TA6<br />
Page 4-38
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.1.7-9: Instrumentation During On-Pad Operations<br />
Temperature, Humidity, Particle Count, and Witness Plate measurements while on the <strong>launch</strong> pad.<br />
Temperature inside<br />
Fairing, T20 and T21<br />
(see table for description,<br />
see fairing schematic<br />
for exact location).<br />
Temperature on Adapter,<br />
TA5 and TA6<br />
(see table for description,<br />
see fairing schematic<br />
for exact location).<br />
Fairing Air Exhaust to Atmosphere<br />
(two at 420 x 500)<br />
Witness Plates<br />
(access door on Fairing boat tail)<br />
Conditioned Air Inlet to Fairing<br />
(Air sample access: continuous<br />
monitoring of air used to measure<br />
temp., humidity, and contamination<br />
levels)<br />
AIR<br />
THERM<br />
AL CONTROL<br />
SYSTEM<br />
(ATCS)<br />
Temperature<br />
Humidity<br />
Particle Count<br />
Page 4-39
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2.2 Launch Pad EGSE Interfaces<br />
The two interface connectors described in Section 4.2.1 are wired to a Mission Specific wiring harness on the adapter<br />
which is connected to the LV flight umbilical harness running the length of the <strong>vehicle</strong> to an interface connector Ø06 at<br />
the bottom of the first stage. From here, ground cabling connects the umbilical to an interface panel in the Vault under<br />
the <strong>launch</strong> pad where the Customer electrical interface equipment is located. As can be seen from Figure 4.2.2-1, there<br />
are test access connectors on the Fourth Stage which permit access to the umbilical from the Mobile Service Tower up<br />
to 8 hours prior to <strong>launch</strong>. These can be used to interface Customer battery charging power supplies on the Mobile<br />
Service Tower with the SC. They can also be used to connect with wiring in the Mobile Service Tower to provide a<br />
parallel path with the flight LV umbilical to reduce overall resistance drop from the SC to Customer GSE for high<br />
current power lines.<br />
The Launch Pad interfaces include connections from the base of the Proton LV (and connections at station 43.85 on<br />
the Mobile Service Tower if required) to ground wiring interfacing with SC EGSE. ILS provides all necessary<br />
electrical harnesses and cables between the SC/LV IFD’s and the SC EGSE interface enables in the Vault and on the<br />
MST. Figure 4.2.2-1 provides a block diagram of the electrical interfaces available between the payload, Launch<br />
Vehicle and the ground systems.<br />
Figure 4.2.2-1: Spacecraft to Launch Vehicle and Ground Systems Electrical Interfaces<br />
O 1A<br />
O 06<br />
Spacecraft<br />
37 or 61 pin Deutsch<br />
Connectors<br />
Mobile Service<br />
Tower<br />
Mission Specific<br />
Wiring<br />
Level 6<br />
Junction<br />
Box X1<br />
Mission Specific Wiring<br />
Vault<br />
Customer<br />
MPS<br />
Notes: 1) Cable length from level 6 of MST to Vault: 400 meters<br />
2) Cable length from SC connectors to O 06: 50 meters<br />
3) Cable length from O 06 to Vault: TBD meters<br />
Bunker Rm 250<br />
Customer<br />
STE<br />
Page 4-40
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2.2.1 EGSE Interface Electrical Constraints<br />
The maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.<br />
SC contractor equipment needs to provide protection against exceeding 100 V at spacecraft umbilical interface. It also<br />
needs to provide continuous monitoring and recording of this bus voltage.<br />
All EGSE interfaces to the SC shall be current limited by the Customer to preclude damage to the LV ground and<br />
airborne systems in the event of a short circuit . SC STE shall shut off power within 0.2 sec if I max is exceeded by 50%.<br />
SC and GSE will be de-energized prior to mating and demating of umbilical connectors for electrical checkouts and<br />
flight mating.<br />
All SC EGSE electrical interface circuits shall be constrained at least 5 minutes prior to liftoff such that there will be<br />
no current flow greater than 100 mA per wire across the T-0 interface .<br />
4.2.3 Telemetry/Command Links<br />
An RF link for telemetry/command is provided between the SC on the pad and the Bunker. The Customer can choose<br />
between one of the 4 links defined in Tables 4.2.3-1a, 4.2.3-1b, 4.2.3-1c and 4.2.3-1d. A block diagram of the SC to<br />
bunker RF/electrical interface is shown in Figure 4.2.3-1.<br />
In order to confirm compatibility with the link, the following is required of the Customer:<br />
a) The SC Checkout Station shall have 1 physical command interface and 1 physical telemetry interface.<br />
b) SC GSE RF interface impedance shall be 50 OHM<br />
c) Uninterrupted operation of RF devices shall not exceed 8 hours, with a 30 minute break before the next 8 hour<br />
session.<br />
d) The SC contractor shall provide to Khrunichev the measured coefficient values for TT &C signals via the RF<br />
window obtained during the RF channel checkup in the integration facility following the Ascent Unit<br />
encapsulation.<br />
e) Prior to installation of the LV+Ascent Unit on the pad and following the delivery of the STE to the bunker, the<br />
SC manufacturer shall verify continuity between command RF link and STE and issue to Khrunichev the<br />
Certificate of Launch Pad Readiness to accommodate the LV and Ascent Unit.<br />
f) After installation of LV+Ascent Unit and prior to the roll-up of service tower, the SC manufacturer, in<br />
conjunction with Khrunichev, shall check out the RF link between the Ascent Unit and STE. Such check out shall<br />
be performed 20 minutes after the mating of the LV aft section. SC contractor to confirm functionality of the RF<br />
link within 45 minutes.<br />
g) At L-6 months, the SC manufacturer shall provide to Khrunichev two connectors for installation by Khrunichev<br />
on the existing RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructions<br />
on cable dressing and cable performances.<br />
Page 4-41
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.3-1a: C-Band RF Link Characteristics<br />
Telemetry Link<br />
Reference Value Note<br />
Frequency range, GHz 3.95 ± 0.2<br />
Bandwidth, MHz > 250<br />
Signal polarization<br />
Radio link output signal power<br />
Left-hand circular<br />
maximum. dBm minus 0.0 with Service Tower rolled back<br />
minus 6.2 with Service Tower in place<br />
minimum, dBm minus 30.0 with Service Tower rolled back<br />
minus 36.2 with Service Tower in place<br />
Radio link gain factor, dB minus 31.0 ±5 with Service Tower rolled back<br />
minus 42.2 ± 2 with Service Tower in place<br />
Gain factor adjustment limit for<br />
30<br />
radio link input, dB<br />
Radio link output signal-to-noise<br />
64<br />
ratio, dB*Hz<br />
Command Link<br />
Reference Value Note<br />
Frequency range, GHz 6.42 ± 0.05<br />
Bandwidth, MHz > 200<br />
Signal polarization<br />
Radio link output signal power<br />
Right-hand circular<br />
maximum, dBW/m2 minus 22.2 with Service Tower rolled back<br />
minus 38.6 with Service Tower in place<br />
minimum, dBW/m2 minus 72.2 with Service Tower rolled back<br />
minus 88.6 with Service Tower in place<br />
Radio link gain factor, dB minus 43.2 with Service Tower rolled back<br />
minus 38.6 with Service Tower in place<br />
Gain factor adjustment limit for radio<br />
link input, dB<br />
30<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
70<br />
Page 4-42
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.3-1b: Ku-Band RF Link 1 Characteristics<br />
Telemetry Link<br />
Reference Value Note<br />
Frequency range, GHz 11.1 ± 0.2<br />
Bandwidth, MHz > 250<br />
Signal polarization Linear vertical<br />
Radio link output signal power with SC antenna input signal power of 0 dBW<br />
maximum, dBm minus 31 with Service Tower rolled back<br />
minus 37 with Service Tower in place<br />
minimum, dBm minus 41 with Service Tower rolled back<br />
minus 41 with Service Tower in place<br />
Radio link gain factor, dB minus 78.3 ± 5 with Service Tower rolled back<br />
minus 80.0 ± 2 with Service Tower in place<br />
Gain factor adjustment limit for radio<br />
link input, dB<br />
30<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
118<br />
Command Link<br />
Reference Value Note<br />
Frequency range, GHz 14.0 ± 0.05<br />
Bandwidth, MHz > 200<br />
Signal polarization Linear, horizontal<br />
Radio link output signal power with SCS antenna input signal power of<br />
3 dBW<br />
maximum, dBW/m2 minus 55.5 with Service Tower rolled back<br />
minus 61.5 with Service Tower in place<br />
minimum, dBW/m2 minus 65.5 with Service Tower rolled back<br />
minus 65.5 with Service Tower in place<br />
Radio link gain factor, dB minus 78.3 with Service Tower rolled back<br />
minus 80.0 with Service Tower in place<br />
Gain factor adjustment limits for radio from -65 to -41<br />
link input, dB<br />
from -67 to -43<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
123<br />
Page 4-43
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.3-1c: K-Band RF Link 2 Characteristics<br />
Telemetry Link<br />
Reference Value Note<br />
Frequency range, GHz 12.2 ± 0.2<br />
Bandwidth, MHz > 250<br />
Signal polarization left-hand circular<br />
Radio link output signal power with SC antenna input signal power of 0 dBW<br />
maximum, dBm 2 with Service Tower rolled back<br />
minus 3 with Service Tower in place<br />
minimum, dBm minus 8 with Service Tower rolled back<br />
minus 7 with Service Tower in place<br />
Radio link gain factor, dB minus 35 ± 5 with Service Tower rolled back<br />
minus 45 ± 2 with Service Tower in place<br />
Gain factor adjustment limit for radio<br />
link input, dB<br />
30<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
118<br />
Command Link<br />
Reference Value Note<br />
Frequency range, GHz 14.0 ± 0.05<br />
Bandwidth, MHz > 200<br />
Signal polarization Right-hand circular<br />
Radio link output signal power with SCS antenna input signal power of 3 dBW<br />
maximum, dBW/m2 minus 31 with Service Tower rolled back<br />
minus 36 with Service Tower in place<br />
minimum, dBW/m2 minus 41 with Service Tower rolled back<br />
minus 40 with Service Tower in place<br />
Radio link gain factor, dB minus 78.3 with Service Tower rolled back<br />
minus 80.0 with Service Tower in place<br />
Gain factor adjustment limits for radio from -50 to -80<br />
link input, dB<br />
from -52 to -82<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
123<br />
Page 4-44
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table 4.2.3-1d: Ku-Band RF Link 3 Characteristics<br />
Telemetry Link<br />
Reference Value Note<br />
Frequency range, GHz 12.45 ± 0.25<br />
Bandwidth, MHz > 500<br />
Signal polarization<br />
Radio link output signal power<br />
Linear horizontal<br />
maximum, dBm minus 8.0 with Service Tower rolled back<br />
minus 14.4 with Service Tower in place<br />
minimum, dBm minus 44.0 with Service Tower rolled back<br />
minus 49.4 with Service Tower in place<br />
Radio link gain factor, dB minus 74.0 with Service Tower rolled back<br />
minus 79.4 with Service Tower in place<br />
Gain factor adjustment limit for radio<br />
link input, dB<br />
30<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
118<br />
Command Link<br />
Reference Value Note<br />
Frequency range, GHz 17.3 ± 0.05<br />
Bandwidth, MHz > 200<br />
Signal polarization<br />
Radio link output signal power<br />
Linear vertical<br />
maximum, dBW/m2 minus 59.0 with Service Tower rolled back<br />
minus 59.1 with Service Tower in place<br />
minimum, dBW/m2 minus 89.0 with Service Tower rolled back<br />
minus 89.1 with Service Tower in place<br />
Radio link gain factor, dB minus 93.0 with Service Tower rolled back<br />
minus 96.0 with Service Tower in place<br />
Gain factor adjustment limit for radio link<br />
input, dB<br />
30<br />
Radio link output signal-to-noise ratio,<br />
dB*Hz<br />
155<br />
Page 4-45
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 4.2.3-1: SC to Bunker RF/Electrical Interface Block Diagram<br />
A048<br />
Launc h Pad<br />
SC<br />
Underground Vault<br />
Interconnect<br />
Panel<br />
4.2.3.1 RF Window Requirements<br />
RF link<br />
Interconnect<br />
Panel<br />
Interconnect<br />
Panel<br />
RF TM<br />
RF TC<br />
Bunker<br />
Figures 4.1.3-1, 4.1.3-3 and 4.1.3-5 show locations of fairing doors and RF window cutouts for the fairing.<br />
Khrunichev<br />
RF<br />
Rack<br />
There are 2 RF window positions in the fairing to take into account the possible view angles required at each of the 2<br />
Proton <strong>launch</strong> pads. When the <strong>launch</strong> pad is designated, 1 out of the 2 windows will be covered with conductive enamel<br />
leaving 1 active window for trans<strong>mission</strong> of the spacecraft T and C signal between the spacecraft and the Bunker. The<br />
RF link between the spacecraft and the ground RF equipment is required before and after Mobile Service Tower<br />
rollback. Prior to rollback, the signals will be transmitted via a repeater on the Mobile Service Tower. Following<br />
rollback, signals will be transmitted directly between the spacecraft and the Bunker antenna. During rollback, there<br />
could be a maximum 20 minute outage of signal as the tower rolls through the line of site between the spacecraft and<br />
the bunker antennas.<br />
Page 4-46
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
4.2.4 Electrical Grounding<br />
All payload preparation areas used by the SC as a <strong>launch</strong> base facilities used by SC and SC EGSE are equipped with<br />
earth referenced steel ground busses with equipment attach points (threaded studs). The resistance between any point<br />
on these bars and the building earth ground is less than 4 ohms. The floor surfaces in the payload and hazardous<br />
payload processing areas is anti-static and connected to the facility grounding system. The SC contractor shall provide<br />
all cables and attachment hardware required to interconnect the SC and support equipment with facility grounds.<br />
4.2.5 Electrical Bonding<br />
The resistance across the spacecraft/adapter separation plane shall be less than 10 milliohms at a current less than 10<br />
milliamps. This may be accomplished either by conductive surface contact between the spacecraft and adapter<br />
interface ring (1666 adapters) or by the use of 2 bonding straps which incorporate a friction contact connector that<br />
releases upon spacecraft separation with a separation force of 40 ± 5 N. Signal and power grounds from the spacecraft<br />
are passed through the umbilical without connecting them to Launch Vehicle structure. Likewise umbilical shield<br />
grounds are isolated from the Launch Vehicle structure.<br />
4.2.6 SC/LV Lightning Protection<br />
All payload preparation areas used by the SC will be equipped with a lightning protection system for direct and<br />
indirect hits. Augmentation of the standard provisions for any necessary SC individual circuit protection shall be<br />
provided by the SC contractor.<br />
4.2.7 Static Discharge<br />
During the entire flight through SC separation, no electrostatic discharge shall occur from either the LV or the SC<br />
surface through the LV to SC interface plane.<br />
4.3 FITCHECK OF MECHANICAL/ELECTRICAL INTERFACES<br />
A fitcheck of electrical/mechanical interfaces with the flight adapter and spacecraft is required at the SC<br />
manufacturer’s facility for first of a kind spacecraft and the first follow-on spacecraft in a series.<br />
Page 4-47
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
5. MISSION INTEGRATION AND MANAGEMENT<br />
ILS provides the Customer a SOW which define the management approach for a Customer Launch Service<br />
Agreement, the deliverables provided to the Customer during the course of the LSA and a schedule for all Mission<br />
Integration activities. This section highlights these provisions.<br />
5.1 MANAGEMENT PROVISIONS<br />
5.1.1 Key Personnel<br />
Immediately after execution of each LSA, ILS and the Customer will designate their respective Mission Managers<br />
who will be responsible for performing all management functions related to the LSA.<br />
ILS shall ensure that personnel necessary for the performance of this contract are made available to the program to<br />
perform the work in a timely fashion and to satisfy requirements of the contract and its exhibits.<br />
5.1.2 Interface Control Document (ICD)<br />
The ICD will be created by ILS based on a generic ICD template and the Customer provided IRD. It will provide the<br />
Customer's technical requirements for the <strong>launch</strong> of their spacecraft and characteristics and constraints of the Launch<br />
Vehicle and Launch Site relating to the interface with the spacecraft.<br />
5.1.3 Schedule Monitoring<br />
ILS will create and maintain an interface activities milestone schedule that provides all key technical interface<br />
milestones necessary for successful completion of the contract. A typical <strong>mission</strong> integration schedule is as shown in<br />
Figure 5.1.3-1a and 5.3.1-1b for a non-recurring and recurring program respectively. It is based on the meeting<br />
schedule in Section 5.1.5.1 and the deliverable milestones provided in Sections 6 and 7. This integrated program<br />
schedule for a particular program will be presented and agreed upon between all Parties at the Kickoff Meeting and<br />
further changes shall be made as necessary and agreed upon at subsequent Technical Interface Meetings (TIM). In case<br />
of changes to internal schedules, the other party will be promptly informed.<br />
Page 5-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 5.1.3-1a: Baseline Integration Schedule (Non-Recurring Program)<br />
ID # Activity Resp<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
7<br />
8<br />
9<br />
10<br />
11<br />
12<br />
13<br />
14<br />
15<br />
16<br />
17<br />
18<br />
19<br />
20<br />
21<br />
22<br />
23<br />
24<br />
25<br />
26<br />
27<br />
28<br />
29<br />
30<br />
31<br />
32<br />
33<br />
34<br />
35<br />
36<br />
37<br />
38<br />
39<br />
Kickoff<br />
Preliminary ICD Review<br />
Preliminary Design Review (PDR)<br />
Launch Site Visits<br />
Operations Review<br />
ICD Sign-Off<br />
Fitcheck<br />
Shock Test<br />
Critical Design Review (CDR)<br />
Launch Vehicle Preshipment Rev<br />
Spacecraft Preshipment Review<br />
Launch Site Acceptance Review<br />
Launch Vehicle Quality Review<br />
Update LV Roll-out Authorization Review<br />
Board LV Loading Authorization Review<br />
Board<br />
IRD (Amendment for specific SC)<br />
Preliminary ICD<br />
ICD prior to meeting to sign<br />
Preliminary SC Dynamic Model SCC<br />
Annex [x] of PDR package, CLA ILS<br />
Report SC Thermal Model<br />
SCC<br />
PDR Package, all except Annex [x] ILS<br />
Inputs to Final Analysis (except SCC<br />
CLA) Final SC Dynamic Model (test SCC<br />
verified) Annex [x] of PDR package; CLA ILS<br />
Report CDR Package; all except Annex [x] ILS<br />
SC Environmental Test Plan SCC<br />
Inputs to Fitcheck & Shock Test PlanSCC<br />
Fitcheck & Shock Test Plan ILS<br />
Fitcheck & Shock Test Report ILS<br />
SC Environmental Test Results SCC<br />
Safety Sub<strong>mission</strong>s<br />
All Certificates<br />
SC Launch Operations Plan SCC<br />
(Preliminary)<br />
SC Launch Operations Plan (Final) SCC<br />
Participants List w/ Passport Info SCC<br />
SC Orbital Data<br />
ILS<br />
Launch Evaluation Report<br />
ILS<br />
Printed: 31AUG98<br />
Management Report<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
SCC<br />
ILS<br />
ILS<br />
SCC<br />
All<br />
ILS<br />
STANDARD INTEGRATION SCHEDULE-NON-RECURRING<br />
Revised: 08/21/98<br />
L-24 L-21 L-18 L-15 L-12 L-9 L-6 L-3 LAUNCH L+2<br />
MEETINGS<br />
Kickoff<br />
ICD Review<br />
PDR<br />
Site Visit No.1 Site Visit No.2<br />
Operations Review #1 Operations Review #2<br />
ICD Sign-Off<br />
Fitcheck<br />
Shock Test<br />
CDR<br />
LV Preship<br />
SC Preship<br />
Site Review<br />
LVQR Update<br />
LV Roll-out Auth<br />
LV Loading Auth<br />
ICD DEVELOPEMENT<br />
IRD (Amendment for specific SC)<br />
Prelim ICD<br />
Signed ICD<br />
PRELIMINARY DESIGN<br />
Preliminary SC Dynamic Model<br />
CLA Report (if required)<br />
SC Thermal Model<br />
PDR Package<br />
CRITICAL DESIGN<br />
Inputs to Final Analysis (except CLA)<br />
Final SC Dynamic Model<br />
CLA Report<br />
CDR Package<br />
SPACECRAFT TESTING<br />
SC Environmental Test Plan<br />
Inputs to Fitcheck & Shock Test Plan<br />
Fitcheck & Shock Test Plan<br />
Fitcheck & Shock Report<br />
SC Environmental Test Results<br />
SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSING<br />
Safety Sub<strong>mission</strong>s<br />
All Certificates<br />
LAUNCH CAMPAIGN<br />
SC Launch Ops Plan (Prelim)<br />
SC Launch Ops Plan (final)<br />
Participants List w/ Passport Info<br />
SC Orbital Data<br />
Launch Eval Report<br />
MANAGEMENT<br />
Page 5-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure 5.1.3-1b: Baseline Integration Schedule (Recurring Program)<br />
ID # Activity Resp<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
7<br />
8<br />
9<br />
10<br />
11<br />
12<br />
13<br />
14<br />
15<br />
16<br />
17<br />
18<br />
19<br />
20<br />
21<br />
22<br />
23<br />
24<br />
25<br />
26<br />
27<br />
28<br />
29<br />
30<br />
Kickoff<br />
ICD Sign-Off<br />
Critical Design Review (CDR)<br />
Fitcheck<br />
Shock Test<br />
LV Preshipment (Quality) Review<br />
SC Preshipment Readiness Review<br />
Operations Review<br />
Launch Site Acceptance Review<br />
Launch Vehicle Quality Review<br />
Update<br />
LV Roll-out Authorization Review<br />
Board<br />
LV Loading Authorization Review<br />
Board<br />
IRD (Amendment for specific SC)<br />
ICD prior to meeting to sign<br />
revision<br />
Inputs to Final Analysis (except<br />
CLA)<br />
Final SC Dynamic Model (Test<br />
Verified)<br />
Annex [X] of CDR package; CLA<br />
Report<br />
CDR Package; all except Annex [x]<br />
SC Environmental Test Plan<br />
Inputs to Fitcheck & Shock Test PlanSCC<br />
Fitcheck & Shock Test Plan<br />
Fitcheck & Shock Test Report<br />
SC Environmental Test Results<br />
Safety Sub<strong>mission</strong>s<br />
All Certificates<br />
SC Launch Operations Plan (final)<br />
Participants List w/ Passport Info<br />
Orbital Data<br />
Printed: 31AUG98<br />
Launch Evaluation Report<br />
Management Report<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
All<br />
SCC<br />
ILS<br />
SCC<br />
SCC<br />
ILS<br />
ILS<br />
SCC<br />
ILS<br />
All<br />
SCC<br />
SCC<br />
ILS<br />
SCC<br />
SCC<br />
ILS<br />
ILS<br />
ILS<br />
MEETINGS<br />
Kickoff<br />
ICD DEVELOPEMENT<br />
IRD<br />
CRITICAL DESIGN<br />
ICD<br />
Inputs to Final Analysis<br />
SPACECRAFT TESTING<br />
ICD Sign-Off<br />
SC Environmental Test Plan<br />
Inputs to Fitcheck & Shock Test Plan<br />
Fitcheck & Shock Test Plan<br />
SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSING<br />
LAUNCH CAMPAIGN<br />
MANAGEMENT<br />
STANDARD INTEGRATION SCHEDULE - RECURRING<br />
CDR<br />
Fitcheck<br />
Shock Test<br />
Final SC Dynamic Model<br />
CDR Package<br />
SC Launch Ops Plan (final)<br />
Fitcheck & Shock Report<br />
CLA Report<br />
LV Preship Rvw<br />
SC Preship Readiness Rvw<br />
Ops Review<br />
Site Review<br />
SC Environmental Test Results<br />
Safety Sub<strong>mission</strong>s<br />
All Certificates<br />
Participants & Info List<br />
Revised: 08/21/98<br />
L-12 L-10 L-8 L-6 L-4 L-2 LAUNCH L+2<br />
LVQR Update<br />
LV Roll-out Auth<br />
LV Loading Auth<br />
Orbital Data<br />
Launch Eval Report<br />
Page 5-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
5.1.4 Documentation Control and Delivery<br />
ILS maintains an internal documentation and configuration control system for all LSAs. Deliverable documentation<br />
shall be maintained under this configuration control system.<br />
All technical correspondence between ILS and the Customer relating to work on the LSA is strictly between the<br />
Customer Mission Manager and the ILS Mission Manager.<br />
5.1.5 Meetings and Reviews<br />
ILS and the Customer will meet as often as necessary to allow good and timely execution of all activities related to<br />
<strong>launch</strong> preparation of each satellite. A preliminary meeting schedule is defined in Section 5.1.5.1 and meeting<br />
schedules will be updated through the course of the contract as part of the interface activities milestone schedule<br />
generated by ILS. Exact dates, locations, agendas, and participation are agreed upon in advance, on a case-by-case<br />
basis, by the ILS Mission Manager and the Customer Mission Manager.<br />
5.1.5.1 Interface Meetings and Reviews<br />
The ILS Mission Manager chairs all meetings unless otherwise specified. ILS will provide meeting minutes at the end<br />
of each meeting signed by ILS and the Customer.<br />
A baseline meeting schedule is provided in Tables 5.1.5.1-1a and 5.1.5.1-1b for a non-recurring and recurring<br />
program respectively. A non-recurring program is one with a first of a kind SC which requires 2 analysis cycles. A<br />
recurring program is one with a similar SC which requires only one analysis cycle and no significant changes to the<br />
<strong>launch</strong> <strong>vehicle</strong> and the <strong>launch</strong> site.<br />
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A description of each type of meeting is provided below:<br />
Table 5.1.5.1-1a: Baseline Meeting Schedule for Non-Recurring Program<br />
Meeting Date* Location<br />
Kickoff L-24 TBD<br />
Preliminary ICD Review L-20 TBD<br />
Preliminary Design Review (PDR) L-14 Moscow<br />
Launch Site Visit No. 1 L-12 Launch Site<br />
Operations Review No. 1 L-12 TBD<br />
ICD Signoff L-11 TBD<br />
Fitcheck L-6 SC Manufacturer<br />
Shock Test L-6 SC Manufacturer<br />
Critical Design Review (CDR) L-6 Moscow<br />
Launch Site Visit No. 2 L-5 Launch Site<br />
Launch Vehicle Preshipment Review L-3 Moscow<br />
Spacecraft Preshipment Review L-2 SC Manufacturer<br />
Operations Review No. 2 L-2 TBD<br />
Launch Site Acceptance Review L-2 Launch Site<br />
Launch Vehicle Quality Review Update (TBC) L-7 days Launch Site<br />
Launch Vehicle Rollout Authorization Review Board L-6 days Launch Site<br />
Launch Vehicle Fueling Authorization Review Board L-1 day Launch Site<br />
* Date: Launch Minus X Months<br />
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Table 5.1.5.1-1b: Baseline Meeting Schedule for Recurring Program<br />
Meeting Date* Location<br />
Kickoff L-12 TBD<br />
ICD Signoff L-10 TBD<br />
Critical Design Review (CDR) L-6 Moscow<br />
Fitcheck L-6 SC Manufacturer<br />
Shock Test L-6 SC Manufacturer<br />
Launch Vehicle Preshipment (Quality)Review L-3 Moscow<br />
Spacecraft Preshipment Readiness Review L-2 SC Manufacturer<br />
Operations Review L-2 TBD<br />
Launch Site Acceptance Review L-2 Launch Site<br />
Launch Vehicle Quality Review Update (TBC) L-7 days Launch Site<br />
Launch Vehicle Rollout Authorization Review Board L-6 days Launch Site<br />
Launch Vehicle Fueling Authorization Review Board L-1 day Launch Site<br />
* Date: Launch Minus X Months<br />
Kickoff<br />
This meeting represents the formal start of the program. A description of overall LSA services will be presented as well<br />
as management organization and preliminary program schedules. The Interface Requirements Document will be<br />
discussed to kickoff generation of the ICD.<br />
Preliminary ICD Review<br />
The preliminary ICD will be reviewed and agreement reached on inputs to begin the Preliminary Analysis cycle.<br />
Preliminary Design Review (PDR)<br />
ILS will present all results of preliminary analyses and compare with ICD requirements.<br />
Launch Site Visit No. 1<br />
This first visit to the <strong>launch</strong> site will provide a first orientation to the Customer. A key goal is to verify compliance with<br />
ICD requirements.<br />
Operations Review<br />
Review of requirements and corresponding implemention for <strong>launch</strong> base operations.<br />
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Review of inputs to Final Analyses<br />
Agreement will be reached at this meeting to all final analysis inputs prior to starting these analyses.<br />
Fitcheck<br />
This is a fitcheck of flight adapter and separation system hardware to the flight spacecraft at the SC manufacturer's<br />
facility.<br />
Shock Test<br />
This is an actuation of the flight type separation system with the flight spacecraft at the manufacturer's facility. It is<br />
done in conjunction with the fitcheck.<br />
Critical Design Review (CDR)<br />
ILS presents all results from the final analysis cycle.<br />
Launch Site Visit No. 2<br />
This is the last visit prior to certification of the <strong>launch</strong> facility for this <strong>mission</strong>.<br />
Launch Vehicle Preshipment Review<br />
This meeting is held at Khrunichev as part of the quality control process. Khrunichev presents the quality status of all<br />
Launch Vehicle hardware per design documentation.<br />
Spacecraft Preshipment Review<br />
This meeting is held at the SC manufacturer's facility and provides a status of the spacecraft readiness to ship to the<br />
Launch Site.<br />
Launch Site Acceptance Review<br />
This meeting is held at the Launch Site prior to SC arrival to confirm the readiness of the Launch Site to begin the<br />
<strong>launch</strong> campaign. The compliance with requirements in the ICD and Operations Plan will be verified.<br />
Launch Vehicle Quality Review Update<br />
At this meeting, the data presented by Khrunichev at the Launch Vehicle Preshipment Review is updated to take into<br />
account all subsequent activities up to L-7 days before <strong>launch</strong> (TBC).<br />
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LV Rollout Authorization Review Board<br />
A meeting is held at the Launch Site to confirm readiness to rollout the LV to the Launch Pad.<br />
LV Fueling Authorization Review Board<br />
A meeting is held at the Launch Site to confirm readiness to load the LV with propellants and confirm SC readiness to<br />
<strong>launch</strong>.<br />
Action Item Control<br />
ILS maintains a centralized action item control system for each LSA.<br />
5.1.6 DTRA Oversight<br />
ILS arranges for Defense Threat Reduction Agency (DTRA) oversight, as necessary for technical interchange<br />
involving foreign nationals.<br />
5.1.7 Quarterly Report<br />
At the end of each contract quarter, ILS provides a report for each LSS covering management and technical progress<br />
including: subcontract status, technical status, <strong>mission</strong> integration schedule, production schedule and status for all<br />
major hardware, action item status, contract deliverable status. Major program issues are summarized and resolution<br />
plans discussed.<br />
5.1.8 Quality Provisions<br />
Refer to Appendix B for a description of Quality Assurance provision in place for Proton <strong>launch</strong> services.<br />
5.1.9 Launch License And Permits<br />
ILS will obtain all necessary Russian Federation permits and approvals required for the processing and <strong>launch</strong> of the<br />
Customer's spacecraft.<br />
The Customer will obtain permits and approvals required to import and export the spacecraft and associated<br />
equipment from its country of origin through the Port of Entry in Russia and Kazakhstan.<br />
5.2 ILS DELIVERABLES<br />
ILS provides the following deliverables during the course of each LSA. The baseline delivery schedule is provided in<br />
Table 5.2-1.<br />
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Table 5.2-1: ILS Deliverable Schedule for a Recurring and a Non-Recurring Program<br />
Recurring Non-Recurring<br />
DOCUMENT Date* Date*<br />
ICD DEVELOPMENT<br />
Preliminary ICD L-21<br />
ICD prior to sign off meeting L-11 L-12<br />
PRELIMINARY DESIGN<br />
PDR package PDR -10 days<br />
CRITICAL DESIGN<br />
CDR package CDR -10 days CDR -10 days<br />
SPACECRAFT TESTING<br />
Fitcheck and Shock Test Plan N/A L-9<br />
Fitcheck and Shock Test Report N/A L-5<br />
SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSING<br />
Safety Certificate L-3 L-3<br />
LAUNCH CAMPAIGN AND LAUNCH<br />
Orbital Data L+0 L+0<br />
Launch Evaluation Report L+2 L+2<br />
MANAGEMENT<br />
Management Report Each quarter Each quarter<br />
5.2.1 ICD Development<br />
5.2.1.1 ICD<br />
ILS shall provide the ICD and will maintain it up to date by issuing revisions as necessary. Preliminary ICD will<br />
contain input data for the Preliminary Design. The signed ICD will contain input data for the Critical Design.<br />
5.2.2 Preliminary and Critical Design<br />
ILS will conduct all performance and <strong>mission</strong> analyses required for the proper implementation of the Customer’s<br />
<strong>launch</strong> <strong>mission</strong>, including those analyses identified below.<br />
The following analyses are conducted during the <strong>mission</strong> integration effort for each satellite <strong>launch</strong> <strong>mission</strong>. For first<br />
of a kind spacecraft, one preliminary and one final analysis cycle will normally be conducted during each satellite<br />
integration effort. For follow-on spacecraft, one analysis cycle will be performed in most cases. Two analysis cycles<br />
will be performed for following on SC where spacecraft and/or <strong>launch</strong> <strong>vehicle</strong> relevant parameters have changed<br />
significantly.<br />
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Each cycle includes the analyses defined in Table 5.2.2-1:<br />
Table 5.2.2-1: Design Review Analyses<br />
Section Title Description<br />
Design and Manufacturing A summary of the LV design concentrating on differences with previous<br />
<strong>vehicle</strong>s. Emphasis is on specificities in adapter and payload compartment<br />
design to meet specific spacecraft payload requirements.<br />
Mission Design The flight design including maneuvers and maneuver sequence, orbit<br />
parameters and dispersions, collision avoidance<br />
Thermal Analysis Integrated thermal analysis of combined operations (ground and flight) for<br />
spacecraft and <strong>launch</strong> <strong>vehicle</strong> hardware to ensure thermal compatibility.<br />
The spacecraft mathematical model is provided by the Customer per the<br />
thermal model specification provided by ILS<br />
Separation Analysis Analysis of spacecraft separation including presentation of pertinent<br />
kinematic parameters and their dispersions during the separation event<br />
CLA/Acoustic/ Shock Loads<br />
Environment<br />
1) Dynamic coupled-loads analysis. The spacecraft mathematical model is<br />
furnished by the Customer according to the ILS provided dynamic model<br />
specification. The following events are analyzed:<br />
a) Liftoff<br />
b) Flight Winds and Gust<br />
c) First/Second Stage Separation<br />
For follow-on satellites of the same configuration, only one verification<br />
coupled loads analysis will be conducted unless significant <strong>launch</strong> <strong>vehicle</strong><br />
configuration changes have occurred (e.g. if first <strong>launch</strong> occurs on Block<br />
DM and next <strong>launch</strong> occurs with Breeze M).<br />
2) Presentation of other load environments including acoustic, shock and<br />
ground transportation loads<br />
Contamination Analysis of ground and flight contamination sources and effect on<br />
spacecraft payload<br />
RF Link and EMC Analysis of the RF link between the bunker and the pad and EMC analysis<br />
verifying compatibility between the spacecraft and LV systems<br />
Clearance Analysis Clearance analysis between the spacecraft and the LV during flight to verify<br />
sufficient dynamic clearances<br />
Venting Analysis Analysis of fairing depressurization during flight<br />
Operations Detailed description of how Khrunichev will meet operational<br />
requirements specified by the Customer in the ICD<br />
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One hardcopy (and a soft copy where possible) of all design documentation will be provided to Customer 10 days prior<br />
to review<br />
Reports will be provided documenting the results of the above analyses. These reports will be provided for each analysis<br />
cycle and include the following topics: summary of results, detail of analyses performed, comparison of analysis results<br />
with ICD requirements. The analyses required may be reduced in scope if agreed between ILS and the Customer.<br />
5.2.3 Spacecraft Testing - Fitcheck/ Shock Test Plan, and Report<br />
ILS will provide an overall plan describing the Fitcheck/Shocktest and a description of the responsibilities and actions<br />
for each of the participants including Khrunichev, ILS, the SC manufacturer and SAAB (if applicable). ILS will<br />
provide a summary report following the Fitcheck & Shocktest documenting the results.<br />
5.2.4 Safety<br />
ILS will prepare a Safety Data Package based on the safety data provided by the Customer Safety Sub<strong>mission</strong>s.<br />
5.2.5 Launch Campaign and Launch<br />
5.2.5.1 Orbital Data<br />
ILS will provide the State Vector data as described in Chapter 2.<br />
5.2.5.2 Launch Evaluation Report<br />
ILS will provide a Launch Evaluation Report for each LSA documenting the results of ground processing of the<br />
spacecraft and the subsequent flight.<br />
5.2.6 Management and Reports<br />
For each LSA, ILS will provide a report each quarter documenting the status of management issues.<br />
5.3 CUSTOMER DELIVERABLES<br />
The Customer provides the following deliverables during the course of each LSA. The baseline delivery schedule is<br />
provided in Tables 5.3-1.<br />
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Table 5.3-1: Customer Deliverable Schedule for a Recurring and a Non-Recurring Program<br />
Recurring Non-Recurring<br />
DOCUMENT Date* Date*<br />
ICD DEVELOPMENT<br />
IRD L-12 L-24<br />
PRELIMINARY DESIGN<br />
Preliminary SC Dynamic Model L-24<br />
SC Thermal Model L-19<br />
CRITICAL DESIGN<br />
Final SC Dynamic Model (test verified model) L-7 L-7<br />
Inputs to Final Analyses (except SC Dynamic Model) L-12 L-11<br />
SPACECRAFT TESTING<br />
SC Environmental Test Plan L-11 L-11<br />
Notching Profile for Finite Element Analysis Sine-test -2 months<br />
SC Environmental Test Results L-4 L-4<br />
Inputs to Fitcheck & Shock Test Plan L-10 L-10<br />
Fitcheck & Shock Test Report L-6 L-5<br />
SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSING<br />
Safety Sub<strong>mission</strong>s (Preliminary and Final SMPSP) L-12 to L-5 L-24 and L-5<br />
Pre-Launch safety certificates L-3 L-3<br />
LAUNCH CAMPAIGN AND LAUNCH<br />
Spacecraft Launch Operations Plan (preliminary) L-16<br />
Spacecraft Launch Operations Plan (final) L-8 L-8<br />
Listing of Campaign Participants with Passport Information L-3 L-3<br />
SC Orbital data L+2 days L+2 days<br />
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5.3.1 ICD Development<br />
The Customer will provide an IRD to ILS with interface requirements describing all pertinent design information on<br />
the spacecraft characteristics, mechanical and electrical interfaces, and constraints necessary to define the integration<br />
tasks and <strong>mission</strong> operation. This will be used to generate the preliminary ICD.<br />
5.3.2 Preliminary and Critical Design<br />
5.3.2.1 SC Dynamic Model<br />
The Customer will provide a SC dynamic model conforming to the ILS dynamic model specification. For the second<br />
analysis cycle, this model will be a test verified version.<br />
5.3.2.2 SC Thermal Model<br />
The Customer will provide a SC thermal model conforming to the ILS thermal model specification.<br />
5.3.2.3 SC Fluid Slosh Model<br />
The Customer shall provide a SC fluid slosh model conforming to the requirements in Appendix C.<br />
5.3.3 Spacecraft Testing<br />
5.3.3.1 SC Environmental Test Plan and Results<br />
The Customer will provide a test plan for ILS approval documenting the tests which will be performed by the SC<br />
manufacturer to demonstrate compatibility with the Proton ground and flight environments. A summary of the results<br />
from these tests will be provided at test completion.<br />
5.3.3.2 Fitcheck/ Shock Test Plan, Procedures and Report<br />
The Customer will provide input to the ILS Fitcheck/Shock Test Plan. The Customer will provide a summary report<br />
following the Fitcheck & Shocktest documenting the results.<br />
5.3.4 Required Safety Data and Certificates<br />
5.3.4.1 Safety Sub<strong>mission</strong>s<br />
The Customer will provide all necessary data required by the Safety Plan to demonstrate that the spacecraft systems,<br />
GSE and procedures are compatible with the Launch Site and flight safety requirements.<br />
5.3.4.2 SC Reliability Certificate<br />
The Customer will provide a certificate to confirm reliability and specifying the SC no-failure probability level for<br />
ground operations and flight.<br />
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5.3.4.3 SC Readiness Certificate<br />
The Customer will provide a certificate to ILS following spacecraft environmental testing to certify that the spacecraft<br />
is capable of withstanding the loads environments imposed on it during ground processing at the Launch Site and<br />
during flight as documented in the ICD and in the CLA report.<br />
5.3.4.4 Certification of Orbital Position Approval<br />
The Customer will provide to ILS a certificate demonstrating that approval has been given by an internationally<br />
recognized authority for the orbital position which the spacecraft to be <strong>launch</strong>ed under this LSA is to occupy.<br />
5.3.4.5 Certificate Approving SC for Space Related Activities (TBC)<br />
The customer will provide a certificate stating that the SC customer is authorized to conduct space related activities.<br />
5.3.4.6 Certificate of Intent to File National Registry(TBC)<br />
The Customer will provide a certificate stating intent to enter SC into the national registry of the owner's country.<br />
5.3.5 Launch Campaign and Launch<br />
5.3.5.1 Spacecraft Launch Operations Plan<br />
The Customer will provide a plan which describes the spacecraft <strong>launch</strong> operations at the Launch Site.<br />
5.3.5.2 Listing of Campaign Participants<br />
The Customer will provide a list of all potential campaign participants 3 months prior to <strong>launch</strong> with all required<br />
passport information. This list will designate primary and backup personnel.<br />
5.3.5.3 Orbital Data<br />
The Customer will provide SC state vector data complying with requirements in Chapter 2 of this Mission Planner’s<br />
Guide.<br />
5.4 SPECIFIC CUSTOMER RESPONSIBILITIES<br />
For each LSA, the Customer has the following responsibilities.<br />
5.4.1 Campaign Duration<br />
The campaign duration from spacecraft arrival to <strong>launch</strong> shall not exceed 42 days.<br />
5.4.2 Spacecraft And Associated Ground Equipment<br />
The Customer will provide at the Launch Site the spacecraft and associated ground equipment and personnel required<br />
to meet the contracted <strong>launch</strong> date.<br />
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5.4.3 Final Spacecraft Data<br />
For <strong>mission</strong>s that employ the Block DM, the Customer shall also provide final spacecraft estimated dry and wet mass<br />
two days prior to Block DM fueling. Prior to encapsulation, the Customer will supply ILS the actual satellite dry and<br />
wet masses.<br />
5.4.4 Spacecraft Readiness<br />
The Customer will provide a readiness to proceed with operations on <strong>launch</strong> minus 17 days (L-17 day), prior to the<br />
start of Combined Operations; prior to rollout to the pad; and prior to fueling of the Launch Vehicle. These dates will<br />
be coordinated with the Baikonur operations schedule.<br />
5.4.5 Removal of SC Support Equipment<br />
Unless prior arrangements have been made, the Customer shall remove from the Baikonur Cosmodrome all of its<br />
ground support equipment using Customer provided charter aircraft within 3 days of <strong>launch</strong>.<br />
5.4.6 Evaluation Of Launch Vehicle And Associated Services<br />
The Customer will provide to ILS, as soon as practical after <strong>launch</strong>, with all relevant available data from the <strong>launch</strong><br />
necessary to assist ILS in evaluating the performance of the <strong>launch</strong> <strong>vehicle</strong> and associated services provided under each<br />
LSA.<br />
5.4.7 Spacecraft Propellants<br />
The Customer will procure spacecraft propellants to support the <strong>launch</strong> campaign and is responsible for shipment of<br />
these propellants to the Port of Entry into Russia (typically St. Petersberg) and through Customs. After the <strong>launch</strong><br />
campaign, the Customer will be responsible for removal of the propellants and associated equipment from the Port of<br />
Entry. ILS will assist the Customer with Customs clearance procedures.<br />
The Customer is responsible for propellant transportation costs.<br />
5.4.8 Connectors<br />
The Customer shall provide to ILS, flight and test connectors per <strong>mission</strong> specific requirements. These connectors will<br />
be used for the assembly of flight and test harnesses.<br />
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5.5 ILS SERVICES AND MATERIAL SPECIFICALLY EXCLUDED<br />
ILS has no obligation to provide the following goods or services:<br />
a) Receiving inspection of spacecraft elements and support equipment upon arrival at the <strong>launch</strong> site<br />
b) Analysis of data generated by the spacecraft through its own telemetry system<br />
c) Any spacecraft test equipment<br />
d) Shipping cost associated with the spacecraft, its components, and support equipment (except while at the <strong>launch</strong><br />
site)<br />
e) Replacement parts for the spacecraft or its support equipment<br />
f) Installation, handling, or other responsibility related to spacecraft pyrotechnic systems or elements<br />
g) Functional operation or installation of any spacecraft systems<br />
h) Any tracking or commanding of the spacecraft after separation from the <strong>launch</strong> <strong>vehicle</strong><br />
i) Spacecraft propellants<br />
j) Propellant sampling analysis. Facilities at or near the Launch Site are not equipped with equipment or technology<br />
necessary for analysis of SC propellants. The Customer must plan for shipment and analysis of samples outside the<br />
Russian Federation if these analyses are required<br />
k) Storage of spacecraft and all associated Customer GSE over 2 months<br />
l) Storage of propellants over 4 months<br />
m) Additional analyses over and above those specified in previous sections, caused by changes to the spacecraft design<br />
which are not in any way attributable to ILS and not required by the terms of this LSA and its' Exhibits<br />
n) Additional analyses over and above those required in previous sections, caused by <strong>launch</strong> postponements requested<br />
by the Customer (unless otherwise specified in the postponement provisions of the LSA)TBC<br />
o) Changes to the LV and/or the <strong>launch</strong> site facilities as described in this Mission Planner’s Guide , caused by<br />
changes to the spacecraft design which are not in any way attributable to ILS and not required by the terms of the<br />
LSA and its' Exhibits<br />
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6. SPACECRAFT AND LAUNCH FACILITIES<br />
6.1 FACILITIES OVERVIEW<br />
The Baikonur Cosmodrome includes several facilities that are used for ILS <strong>launch</strong> campaigns (Figure 6.1-1). These<br />
facilities include:<br />
a) Yubeleini Airfield—Spacecraft (SC) and ground support equipment (GSE) arrival and departure<br />
b) Area 31 and Area 92A-50—SC preparation and encapsulation<br />
c) Building 92-1—Final integration of the Ascent Unit to Stages 1-3 of the Launch Vehicle Proton K<br />
d) Building 92A-50—Final integration of the Ascent Unit to Stages 1-3 of the Launch Vehicle Proton M<br />
e) Launch Complex Area 81, Pad 23 or 24—SC <strong>launch</strong><br />
The sub-sections that follow provide brief descriptions of these facilities. More in-depth descriptions of these same<br />
facilities are provided later in this section.<br />
6.1.1 Yubeleini Airport<br />
Yubeleini Airport is located at the Baikonur Cosmodrome and is used for receiving the charter aircraft carrying the SC<br />
and GSE, as well as charter flights with campaign personnel. It is an internationally rated airport with a single 4.5 km<br />
long, 84 m wide landing strip oriented 60 degrees/240 degrees relative to North. The airport has an elevation of<br />
approximately 100 m above sea level.<br />
A 140 m by 420 m pad is available next to a railhead for unloading aircraft and transferring equipment to rail convoys.<br />
Prior to aircraft arrival, this area is cleared and ground handling equipment is positioned. The pad also is equipped<br />
with stationary spotlights for use in night operations.<br />
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Figure 6.1-1: Baikonur Facilities Map<br />
92A-50<br />
Facility<br />
Bldg 92-1<br />
Railline:<br />
Road:<br />
Hotel<br />
Complex<br />
Proton Launch Complex 81<br />
(Pads 23 & 24)<br />
Proton Launch<br />
Complex 200<br />
(Pads 39 & 40)<br />
Yubileini Airfield<br />
Area 254<br />
Area 31<br />
Propellant Storage Facility Oxygen and Nitrogen Plant<br />
Tyura-Tam Rail Station<br />
Krainy Airfield<br />
Syr dar'ya River<br />
Hwy To Tashkent<br />
BAIKONUR TOWN<br />
(LENINSK)<br />
Saturn Measurement Station<br />
approximately 90 km- figure not to scale<br />
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6.1.2 Building 92A-50<br />
Building 92A-50 contains all facilities necessary for processing a SC from its arrival through its integration with the<br />
Fourth Stage and encapsulation. The SC and GSE arrive at Hall 102, where the containers are cleaned. The SC<br />
container is transported on the railcar or on its own wheels into Hall 101, where the SC is normally removed from the<br />
container. The SC is then transported into Hall 103A on the SC transporter where it is installed onto its fueling/test<br />
stand. The SC remains in this hall for all subsequent testing and fueling operations. Following fueling, the SC is<br />
transported back to Hall 101 on a special transport dolly, where it is integrated with the Fourth Stage and<br />
encapsulated.<br />
6.1.3 Area 31 Facilities<br />
Area 31 includes three facilities that serve important functions in preparing spacecraft for <strong>launch</strong>.<br />
6.1.3.1 Building 40, Hall 100; Area A, B and C<br />
Building 40, Hall 100 is a large high-bay that is divided into three main zones: Area A, Area B and Area C:<br />
a) Area A is a Class 100,000 clean area used for final preparation of the Ascent Unit and its components. It is also<br />
used for integrating the SC with the flight adapter and mating the SC to the Fourth Stage. Final encapsulation of<br />
the SC also takes place in this area.<br />
b) Area B is a Class 100,000 airlock used to transition material from Area C into Area A and into Rooms 119, 120,<br />
and 121 of Building 40D.<br />
c) Area C encompasses approximately half the building and is used for SC and GSE loading/unloading, as well as<br />
the Fourth Stage and Ascent Unit components.<br />
6.1.3.2 Building 40D, Rooms 119, 120, 121, and Offices<br />
Building 40D is a three-story facility adjoining Building 40 that includes satellite preparation, ground equipment, and<br />
personnel/office areas:<br />
a) Room 119 is a Class 100,000 cleanroom used for SC testing prior to loading. SC pressurization may occur in this<br />
room prior to transfer to the Filling Hall (Building 44) with a portable blast shield in place.<br />
b) Room 120 is a Class 100,000 cleanroom used as a control room for test operations conducted in Room 119. Most<br />
electrical support equipment is located in this room.<br />
c) Room 121 is a Class 100,000 cleanroom used for SC equipment storage, as required.<br />
d) Offices are provided on the second and third floors for the SC Customer and ILS personnel.<br />
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6.1.3.3 Building 44 (Fueling Hall)<br />
Building 44 is a propellant loading facility that is used for loading the SC as well as the Fourth Stage. It is divided into<br />
three main halls.<br />
a) Room 1: Propellant Loading Area. A Class 100,000 clean tent is installed in this room to receive the SC for fuel<br />
and oxidizer loading operations. Oxidizer propellant cylinders are thermally conditioned in this room in special<br />
chambers prior to loading.<br />
b) Room 2: SC Removal from Container and Container Storage Operations. This room is used for agitation of<br />
propellant cylinders, storage of the Thermal Transport Container, and storage of the GSE.<br />
c) Room 3: Storage of Propellant Cylinders. Storage of fuel propellant cylinders occurs in this room in prior to<br />
loading.<br />
In addition to the areas described above, a control room on the second floor (Room 58) houses SC Customer electrical<br />
ground support equipment that is used to monitor spacecraft telemetry, charge SC batteries, and communicate with<br />
the propellant-load team.<br />
Fuel and oxidizer decontamination rooms are available for cleaning the GSE. These are located at either end of the<br />
building and are supplied with water and nitrogen gas sources.<br />
6.1.4 Building 92-1 and the Proton Launch Zone (Area 81)<br />
Following encapsulation, the Ascent Unit is transported to Building 92-1 for integration with Stages 1-3 of the<br />
Launch Vehicle. For the Breeze M configuration, the Launch Vehicle will return to Building 92A-50, Hall 111 for<br />
final electrical verification. Following integration, the Launch Vehicle is transported to the Proton Launch Complex<br />
(Area 81) for erection and <strong>launch</strong>. At the Launch Complex, two areas are used for Spacecraft GSE. Rooms 64 and 76<br />
(underground vaults) accommodate SC customer equipment providing power to the SC while on the pad. In the<br />
Bunker, which is approximately 1 km from the <strong>launch</strong> pad, Room 250 is used for installation of SC customer electrical<br />
GSE required for <strong>launch</strong>.<br />
6.1.5 Hotels<br />
The Hotel Kometa, Hotel Fili, and Hotel Polyot, which are located in Area 95 near the Launch Complex, are used to<br />
house personnel during a <strong>launch</strong> campaign.<br />
6.2 SPACECRAFT PROCESSING FACILITIES<br />
This section describes the spacecraft (SC) processing facilities, which provide the capability to perform all required<br />
operations from receipt of the SC through its encapsulation in preparation for <strong>launch</strong> on the Proton Launch Vehicle<br />
(LV) at the Baikonur Cosmodrome. These operations include off-loading in the SC technical zone, testing, fueling,<br />
mating to the Fourth Stage, payload encapsulation, and LV integration.<br />
The SC processing facilities and Proton <strong>launch</strong> complex are located at the Baikonur Cosmodrome in the Republic of<br />
Kazakhstan in Central Asia, approximately 2,000 km southeast of Moscow. The annual temperature averages 13 o C,<br />
ranging from -40 o C in winter to 45 o C in summer. Figure 1.3-1 depicts the overall layout of the Cosmodrome, showing<br />
the facilities that are used for ILS <strong>launch</strong> campaigns. The specific SC processing areas at Cosmodrome described<br />
include:<br />
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a) Area 92, Building 92A-50 (Section 6.2.1)<br />
b) Area 31, Buildings 40/40D (Section 6.2.2)<br />
c) Area 31, Building 44 (Section 6.2.3)<br />
d) Area 254, Building 254-1 (Section 6.2.4)<br />
e) Area 92, Building 92-1 (Section 6.2.5)<br />
The Baikonur Cosmodrome is equipped with spur-railroad service lines that are used for most transport. Specialized<br />
equipment is available for fueling, handling of compressed gases, and SC integration with the LV.<br />
The main buildings within the technical complex are the integration and testing facilities. Assembly and integration of<br />
the various stages of the Proton LV are carried out in the Launch Vehicle Integration Building. Spacecraft preparation,<br />
testing, and integration with the <strong>launch</strong> <strong>vehicle</strong> fourth stage and the fairing are accomplished either in Area 92,<br />
(Building 92A-50) or in Area 31, (Buildings 40/40D), which are both Spacecraft Processing Areas. Spacecraft fueling<br />
and pneumatics pressurization are accomplished in either Area 92, (Building 92A-50) or Area 31, (Building 44) the<br />
Spacecraft Fueling Hall). The LV fourth stage with the mated spacecraft and fairing are transferred to the LV<br />
Technical Zone in Area 92, (Building 92-1), where they are horizontally mated to the assembled LV Stages 1-3. The<br />
integrated LV is then transported to the Proton Launch Zone (Area 81) for erection, checkout, and <strong>launch</strong> (see Section<br />
6.3: Launch Complex Facilities).<br />
6.2.1 Facility 92A-50<br />
Facility 92A-50 is located in Area 92 of the Baikonur Cosmodrome. The facility was modified and outfitted for the<br />
specific needs of SC Customers, and provides the capability to perform required operations in one conveniently located<br />
area. These operations include SC offloading, testing, fueling, mating to the Fourth Stage, and payload encapsulation.<br />
This section describes the specific functional areas included within Facility 92A-50, as well as the equipment and<br />
services available to SC Customers for pre-<strong>launch</strong> payload processing.<br />
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6.2.1.1 Facility Layout and Area Designations<br />
Building 92A-50 has been expressly modified and outfitted to efficiently complete all SC processing and encapsulation<br />
in a single building. The halls/rooms, facility systems, and equipment are sized to accommodate SCs of up to<br />
approximately 4.5 m diameter, 10.0 m height, and loaded weights of up to 8,700 kilograms. Figure 6.2.1.1-1 depicts<br />
the overall arrangement of all areas within the building used for commercial programs.<br />
A SC, in the manufacturer’s shipping container, is delivered into Hall 101 on railcar, after having been cleaned in the<br />
Receiving Area. In Hall 101, the container is removed from the railcar and placed on the floor. After the railcar is<br />
removed and the environment reestablished, the SC is removed from the shipping container and placed on a<br />
transporter to be moved into the Processing and Fueling Hall (Hall 103A). Once there the SC is placed on the fueling<br />
island and require no further movement in order to complete all necessary standalone assembly, checkout, propellant<br />
loading, and pneumatics servicing. When ready, the SC is moved by special transport dolly to the Integration Hall<br />
(Hall 101) for mating to the Proton upper stage and encapsulation inside the nose fairing.<br />
Building 92A-50 is approximately 229 m long and 147 m wide. Only a portion of the building is used for commercial<br />
programs.<br />
The Receiving Area (Hall 102) is the primary entrance for the SC and associated equipment, and is located on the east<br />
side of the building.<br />
The main entry into 92A-50 for ILS and SC processing personnel is next to Hall 103A, on the west end of the<br />
building, near the Change Room Area. An additional entrance on the north side of Hall 103A is used for Control<br />
Room equipment delivery.<br />
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Figure 6.2.1.1-1: Building 92A-50 General Arrangement<br />
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6.2.1.2 Receiving/Storage Area - Hall 102<br />
The Receiving Area (Hall 102), or the Integration Area (Hall 101), may be used to offload the SC Container from its<br />
transport railcar. External access for the SC and its GSE is provided by rail through two locally controlled, exterior<br />
sliding doors located in the hall’s east wall. An area is provided outside of Hall 102 for initial wash-down of the railcar<br />
and SC container prior to its entry into the building.<br />
An overhead crane is provided in Hall 102 to transfer SC GSE from their railcars to the floor for storage and<br />
offloading.<br />
Container cleaning is accomplished in Hall 102. An area of approximately 240 square meters is provided for non-Class<br />
100,000 storage of non-hazardous items.<br />
The overall clear dimensions of Hall 102 are approximately 70.5 m by 36 m. The clear ceiling height is 25.85 m, and<br />
the height of the overhead 50 T crane hook is 18.01 m. The SC unloading area is approximately 8.85 m wide and 34.1<br />
m long.<br />
When ready, the SC Container on the railcar is moved, via railcar, from Hall 102 to the inside of Hall 101 where it is<br />
transferred to an air pallet for the move into the Processing and Fueling Area (Hall 103A). Alternatively, the SC may<br />
be moved using its own wheeled container, a transport dolly.<br />
6.2.1.3 Integration Area - Hall 101<br />
Once the SC has been processed and fueled in Hall 103A, it is transported to the Integration Area (Hall 101), which is<br />
a Class 100,000 cleanroom. The Integration Hall is used to assemble the Ascent Unit, which involves the following<br />
operations:<br />
a) mating the SC, Adapter, and Fourth Stage<br />
b) checking system continuity<br />
c) rollover of the fourth stage to horizontal<br />
d) encapsulation within the Nose Fairing<br />
Overhead bridge cranes are used to transfer the SC from the Transport Dolly to the rollover fixture, as well as<br />
transferring the integrated Ascent Unit from the rollover fixture to a railcar for delivery to the Integration Facility<br />
92-1.<br />
This Hall is also used for electrical checkouts of the Breeze M stage prior to fueling.<br />
Hall 101 is 34.5 m wide and 107 m long. It has a full-height wall and ceiling facing as well as door sealing, thermal<br />
insulation, and an anti-static floor coating.<br />
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6.2.1.4 Spacecraft Processing and Fueling Hall - Hall 103A (Room 4101)<br />
Hall 103A (also referred to as Room 4101), the Processing and Fueling Hall, is used for pre-encapsulation SC<br />
processing, including loading propellants and servicing pneumatics in the SC (Figure 6.2.1.4-1). Access to Hall 103A<br />
is made available from Hall 103 through two sliding doors with a clear opening that is 9.5 m wide by 11.95 m high. A<br />
15 MT overhead bridge crane, is provided.<br />
An 8 m by 8 m fueling island, located on the west side of Hall 103A, is used for oxidizer and fuel transfer operations. It<br />
is surrounded by a grating-covered trench, which drains any fuel or oxidizer spills into separate waste tanks. The<br />
grating permits the passage of wheeled dollies.<br />
The floor of Hall 103A has an anti-static coating and a load rating of 10 MT (3,000 kg/cm 2 ) per truck axle. All<br />
finishes in Hall 103A use materials that do not react with propellants.<br />
The wall between Hall 103A and Hall 103 includes a pair of large doors designed to withstand a 60 kilogram per<br />
square meter overpressure load.<br />
Rooms 4114 and 4116 provide rapid egress routes from Hall 103A, and the Pressurization Airlock (Room 4110)<br />
provides the standard egress route. Rooms 4114 and 4116 each have three emergency showers and eyewashes. The<br />
SCAPE Shower Areas (Rooms 4121 and 4122) have showers for post-operation clean-up. A parking lot for<br />
ambulances and fire trucks is located next to the rapid egress routes. The Pressurization Airlock (Room 4110) and the<br />
space between the double doors between Hall 103 and 103A are pressurized with clean air in order to isolate Hall 103A<br />
during propellant loading operations.<br />
The clear dimensions for the SC Processing and Fueling Hall 103A are 16.5 m wide by 22 m long.<br />
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Figure 6.2.1.4-1: Building 92A-50 Spacecraft Processing and Fueling Area<br />
.<br />
BUILDING 92A-50<br />
SPACECRAFT PROCESSING &<br />
FUELING AREA<br />
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6.2.1.5 Fuel and Oxidizer Conditioning Rooms - Rooms 4112 & 4105<br />
Room 4112, the Fuel Conditioning Room, is used for temporary storage of the campaign fuel (e.g., MMH, N 2H 4)<br />
and thermal conditioning of the fuel before loading. Room 4105, the Oxidizer Conditioning Room, is used for<br />
temporary storage of the campaign oxidizer (e.g.N2O4) and thermal conditioning of the oxidizer before loading (Plate<br />
3.1-6). Both rooms have the capability to collect and dispose of propellant spills. Room 4112 contains no materials<br />
that react with fuel, and Room 4105 contains no materials that react with oxidizer.<br />
The floors in both rooms have an anti-static coating and a load rating of 10 MT(3,000 kg/cm 2 ) per truck axle. The<br />
floor elevations are the same as Room 4101.<br />
Room 4105 and 4112 are approximately 5.7 m long and 4.4 m wide, and both have clear ceiling heights of 2.9 m.<br />
6.2.1.6 Fuel and Oxidizer Equipment Decontamination Rooms - Rooms 4111 & 4115<br />
Room 4111, the Fuel Equipment Decontamination Room, is used to decontaminate fuel loading. Room 4115, the<br />
Oxidizer Equipment Decontamination Room, is used to decontaminate oxidizer loading equipment. Both rooms have<br />
the capability to collect and dispose of propellant spills. Room 4111 contains no materials that react with fuel, and<br />
Room 4115 incorporates no materials that react with oxidizer.<br />
Rooms 4111 and 4115 are both 6.1 m long and 4.1 m wide, and both have clear ceiling heights of 2.95 m.<br />
6.2.1.7 Control Room - Room 4102<br />
Room 4102, the Control Room, is used for monitoring and controlling SC processing and fueling activities in Hall<br />
103A, as well as SC integration into the Ascent Unit in Hall 101.<br />
A blast-resistant (bulletproof) viewing window is provided between the Control Room and Hall 103A for monitoring<br />
all processing and fueling operations. The wall between 103A and the Control Room is a welded, reinforced steel<br />
structure that provides a hermetic seal.<br />
The Control Room is 4.9 m by 12.9 m in overall dimension, with a clear ceiling height of 3. 1 m. An equipment entry<br />
vestibule, with inner and outer doors 2.9 m wide and 2.9 m high, is provided to facilitate equipment movement into<br />
the Control Room.<br />
The floors of the Control Room and all associated access corridors are designed for wheeled dollies. Forklifts may be<br />
used to bring equipment into the vestibule of the room, but they are not permitted to operate in the Control Room<br />
itself. Temporary ramps are available to aide moving items from the entrance vestibule into the Control Room.<br />
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6.2.1.8 Entrance/Lobby Area<br />
The Entrance/Lobby Area includes the following rooms and features:<br />
a) a street-level entrance<br />
b) lobby area (Room 0301)<br />
c) cloak room (Room 0302)<br />
d) SC Customer security checkpoint with viewing windows and security sensor alarm panel (Room 0318)<br />
e) tool storage room (Room 0319)<br />
f) restroom (Room 0320)<br />
Because these rooms serve general purpose administrative functions, they have an office-type environment.<br />
6.2.1.9 Change Room Area<br />
The Change Room Area consists of several rooms (0303 – 0317), including independent men’s and women’s<br />
restrooms, change and shower areas, a storage and issue room for cleanroom garments, a Personal Protective<br />
Equipment storage and donning room, and a corridor with an air shower. Clean passage is available from the Change<br />
Rooms to either Hall 103 or Hall 103A.<br />
6.2.1.10 Pressurization Airlock - Room 4110<br />
The Pressurization Airlock provides clean access between the Airshower (Room 0317) and Hall 103A. During<br />
propellant loading operations in Hall 103A, the Airlock is pressurized slightly more that Hall 103A to prevent vapor<br />
migration from Hall 103A. SCAPE suited personnel use the airlock to access the SCAPE Showers and Doffing<br />
Rooms, and the Pressurization Airlock and Corridor to the Airshower and Change Rooms can be used as an<br />
emergency egress route from Hall 103A, if necessary.<br />
Room 4110 is 1.4 m by 3.5 m wall-to wall.<br />
6.2.1.11 SCAPE Doffing Rooms and Showers - Rooms 4108, 4109, 4121 and 4122<br />
The SCAPE Doffing Rooms; Rooms 4108 for Fuel and 4109 for Oxidizer, are available for donning and doffing<br />
Personal Protective Equipment (PPE) for a propellant loading operation. As necessary, SCAPE Showers, Rooms 4121<br />
for Fuel and 4122 for Oxidizer, are available to decontaminate the PPE suits before doffing. The dedicated showers<br />
are plumbed to the respective liquid waste tanks.<br />
Room 4108 is approximately 1.65 m by 3.0 m and Room 4109 is approximately 1.9 m by 5.4 m. Room 4121 is 1.2 m<br />
by 3.4 m and Room 4122 is 1.2 m by 3.4 m.<br />
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6.2.1.12 Clean Storage Hall - Hall 103<br />
Hall 103, the Clean Storage Hall, provides accessible storage for clean items supporting SC processing. It also provides<br />
a Class 100,000 corridor between Halls 101 and 103A.<br />
A refrigerator/freezer is available for consumable storage. The unit has an internal volume of approximately 1 cubic<br />
meter, and its operating temperature is adjustable from -18 to + 10°C.<br />
The wall-to-wall dimensions of the Clean Storage Hall (Hall 103) are 17.5 m by 31.8 m at floor level, and the ceiling<br />
height is 15.0 m. At heights greater than 3 m above the floor, the width of Hall 103 restricted by HVAC ducting to<br />
about 16 m.<br />
6.2.1.13 Ordnance Storage<br />
Khrunichev provides limited storage of ordnance required to support a <strong>launch</strong> campaign. The Ordnance Storage Room<br />
may be accessed through a door located in the north wall of Hall 101.<br />
Ordnance to be stored must meet the following criteria:<br />
a) A maximum (TNT equivalent) quantity of 50 grams requiring a volume no more than 60 cm by 60 cm by 60 cm<br />
may be stored in accordance with Russian Federation Standards<br />
b) Only insensitive explosives are permitted, and each item must be individually packaged in U.S. Department of<br />
Transportation-approved shipping and storage containers<br />
c) The SC Customer must provide a certificate of conformance to the Hazard of Electromagnetic Radiation to<br />
Ordnance (HERO) Specification (MIL-I-23659)<br />
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6.2.1.14 Offices and Conference Room Area - Rooms 1202 through 1209<br />
An Office/Conference Room Area (Rooms 1202 - 1209) is located on the second floor of Building 92A-50 (Figure<br />
6.2.1.14). The functions of the seven constituent rooms are:<br />
a) 1202, Facility Management/Security Control and Medical Office<br />
b) 1203, SC Mission Management Office<br />
c) 1204, Break Room<br />
d) 1205, ILS Office<br />
e) 1206, SC Contractors Office<br />
f) 1207, SC Customer’s Office<br />
g) 1208/1209, Conference Room<br />
The clear dimensions of these rooms are as follows:<br />
a) Offices - Rooms 1202 through 1207 are 8.9 m by 5.9 m<br />
b) Conference Room 1208/1209 is 8.9 m by 11.9 m<br />
c) The clear height of all rooms is 3.1 m<br />
Restrooms are accessible from the corridor serving the Office/Conference Room Area; general access to the area is via<br />
stairs from the “street” entrance to the Change Room Area. Only essential personnel are permitted in the<br />
Office/Conference Room Area during propellant transfer operations.<br />
Two egress routes are available from the area: the normal route at the western end of the room block that exits to the<br />
“street” entrance to the Change Room Area; and an emergency evacuation route that exits east through the<br />
Khrunichev area of the building. A 2,000 kg capacity freight elevator, with a 1.8 m wide by 2.2 m high door opening<br />
and floor measuring 2.0 m by 3.0 m, is also available in the Khrunichev work area.<br />
6.2.2 Area 31, Buildings 40/40D General Description<br />
Buildings 40 and 40D are located in Area 31 of the Baikonur Cosmodrome. These facilities were modified and<br />
outfitted with the specific needs of SC manufacturers and customers in mind and, along with Building 44, provide the<br />
capability to perform all required SC operations, including off-loading, testing, fueling, mating to the Fourth Stage,<br />
and payload encapsulation. Figure 6.2.2-1 provides the general layout of Buildings 40/40D and their location relative<br />
to the Propellant Filling Hall (Building 44).<br />
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Figure 6.2.2-1: Spacecraft Processing Zone (Area 31)<br />
Building 44<br />
Propellant<br />
Filling Hall<br />
Rm.1<br />
P A S - 5 -<br />
0 9 1<br />
SC fueling area<br />
300-500 m<br />
Area C<br />
H a l l 1 0 0<br />
Building 40<br />
Common Hal l<br />
Building 40D<br />
Preparation Hall<br />
A scent Uni t<br />
Int egrati onA r ea<br />
This section describes the specific areas of Buildings 40/40D as well as the equipment and services available to SC<br />
Customers for pre-<strong>launch</strong> payload processing. Payload processing refers to the final preparation of space payloads,<br />
upper stages, fairings, and related spaceflight support equipment. It is intended to provide sufficient information to<br />
enable customers, and potential customers, to make detailed plans for payload processing activities. It also serves as a<br />
useful reference for the facility areas, equipment, and services available during actual payload processing operations.<br />
6.2.2.1 Facility Layout and Area Designations<br />
Building 40, consists primarily of Hall 100, which is divided into Areas A, B, and C. This facility is used to perform<br />
the following operations:<br />
a) Off-loading/loading of the SC shipping container and support equipment<br />
b) Processing of the Fourth Stage, the LV adapter system, and the Fairing<br />
Area A<br />
Area B<br />
c) Ascent Unit integration (i.e., mating the Fourth Stage, the LV adapter system, the SC, and the Fairing, followed<br />
by electrical checkouts)<br />
d) Verification of the SC (integrated in the Ascent Unit)<br />
e) Transfer of the SC onto the rail transporter<br />
Building 40D is a three-story facility that adjoins Building 40 and includes satellite preparation, ground equipment,<br />
and personnel/office areas.<br />
Details of Buildings 40 and 40D are described and illustrated in the paragraphs that follow.<br />
Rm.<br />
119<br />
SC<br />
Area<br />
Rm.<br />
120<br />
Rm.<br />
121<br />
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6.2.2.2 Building 40, Hall 100 (Common Hall)<br />
Hall 100 is a large high-bay that is divided into three main zones: Area A, Area B, and Area C. SC are delivered by rail<br />
in the manufacturer’s shipping container to Area C and transferred into environmentally controlled Class 100,000<br />
clean areas (Area B or Room 119) for SC processing and integration.<br />
Hall 100 is approximately 120 m long and 30 m wide, with a height of over 18 m. The total usable area is<br />
approximately 2,500 square meters. The room has 8.4 m wide, 10 m high doors to accommodate transport <strong>vehicle</strong>s.<br />
Two overhead traveling bridge cranes are available for lifting and loading operations. This hall houses facilities for<br />
vertical SC integration and the necessary mechanical ground support equipment (GSE) for integration operations with<br />
the Fourth Stage, the LV adapter system, the Fairing, and the assembled Ascent Unit. Figure 6.2.2.2-1 shows the<br />
hall’s general arrangement.<br />
Figure 6.2.2.2-1: Building 40, Hall 100 (Common Hall) Layout<br />
30 m<br />
Thermal<br />
Control Car<br />
Locomotive<br />
Rooms:<br />
100 = Common Hall<br />
119 = Payload Preparation Hall<br />
120 = STE Lowbay<br />
121 = Storage<br />
6.2.2.2.1 Area A<br />
Railway<br />
Transporter<br />
with SC<br />
Area C<br />
120 m<br />
Room 100<br />
Ascent Unit<br />
Integration<br />
Stand<br />
Area A<br />
Area B<br />
Cleanroom for<br />
Ascent Unit<br />
Two Traveling Cranes Assembly<br />
Hook Height = 14.5 m<br />
Load Carrying Capacity 10/50 Tons<br />
Traveling<br />
Crane Above<br />
Area A, which measures approximately 11 m by 47 m, is a Class 100,000 clean area used for final preparation of the<br />
Ascent Unit and its components. It also is used for integration of the SC with the SC flight adapter and then with the<br />
Fourth Stage. Final encapsulation of the spacecraft takes place in this zone. It is separated from Area B by an 8 m high<br />
partition that has an upper opening approximately 18 m long and 4 m high used for transferring cargo by crane<br />
between Areas A and B.<br />
6.2.2.2.2 Area B<br />
Area B measures approximately 19 m by 47 m. It is, essentially, a Class 100,000 airlock used to transition material<br />
from Area C into the Class 100,000 Area A cleanroom, as well as through a door into Building 40D, Room 119.<br />
6.2.2.2.3 Area C<br />
Area C measures approximately 30 m by 73 m and constitutes approximately half of Building 40. It is not strictly<br />
environmentally controlled and is used for loading/unloading SC and support equipment, as well as the Fourth Stage<br />
and Ascent Unit components.<br />
Room<br />
119<br />
Room<br />
120<br />
Room<br />
121<br />
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6.2.2.3 Building 40D, First Floor<br />
The first floor of Building 40D contains the SC preparation hall consisting of Rooms 119, 120, and 121. These rooms<br />
are environmentally controlled, Class 100,000 clean areas. Details of this layout are provided in Figure 6.2.2.3-1. The<br />
total usable area of Rooms 119, 120, and 121 is approximately 400 square meters. Change rooms and restrooms also<br />
are located on the first floor.<br />
Figure 6.2.2.3-1: Building 40D, First Floor Layout<br />
12 m<br />
PA S- 5 - 0 2 2<br />
5.5 m<br />
(H = 7 m)<br />
Room 119<br />
Room 101<br />
Room<br />
122<br />
5. 5 m<br />
2 m<br />
3 m<br />
13 m<br />
114<br />
Room<br />
111<br />
Room<br />
112<br />
Room<br />
104<br />
Room<br />
109<br />
Room 110<br />
Main Entrance<br />
Room<br />
108<br />
4 m<br />
(H = 5 m)<br />
Room 120<br />
H = 4.5 m<br />
Room 121<br />
12 m<br />
9 m<br />
16 m<br />
Rooms:<br />
101 - Personnel changing<br />
room<br />
104 - Air shower<br />
108 - Security Office<br />
109 - Medical Office<br />
112 - Security Checkpoint<br />
116 - Restroom<br />
120 - Control Room<br />
121 - Storage<br />
122 - ILS Storage<br />
The normal personnel entrance into Building 40D is through an outside door past a security checkpoint. This building<br />
is the only one completely controlled by U.S. security for the duration of a <strong>launch</strong> campaign. Security video monitors<br />
are located in the first floor Security Office (Room 108). Access to cleanroom areas (Room 119, 120, 121) is through<br />
an air shower (Room 104) into Room 121.<br />
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6.2.2.3.1 Spacecraft Processing Room 119<br />
Room 119 is used to place support equipment and to prepare the SC prior to propellant loading. The room is an<br />
environmentally controlled Class 100,000 cleanroom; details of its layout are provided in Figure 6.2.2.3-1. The room<br />
is approximately 15.8 m long, 13.8 m wide, and 10 m high. A 4.85 m wide door that is 7 m high provides access to<br />
Hall 100 for the SC. An overhead traveling bridge crane also is available in Room 119.<br />
A header rack to supply compressed air, nitrogen, and helium (at 240 and 350 bars), a blast shield, and service units<br />
(to provide access to the SC) are provided to support stand-alone SC processing.<br />
6.2.2.3.2 Control Room 120<br />
Room 120 is used to place electrical support equipment. The room is an environmentally controlled Class 100,000<br />
cleanroom that is approximately 12 m long, 9 m wide, and 10 m high (Figure 6.2.2.3-1). The room has a 3.4 m wide,<br />
5 m high door that opens on to Room 119.<br />
6.2.2.3.3 Storage Room 121<br />
Room 121 is used to place GSE or other spacecraft-related equipment as required. Like the other rooms in this block,<br />
Room 121 is an environmentally controlled Class 100,000 cleanroom. It is approximately 12 m long, 7 m wide, and<br />
10 m high.<br />
6.2.2.3.4 Support Rooms/Areas<br />
Support rooms on the first floor of Building 40D are shown in Figure 6.2.2.3-1 and include:<br />
a) Personnel changing room (Room 101)<br />
b) Air shower (Room 104)<br />
c) Clean/used cleanroom garment storage<br />
d) Toilet, washroom, and showers<br />
The walls and ceilings of all support rooms are plastered and coated with water-based paint. The walls of toilet and<br />
shower rooms are faced with ceramic tiles throughout their entire height. The floors of the support rooms are cast-inplace<br />
concrete; the floors of the toilets and shower room are covered with ceramic tiles.<br />
SC loading equipment, packing materials, and containers are stored in Area C or in Building 14, which is located near<br />
Building 40. Sealand containers and pallets are stored outside Building 40.<br />
6.2.2.4 Building 40D, Second/Third Floors<br />
The second and third floors of Building 40D serve as offices. Figures 6.2.2.4-1 and 6.2.2.4-2 provide details of the<br />
physical arrangements of these areas.<br />
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Figure 6.2.2.4-1: Building 40D, Second Floor Layout<br />
PA S- 5 -023<br />
Room 201<br />
Room 202<br />
Room 203<br />
Room Rm Room Rm<br />
207 207 206<br />
Figure 6.2.2.4-2: Building 40D, Third Floor Layout<br />
P A S- 5 - 0 2 4<br />
Room 301<br />
Room<br />
303<br />
Room<br />
31 3<br />
Room Rm<br />
Room<br />
312 31<br />
311<br />
1<br />
Room<br />
305<br />
Room<br />
205<br />
Room 3 04<br />
Room 203<br />
Room<br />
308<br />
Room 204<br />
Rooms:<br />
201 - ILS Office<br />
203 - 50 Hz Power<br />
Conditioning<br />
Equipment<br />
204 - Security Office<br />
206/7 Restroom<br />
208 - Khrunichev<br />
Storage<br />
Room 305<br />
Room 307 Room 306<br />
Rooms: 301/307 = SC Customer Offices<br />
304 = Automatic Telephone Exchange<br />
305 = SC Customer Breakroom<br />
306 = Conference Hall<br />
311/312 = Restroom<br />
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The second floor of Building 40D is used by the ILS team. Room 201 (48 m 2 ) is used as the ILS Office Area, and<br />
Room 204 (32 m 2 ) is used for ILS Security. Room 203 houses 50 Hz power racks and is reserved for Khrunichev<br />
engineers. Rooms 206 and 207 are restrooms/washrooms, and Room 208 is used for Khrunichev storage.<br />
The third floor of Building 40D is used by the SC Customer‘s team. It has two available offices: Rooms 301 (48 m 2 )<br />
and 307 (15 m 2 ). Room 305 (42 m 2 ) is normally used as a breakroom. Room 306 (42 m 2 ) is used as a conference<br />
room. Rooms 311 and 312 are restrooms, and Room 304 contains Khrunichev communications switch gear.<br />
6.2.3 Area 31, Bldg 44 - General Description<br />
Building 44 is a SC propellant loading facility located in Area 31 approximately 300 meters from Building 40. This<br />
facility provides the capability to safely perform SC pre-filling operations, propellant thermal conditioning, and SC<br />
fueling in a clean environment. The fueling station has been upgraded to support the loading of the SC with<br />
appropriate propellants. (Propellants are supplied by the SC Customer and are loaded using the SC Customer’s<br />
equipment.)<br />
This section describes the specific areas of Building 44, as well as the equipment and services, available to SC<br />
Customers for payload propellant filling operations.<br />
6.2.3.1 Facility Layout and Area Designations<br />
Building 44, otherwise known as the Propellant Filling Hall, is an 18 m wide by 72 m long highbay divided into three<br />
equal size bays (Figure 6.2.3.1-1).<br />
Commercial SC enter Building 44 through Room 3 and use Filling Room 1 for SC propellant loading. The SC is<br />
placed on the Customer's fueling stand in a clean tent and secured to a tray platform. Room 1 includes a 5.5 m wide by<br />
7 m high sliding door and two emergency exits. Two emergency exits are equipped with panic hardware (each lock<br />
normally keeps the door closed but is immediately released when pressed from the inside).<br />
In addition to the main high-bays, a Control Room on the second floor (Room 58) is used for electrical ground<br />
support equipment to monitor the SC. Fuel and oxidizer decontamination rooms are available for decontamination of<br />
support equipment following loading. Areas for propellant thermal conditioning also are available.<br />
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Figure 6.2.3.1-1: Building 44 (Filling Hall) Floor Plan<br />
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6.2.3.2 Building 44, Filling Hall Room 1 (Clean Tent)<br />
Room 1 in Building 44 is used for servicing of propellants and is approximately 24 m long, 18 m wide, and 15 m high.<br />
The SC entrance doors are 7 m high and 5.5 m wide.<br />
A clean tent is installed in Room 1 to provide a clean zone for the SC and propellant support equipment. The clean<br />
tent consists of two parts: the SC Zone, which measures approximately 9.2 m by 7.6 m; and the Ground Support<br />
Equipment (GSE) Zone, which measures approximately 4.6 m by 6.6 m. Two configuration options are available to<br />
the SC Customer for the clean tent: a pass-through corridor may be installed between the SC and GSE zones to<br />
facilitate personnel passage between the two areas; or a blast shield may be installed between the SC and GSE zones to<br />
prevent personnel passage between the two areas. The clean tent has three exits with ramps.<br />
The walls of the clean tent are made of a transparent, non-particulate-forming, anti-static polymer that is compatible<br />
with propellant vapors. The walls are mounted on a frame made of stainless steel vertical beams and cross-members.<br />
The floor of the tent is made of aluminum, with a stainless steel center section, designed to interface with an SC or SC<br />
adapter. The aluminum floor has channels that are approximately 2 m apart and 3 to 8 cm deep that provide a pathway<br />
for liquid spills to drain to one end of the tent and to the spill tanks located underneath the floor of Building 44. The<br />
oxidizer sump drain is located in Room 2, and the fuel sump drain is located in Room 1.<br />
Conditioned air is supplied to the SC and GSE area through the side walls of the tent at two locations to achieve a<br />
Class 100,000 cleanroom. The air flows into the tent and leaves through outlets at the bottom of the walls. The tent is<br />
pressurized to between 9.8 and 28.6 Pa relative to the outside of the tent. HEPA filters are installed at the inlets of the<br />
stainless steel air ducts. Lighting in the clean tent is provided at an illumination of 100 dekalux. The top and front top<br />
wall of the tent may be opened to facilitate entry of the SC and GSE after tent installation.<br />
6.2.3.3 Building 44, Room 58 (Control Room)<br />
Room 58 on the second floor of Building 44 is the SC Control Room. Room 58 is approximately 6 m wide and 18 m<br />
long and has a window that is covered with explosion-proof glass to facilitate observation of fueling operations. An<br />
umbilical cable may be laid via a blast-proof penetration located between Room 58 and Fueling Hall Room 1.<br />
Telephone, CCTV television monitors, and vapor detection indicators are also provided in Room 58.<br />
6.2.3.4 Building 44, Room 26 and 27 (Fuel/Oxidizer Decontamination Rooms)<br />
Rooms 26 and 27 are available as propellant de-servicing areas to allow the SC Customer to decontaminate SC<br />
propulsion GSE after propellant loading operations. Room 26, the fuel decontamination room, is approximately 5 m<br />
wide and 8 m long and is located adjacent to Filling Hall Room 1. Room 27, the oxidizer decontamination room, is<br />
approximately 9 m wide and 9 m long and is located adjacent to Filling Hall Room 3. Access to these rooms is<br />
provided through exterior entrance doors only.<br />
Each decontamination room is supplied with gaseous nitrogen, which may be used to purge SC ground support<br />
equipment. Floor drains in each room are connected to the facility’s appropriate contaminated water sump tanks.<br />
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6.2.3.5 Building 44 Support Rooms/Areas<br />
There are three support rooms on the first floor of Building 44 that are used by SC personnel. These facilities include:<br />
a) Room 11, which provides the entrance to Room 1 and the locker room<br />
b) Room 12, which is used as the staging area for SCAPE personnel<br />
c) Room 14, which is used as the SCAPE suit donning area<br />
6.2.4 Building 254 (TBD)<br />
6.2.5 Area 92, Building 92-1 - General Description<br />
The Proton LV Processing Room, in Building 92-1 at Area 92, is used to horizontally mate the assembled <strong>launch</strong><br />
<strong>vehicle</strong> stages and strap-on elements.<br />
6.2.5.1 Facility Layout & Area Designations<br />
Building 92-1 is an industrial-type, unprotected building constructed from concrete, brick, and welded/riveted steel<br />
that measures approximately 50 m wide and 120 m long (Figure 6.2.5.1-1). The building includes the<br />
integration/assembly room and two laboratory annexes adjoining the assembly room. The building is provided with<br />
heating, ventilation, and fire-fighting systems, fire and security alarms, and special lighting. An overhead crane is<br />
available for handling the integrated LV.<br />
The Proton LV Processing Room in Building 92-1 is used to:<br />
a) Transfer the Ascent Unit from its railway transportation unit to a mating stand.<br />
b) Mate the Ascent Unit to the Proton LV<br />
c) Conduct electrical checks of the LV transit circuits and verify the hardware links between the Ascent Unit and the<br />
LV<br />
d) Charge the SC’s onboard batteries<br />
e) Install the Thermal Cover on the Ascent Unit<br />
f) Transfer the integrated LV to the erector<br />
g) Prepare the integrated LV for transport to the <strong>launch</strong> complex<br />
The Proton LV Processing Room is approximately 30 m wide, 119 m long, and 22.9 m high. The room has three rail<br />
tracks with a gauge of 1,524 mm. The central track and one of the lateral tracks are throughways, while the third track<br />
terminates in the Processing Room. Along the building’s central axis, there are three double-leaf gates measuring 4.7<br />
by 5.6 m, and one electric-operated rolling (central) gate measuring 8.4 by 10 m.<br />
Temperature and humidity in the room are maintained between 13 and 27 o C and between 30 and 60 percent,<br />
respectively. While it is in the room, the temperature and humidity of the SC are maintained within required levels by<br />
using the Thermal Conditioning Railcar, if necessary.<br />
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Figure 6.5.2.1-1: Building 92-1<br />
6.3 LAUNCH COMPLEX FACILITIES<br />
Following integration of the Proton Launch Vehicle (LV), in Building 92-1, the <strong>launch</strong> <strong>vehicle</strong> is transported to the<br />
Proton Launch Zone, Area 81, for erection, checkout, and <strong>launch</strong>. This section describes the facilities that are used for<br />
erection, checkout, and <strong>launch</strong> of the Proton LV at the Baikonur Cosmodrome. In particular, areas available at the<br />
<strong>launch</strong> complex for Spacecraft (SC) equipment and personnel are described. These areas may be used by the SC<br />
Customer to test and monitor the SC and to charge SC batteries during final LV checkout and <strong>launch</strong>.<br />
6.3.1 Area 81, Launch Pad 23 - General Description<br />
Following integration, the Proton LV is transported to the Launch Complex, Pad 23, located in Area 81 for erection<br />
and <strong>launch</strong>. Figure 6.3.1-1 shows a plan view of the <strong>launch</strong> complex. The <strong>launch</strong> area includes the following physical<br />
facilities, units, and systems that support processing and <strong>launch</strong>ing of the Proton LV:<br />
a) Launch structure with <strong>launch</strong> pad (including underground vault)<br />
b) LV Mobile Service Tower<br />
c) Bunker (Room 250)<br />
d) Facilities for support systems<br />
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Issue 1, Revision 4, March 1, 1999<br />
Figure 6.3.1-1: Proton Launch Zone, Area 81<br />
PUG968<br />
Launch<br />
Pad 23<br />
600 m<br />
Bunker<br />
Rooms 250/251<br />
Launch<br />
Pad 24<br />
340 m<br />
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6.3.2 Facility Layout & Area Designations<br />
6.3.2.1 Launch Structure with Launch Pad (Including Underground Vault)<br />
The <strong>launch</strong> structure and vault house equipment that supports the pre-<strong>launch</strong> processing of the LV. They provide<br />
electric, pneumatic, and hydraulic links between the ground system testing equipment and on-board LV hardware via<br />
the LV transit cables and pipes. The <strong>launch</strong> structure is designed to withstand LV first-stage engine plume<br />
impingement. The <strong>launch</strong> pad is intended for LV installation, erection, securing the LV in a vertical position.<br />
6.3.2.1.1 Room 64 - Vault<br />
Room 64 (Vault), along with Room 250 (Bunker), can be used to house the SC Customer’s ground support equipment<br />
(GSE). Room 64 measures approximately 5.1 by 5.6 m and is equipped with 60 Hz electrical power, grounding, and<br />
communications services. All <strong>launch</strong> campaign operations requiring the presence of personnel in the vault must be<br />
completed prior to the start of LV fueling, which occurs approximately 7 hours prior to <strong>launch</strong>. All personnel are<br />
required to leave the vault by this time, and from then on, all vault equipment must be controlled remotely from the<br />
bunker (Room 250). Additional details about Room 64 are provided in Drawing 4.1-1. Plate 4.1-1 provides a<br />
photographic overview of the vault’s interior.<br />
6.3.2.1.2 Room 76 - Vault<br />
Room 76 can be used to house the SC Customer’s GSE. Room 76 measures approximately 5.4 m by 10.8 m and is<br />
equipped with 60 Hz electrical power, grounding, and communications services. All <strong>launch</strong> campaign operations<br />
requiring the presence of personnel in the vault must be completed prior to the start of LV fueling, which occurs<br />
approximately 7 hours prior to <strong>launch</strong>. All personnel are required to leave the vault by this time, and from then on, all<br />
vault equipment must be controlled remotely from the bunker (Room 250).<br />
6.3.2.2 Mobile Service Tower<br />
The Mobile Service Tower provides access to the SC and LV and houses equipment to support SC and LV pre-<strong>launch</strong><br />
processing and <strong>launch</strong>. The Mobile Service Tower includes service platforms, a gallery, service fixtures, two cargopassenger<br />
elevators (500 kg rated load capacity each), and two cranes (rated load capacity 500 and 5,000 kg,<br />
respectively). An overall view of the Mobile Service Tower is provided in Figure 6.3.2.2-1.<br />
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Figure 6.3.2.2-1: Proton Mobile Service Tower<br />
30.515 m<br />
30.400 m<br />
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6.3.2.3 Rooms 250/251 - Bunker<br />
The Bunker (Rooms 250/251) is used to support a Proton <strong>launch</strong>. It is located 1.3 km from the <strong>launch</strong> pad and<br />
provides protection for personnel and equipment during the <strong>launch</strong>. The bunker houses the LV and SC system test<br />
equipment and GSE required for pre-<strong>launch</strong> operations and monitoring of SC readiness. Air temperature and<br />
humidity inside the bunker are controlled by an air-conditioning unit.<br />
Room 250 is allocated for installation of SC equipment, and Room 251 is the ILS Security Office.<br />
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7. LAUNCH CAMPAIGN<br />
7.1 ORGANIZATIONAL RESPONSIBILITIES<br />
In any given <strong>launch</strong> campaign several organizations are involved. What follows is a brief overview of the responsibilities<br />
of these organizations.<br />
7.1.1 Khrunichev<br />
a) Overall responsibility for coordinating work performed at the <strong>launch</strong> complex by the Strategic Rocket Forces (and<br />
NPO Energia if Block DM used)<br />
b) Engineering support and quality inspection for all testing performed on Stages 1-3 of the Launch Vehicle, as well<br />
as the adapters and fairing. For <strong>launch</strong>es with the Breeze M Khrunichev is also responsible for Breeze M<br />
engineering, inspection and test<br />
c) Maintenance of Buildings 92A-50 (Halls 103A, 103, and 101) and 40/40D; and the hotel complex<br />
d) Transportation and food services<br />
e) Coordinating Baikonur and Leninsk medical services with Strategic Rocket Forces<br />
f) Launch complex security<br />
7.1.2 Strategic Rocket Forces<br />
a) Maintenance of Building 44, Building 92-1, Building 92A-50 (Hall 102), and the Launch Complex<br />
b) Provision of technicians for performing <strong>launch</strong> <strong>vehicle</strong> testing<br />
c) Provision of quality inspectors<br />
d) Launch Vehicle operations from integration in Building 92-1 through erection on pad and <strong>launch</strong><br />
e) Launch complex security<br />
7.1.3 Energia (Block DM <strong>launch</strong>es only)<br />
a) Checkout and processing of the Fourth Stage at the Baikonur Cosmodrome from its arrival through the <strong>launch</strong>.<br />
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7.1.4 ILS<br />
a) Prime interface between the Customer and Khrunichev<br />
b) Coordinating campaign schedules and operations with the SC Customer and Khrunichev<br />
c) Logistics<br />
d) Safety overview as an advisory function to SC and Customer management, as well as physical security of SC assets<br />
while at Baikonour processing facilities<br />
e) Translation and interpretation services<br />
f) Medical doctor and SOS evacuation<br />
7.1.5 Spacecraft Customer<br />
a) Spacecraft checkout and processing at the Baikonur Cosmodrome<br />
7.2 CAMPAIGN ORGANIZATION<br />
7.2.1 Contractual and Planning Organization<br />
The fundamental contractual relationships among the principal parties in any given <strong>launch</strong> campaign are as follows:<br />
a) ILS is a contractor to the SC Customer<br />
b) Khrunichev is a subcontractor to ILS<br />
All matters that could potentially affect the terms of the Launch Service Agreement (the contract) between a SC<br />
Customer and ILS must be dealt with by the SC Customer and ILS. Matters affecting the terms of the subcontract<br />
between ILS and Khrunichev must be dealt with by ILS and Khrunichev. In particular, any issues involving possible<br />
additional costs must be mutually agreed upon through these contractual relationships.<br />
ILS will coordinate all logistics support and operations planning with both the SC Customer and Khrunichev.<br />
ILS may assign a safety engineer to monitor any given operation to ensure that all activities are carried out in<br />
conformance with the mutually agreed upon safety plan. This safety engineer is present for all hazardous operations.<br />
Figure 7.2.1-1 provides a diagrammatic representation of a typical operational organization used during a <strong>launch</strong><br />
campaign.<br />
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Figure 7.2.1-1: Organization During Launch Campaign<br />
Customer MM<br />
SC Subcontractor<br />
ILS MM<br />
SC Ops Manager ILS Ops Manager<br />
7.2.2 Organization During Combined Operations<br />
Khrunichev<br />
Program Director<br />
Khrunichev<br />
Ops Manager<br />
Russian Strategic<br />
Rocket Forces<br />
Other Subcontractors<br />
Combined operations are those operations involving some combination of SC Customer organization, ILS, and<br />
Khrunichev personnel (e.g., Khrunichev adapter mating, Fourth Stage to payload mating, encapsulation, Ascent Unit<br />
checkout, Ascent Unit integration to Stages 1-3, any operations on the pad which require payload faring access). For<br />
each such combined operation, one operation leader is assigned, either from Khrunichev or the SC Customer. This<br />
individual directs the operation and ensures that it is carried out in conformance with the mutually agreed upon<br />
procedures.<br />
For each operation, one person from the SC Customer organization, ILS and Khrunichev is designated as team leader<br />
for their respective organizations. Agreements among organizations can only be reached among these three team<br />
leaders.<br />
Security personnel from either or both ILS and the Strategic Rocket Forces may be present during any operation if<br />
required by the Security Plan.<br />
ILS provides at least one interpreter for each combined operation. Special training is conducted with the SC Customer<br />
and Russian personnel for joint crane operations to ensure reliable communications between English and Russianspeaking<br />
personnel.<br />
Either Khrunichev or the SC Customer provides a Quality Assurance Representative for each operation who<br />
documents any test discrepancies on a Quality Assurance Report.<br />
Strategic Rocket Forces personnel implement many of the operations at the Baikonur Cosmodrome. The Strategic<br />
Rocket Forces is a Khrunichev subcontractor and, as such, coordinates directly with Khrunichev and NOT with the SC<br />
Customer or ILS personnel.<br />
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7.2.3 Planning Meetings<br />
ILS maintains the master schedule for combined operations planning and reviews it with the SC Customer and<br />
Khrunichev at a daily scheduling meeting. At this meeting, the current operations for the day are reviewed and the next<br />
3 day’s operations are agreed upon. Following each meeting, ILS revises the master schedule for the following day. At<br />
a minimum, the following personnel must attend the daily scheduling meetings:<br />
a) ILS Mission Manager or Launch Operations Manager<br />
b) SC Customer Launch Campaign Manager<br />
c) Khrunichev Mission Manager<br />
d) Strategic Rocket Forces representative in charge of the facility<br />
At certain stages in the campaign, ILS holds meetings to give the go-ahead for critical phases of a campaign. These<br />
critical phases include:<br />
a) SC offload and move to Integration Hall<br />
b) SC Processing and Propellant Loading<br />
c) SC Encapsulation<br />
d) Launch<br />
Two critical meetings require high-level concurrence prior to proceeding to the next phase of the campaign. These are:<br />
a) Vehicle Readiness for Transport to Pad—5 days prior to <strong>launch</strong><br />
b) Vehicle Readiness for Propellant Load—10 hours prior to <strong>launch</strong><br />
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7.3 COUNTDOWN ORGANIZATION<br />
The countdown organization is illustrated in Figure 7.3-1.<br />
Figure 7.3-1: Countdown Organization<br />
SC Customer<br />
Mission<br />
Director<br />
ILS<br />
Mission<br />
Manager<br />
Energia<br />
Mission<br />
Director<br />
(if Block DM<br />
used)<br />
Khrunichev<br />
Mission<br />
Director<br />
KBOM*<br />
Readiness<br />
Review Board<br />
Authorization Strategic Rocket<br />
Forces Launch<br />
Commander<br />
Launch<br />
Authorization<br />
SC Ready<br />
L - 10 min<br />
Spacecraft<br />
Director<br />
(SC Customer)<br />
Automatic<br />
Computer-Controlled<br />
Sequencer<br />
LV Ready<br />
L - 5 min<br />
Stages 1-3<br />
Director<br />
(Khrunichev)<br />
LIFT-OFF<br />
COMMAND<br />
Fourth Stage<br />
Ready<br />
L - 2 min<br />
Fourth Stage<br />
Director<br />
(Energia) if<br />
Block DM<br />
used)<br />
*Design Bureau for General Machine Engineering (Launch Site) Note: Launch Commander has sole authority to abort <strong>launch</strong><br />
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The Strategic Rocket Forces directs the countdown, which follows a pre-approved script. The Launch Commander<br />
receives authorization to <strong>launch</strong> from the Readiness Review Board, which consists primarily of the five entities shown<br />
in Figure 7.3-1.<br />
Certain organizations have pre-assigned abort capability. Each of these organizations is asked to acknowledge the<br />
readiness of their subsystems on Launch Day according to the Launch Day Script. These subsystem readiness checks<br />
are as follows:<br />
a) Stages 1-3 Readiness: Khrunichev<br />
b) Fourth Stage Readiness: Khrunichev for Breeze M, Energia for Block DM<br />
c) SC Readiness: SC Customer<br />
Each organization designates a single individual to provide the above authorities on Launch Day, and each<br />
representative is vested with abort authority over the <strong>launch</strong> sequencer for their respective area of responsibility. For<br />
example, the SC Customer may abort the start sequence as late as 2.5 seconds prior to lift-off contact.<br />
7.4 ABORT CAPABILITY<br />
There is a ground circuit that, when closed, sends a command to an onboard switch that, in turn, activates on-board<br />
power to the first three stages and begins the sequence for first stage engine ignition. Normally, readiness signals are<br />
given in the following sequence:<br />
a) Launch - 10 minutes: SC Ready<br />
b) Launch - 5 minutes: Stages 1-3 Ready<br />
c) Launch - 2 minutes: Fourth Stage Ready<br />
Once all readiness signals are received, the command circuit is ready and is waiting only for the Launch Vehicle Start<br />
Command to be issued automatically by the Start Timer Mechanism. This command is issued at 2.5 seconds prior to<br />
liftoff, at which time the final relay closes, sending a command transferring to on-board power and initiating the First<br />
Stage engine ignition sequence.<br />
The SC Customer is capable of aborting the Launch Vehicle Start Command issued by the Start Timer Mechanism.<br />
The <strong>launch</strong> aborts if this circuit is open when the Launch Vehicle Start Command is issued. If the circuit is opened<br />
following issuance of the Launch Vehicle Start Command, the <strong>launch</strong> can no longer be aborted. The SC Customer can<br />
abort the <strong>launch</strong> countdown up to 2.5 seconds prior to lift-off contact.<br />
The Fourth Stage Manager can abort the <strong>launch</strong> by removing power from the command circuit. He has this capability<br />
up to 2.5 seconds prior to lift-off contact.<br />
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7.5 LAUNCH OPERATIONS OVERVIEW<br />
This section provides an overview of Launch Vehicle (LV) and Spacecraft (SC) processing.<br />
Figure 7.5-1 provides a similar schedule assuming the use of Building 92A-50 for SC processing. Figure 7.5-2 provides<br />
details of the SC processing timeline using Area 31 facilities.<br />
The typical duration of a <strong>launch</strong> campaign from SC arrival to <strong>launch</strong> is 30 days, depending on SC manufacturer and<br />
Customer requirements.<br />
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Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50<br />
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Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50 (Continued)<br />
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Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31<br />
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Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31 (Continued)<br />
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7.5.1 Launch Vehicle Processing<br />
The Proton LV stages and fairings are built in Moscow by Khrunichev and transported by rail to the Baikonur<br />
Cosmodrome typically well in advance of the SC delivery date. After transportation of the Proton’s stages and fairing<br />
by rail, LV build-up takes place in an integration and test facility (Building 92-1) capable of supporting four<br />
simultaneous Proton assembly and checkout operations. The Fairing is moved either to Building 40 or to Building<br />
92A-50, depending upon which facility is to be used for SC integration, prior to SC arrival, there it is stored and<br />
cleaned in preparation for encapsulation. The Fourth Stage is delivered to Area 254 for pre-<strong>launch</strong> checkout and<br />
testing. The Fourth Stage is then delivered to the Building 44 in Area 31, the propellant fueling hall, where either<br />
kerosene is loaded (Block DM) or MMH and N 2O 4 (Breeze M). It is then moved to either Building 40 or 92A-50 for<br />
integration with the SC. Payload adapters typically are delivered shortly before the processing cycle and cleaned and<br />
prepared in the Integration Hall (100A or 101) of whatever processing area is being used for that program.<br />
In advance of SC arrival, Payload Processing Facilities undergo facility activation and certification. Buildings 40,<br />
40D, 44, or 92A-50 are verified to meet environmental control and cleanliness requirements, in addition to<br />
commodities and power support requirements usually a week prior to the SC arrival date.<br />
7.5.2 Spacecraft Preparations Through Arrival<br />
Prior to campaign start, SC propellants are shipped by rail from St. Petersburg to the Baikonur Cosmodrome. They<br />
are stored in the same temperature-controlled railcars used for transport from St. Petersburg until required for fueling.<br />
A pre-determined number of hours prior to fueling, as defined by the spacecraft manufacturer, propellant containers<br />
are transferred to a storage/conditioning room for temperature stabilization.<br />
The SC and its ground support equipment (GSE) arrive at Yubeleini Airport via a SC Customer-chartered aircraft,<br />
where they are offloaded and loaded onto railcars. These operations are supported by Khrunichev-supplied mobile<br />
cranes, K-loader, and 5 and 15-ton forklifts as required. After the SC container is placed on a railcar, it may be<br />
connected to a thermal control railcar via two air duct flanges (inlet and outlet air flow) to provide thermal<br />
conditioning during transport. Some spacecraft containers are completely self-contained thermally and<br />
environmentally and do not require this support option. The thermal car also provides a dynamic load monitoring<br />
system, so use of the car may be limited to use of this system. SC Customer personnel effects may be transported<br />
directly to the hotel by truck.<br />
7.5.3 Area 31 (Buildings 40, 40D, 44) - Spacecraft Testing, Fueling, and Ascent Unit integration<br />
The SC and its GSE are transported by rail convoy approximately 70 km to Area 31; transport duration is<br />
approximately 7-10 hours, depending on time of day of transport.<br />
The SC and its GSE are offloaded outside of Building 40, Room 100, Zone C using 5 and 15-ton forklifts. The<br />
Building 40 bridge crane may be used for offloading large GSE such as the SC container. Spacecraft GSE and<br />
container handling proceed as required for each unique spacecraft operations flow.<br />
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A typical flow might progress as follows:<br />
a) SC container offloaded from railcar in Rm 100C, using overhead bridge crane, and container lid removed<br />
b) SC rotated to the vertical position and moved into Rm 119, placed on transporter<br />
c) SC electrical checkout equipment moved into Rm 119<br />
d) SC standalone testing, electrical and mechanical systems<br />
Note: A portable blast shield is available for use if required for high-pressure leak checks<br />
Following Fourth Stage filling in Building 44 (Filling Hall), Room 1 of the Filling Hall is readied to receive the SC.<br />
This includes setting up the clean tent and all fluid/gas systems. The SC may be bagged to protect it during any brief<br />
exposures to less than Class 100 000 environments. It is then rolled out of Building 40D Room 119 to Building 40,<br />
Room 100, Zone B, where it is transferred to the support on the base of the Thermal Transport Container. The SC<br />
maintains its vertical orientation, as the Thermal Transport Container is designed to interface to the SC adapter<br />
mating surface. This container is used to transfer the SC to the Filling Hall for propellant fueling. Once on the<br />
container base, the SC container cover is installed, and the container is lifted onto the railcar in Zone C for transport.<br />
The thermal control railcar may be used to provide conditioned air and dynamic load data during transportation. Use<br />
of this temperature control system may be waived by the SC Customer since the nominal time for transportation to the<br />
Filling Hall is of short duration (less than 1 hour). The Thermal Transport Container has been specifically designed to<br />
provide thermal stability for an extended period of time.<br />
The SC is moved to Building 44 (Filling Hall) 250 meters away, and the Thermal Transport Container cover is<br />
removed in Room 2 and placed on support stands. The railcar is then rolled into Room 1 and the SC is hoisted from<br />
the Thermal Transport Container base onto the SC loading stand already installed in the clean tent. The clean tent<br />
ceiling is closed and a Class 100 000 environment is established.<br />
Propellants are loaded in accordance with the SC Manufacturer’s procedure. The facility provides passive vapor vent<br />
and active vapor extraction, as well as liquid waste disposal and commodities such as water, breathing air, and GN2.<br />
Propellant GSE is decontaminated using rooms in the Filling Hall specifically dedicated to this purpose. Hot and cold<br />
GN2 purge is available as required.<br />
Following fueling, the process is reversed and the SC is returned to the Thermal Transport Container and then moved<br />
back to Building 40, Room 100 for further testing and processing. Electrical checkouts are made on the adapter<br />
electrical harnesses. The SC is moved from Zone B to Zone A and placed onto the flight adapter (s).<br />
The SC and its adapter is then mated to the Fourth Stage. For certain SC, this may require the use of an integration<br />
stand commonly referred to as the “Gallows”. The Gallows allows taller payloads to be installed onto the Fourth<br />
Stage without structurally modifying Building 40D. The flight clamp band is installed and tightened to the correct<br />
flight-specific tension.<br />
Following integration, an end-to-end electrical checkout is performed. The composite Ascent Unit is rotated to the<br />
horizontal for encapsulation with the two fairing halves. Following encapsulation, an RF GO/NO GO test is<br />
performed through the RF window. The Ascent Unit is then transferred to its railcar for transport to Area 92. During<br />
transport, the Ascent Unit is thermally conditioned using the thermal conditioning railcar; input air to the Ascent Unit<br />
is provided at the base of the Fourth Stage and output occurs at the Fairing nose.<br />
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7.5.4 Area 92 (Building 92A-50) - Spacecraft Testing, Fueling, and Ascent Unit Integration<br />
In this processing scenario, the SC and its GSE are transported by rail convoy approximately 30 km to Building 92A-<br />
50, Room 102 , where external cleaning of the SC container is performed in Hall 102. Final cleaning is completed in<br />
this area prior to moving the SC to the Integration Hall (Hall 101). Hall 101 is a Class 100,000 cleanroom, and the<br />
SC container cover may be removed here or in Hall 103A as required by the unique <strong>mission</strong>-specific SC processing<br />
flow.<br />
A typical flow is as follows:<br />
a) Move the SC container into Hall 101 on railcar<br />
b) Remove container from the railcar and place on the floor in Hall 101<br />
c) Remove the container lid<br />
d) Roll container into 103A if on wheels or<br />
e) Lift SC Container onto KhSC-supplied air pallet or Customer-supplied transport dolly to roll into 103A<br />
f) Move SC on transporter into Hall 103A for processing<br />
Note: Container lid removal may also be accomplished in Hall 103A as required by the Customer.<br />
The SC container and lid may be stored in Hall 101. Electrical test equipment is brought into the control room by<br />
means of an external door which opens directly into the control room loading area. This is a small buffer zone between<br />
two sets of double doors with a concrete floor.<br />
After container removal, SC electrical testing, pneumatic testing, and propellant fueling occurs in Hall 103A. Passthroughs<br />
from the control room are available for cabling. These cable feeds are verified to be leak-tight prior to<br />
propellant operations. The 60 Hz, 120 Vac power source is provided by an UPS. Typically the UPS is activated the<br />
week prior to SC arrival and not deactivated until the SC leaves the facility and all parties agree that no further<br />
requirement for it exists. A portable blast shield is available for high-pressure tests.<br />
For propellant operations, the facility is configured with liquid waste aspirators, passive vent scrubbers, and a vapor<br />
detection system which alarms locally, in the control room, and at the Security Command Post. Breathing air is<br />
supplied by a single source which is sampled prior to operations. GN2, water, and shop air are provided on demand. A<br />
fire suppression system which will arm but not release on alarm is also active in Hall 103A. The command to activate<br />
the suppression system deluge is made in the control room, and is not an automatic function of the alarm system. LN2<br />
is available with 24 hr call-up.<br />
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The loaded SC is transported back to Hall 101 using the Transport Dolly. The SC is lifted from the transporter and<br />
mated to the SC adapter using the 50T bridge crane. The SC/adapter unit is then lifted and mated to the Fourth Stage<br />
which has been previously installed in the Integration Stand. The adapter clampband is installed and tensioned.<br />
Incremental electrical continuity tests are performed at each phase, with the final check being an end-to-end test with<br />
the SC mated to the Fourth Stage.<br />
After SC integration, final closeout operations and photographs are performed. The combined Fourth Stage/SC is<br />
rotated to the horizontal position on the integration stand and is encapsulated by the bi-conic payload Fairing. After<br />
the upper fairing half is emplaced, an RF Go/NoGo test is performed to ensure that the SC link has not been disturbed<br />
and that the RF window is transparent to RF. This is performed as soon as the fairing half is mechanically emplaced<br />
and before continuation of encapsulation sealing operations. If any anomaly is found, the fairing may be removed<br />
relatively easily at this point. After determining a good RF signal, encapsulation is completed.<br />
After encapsulation and required RF testing, the integrated Ascent Unit is placed on a railcar for transport to Building<br />
92-1 (Integration Hall) for integration with the Proton Launch Vehicle. At this point, the processing flow for all SC<br />
becomes a common flow, independent of SC Processing Facility.<br />
7.6 LAUNCH VEHICLE INTEGRATION (BUILDING 92-1) THROUGH LAUNCH PAD<br />
OPERATIONS<br />
The Ascent Unit is transported to Building 92-1, where it is disconnected from the thermal conditioning car and lifted<br />
onto the integration dolly. Here, it is brought together horizontally with the mating surface of Stage Three of the<br />
integrated Proton <strong>launch</strong> <strong>vehicle</strong> (LV). An end-to-end electrical check is performed on the SC/LV cables. The<br />
integrated LV is then transferred to the transportation/erection fixture and a thermal blanket is installed over the<br />
Fairing to protect the payload from temperature extremes during periods when there is no active thermal control. A<br />
typical <strong>launch</strong> flow requires 4 days in the Integration Hall. Activities in the Integration Hall are controlled by the<br />
Strategic Rocket Forces and are based on the LV Launch schedule.<br />
For LVs with a Block DM, the LV will be transported out to the pad directly from 92-1 at L-5. Fr LVs with the<br />
Breeze M, the LV will be transported first to 92A-50 for final integrated Breeze M tests. This will be followed by<br />
transportation to the Breeze M Filling Station for top off of the low pressure MMH and N2O4 reservoirs on its way to<br />
Launch Pad 24.<br />
The first of two Inter-governmental Com<strong>mission</strong> Meetings is held on Day L-6, prior to <strong>vehicle</strong> roll-out to the pad, to<br />
ensure all agencies are go for pad roll-out. All LV agencies, including the SC Manufacturer will be called upon to<br />
provide a <strong>launch</strong> readiness statement.<br />
The LV along with the thermal control railcar is moved to the <strong>launch</strong> pad. The <strong>vehicle</strong> is erected, and the liquid<br />
thermal conditioning system is activated. An RF check of the SC telemetry and command links is performed from the<br />
Bunker (Room 250) prior to Mobile Service Tower rollup. Once the Mobile Service Tower is in place, the Ascent Unit<br />
air-conditioning system is activated and the liquid system is turned off.<br />
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7.7 LAUNCH PAD OPERATIONS<br />
The Integrated LV is transported to the Launch Pad and erected in one piece at L-5 using the Launch Vehicle<br />
Transporter. From L-5 on, the <strong>launch</strong> schedule is driven by the Strategic Rocket Forces overall countdown schedule.<br />
Coordination of SC-related pad activities is performed through the 7/701 Script. The 7/701 Script is generated by<br />
Khrunichev with SC/Customer input and should include all pad access requirements and requirements for RF<br />
radiation and commanding the SC. Operator “7” is the Khrunichev Program Director while Operator “701” is the SC<br />
Point of Contact. Information required from the SC/Customer: RF radiation, battery charging, SC commanding, pad<br />
access. Note that ILS functions such as pad access will also be coordinated through this script. Figure 7.7.1 provides a<br />
typical detailed on-pad operations flow.<br />
The SC Customer is expected to participate in a Launch Countdown Rehearsal on Day 3 on pad. This countdown<br />
rehearsal is supported by a full booster <strong>vehicle</strong> <strong>launch</strong> crew countdown and requires the SC Customer to indicate SC<br />
readiness to go at the required time. The rehearsal also includes a planned abort on a SC No-Go condition. SC full<br />
fidelity countdown rehearsal is not required for this exercise, simply the operation of the readiness switch at the<br />
planned time in accordance with the 7/701 script.<br />
The second Inter-Governmental Com<strong>mission</strong> Meeting is held at T-8 hours, to ensure all agencies are go for <strong>launch</strong><br />
prior to propellant load of the booster <strong>vehicle</strong>. At T-8 hours, the <strong>launch</strong> pad is cleared of all non-essential personnel,<br />
and at T-6 hours, propellant load commences. At T- 2.5 hours, the pad is open for final closeouts and service tower<br />
removal. At T-2 hours, all personnel are cleared from the hotel areas and should be in their final positions for <strong>launch</strong><br />
i.e. Bunker, Viewing Area, and Comm Center. Note that personnel in the Bunker and the Comm center should be<br />
limited to essential personnel only.<br />
The SC Customer participates in the final countdown by sending a SC readiness to <strong>launch</strong> signal at T-10 minutes, as<br />
noted in the Countdown Organization discussion.<br />
Active commanding of the SC is prohibited during critical booster <strong>vehicle</strong> functions. Propellant fueling of the booster<br />
<strong>vehicle</strong>, which starts at T- 8 hours, is one such time-frame. Other typical RF silence and no-command times are<br />
shown in Figure 7.7-2.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Figure 7.7-1: Typical Launch Pad Operations Timeline<br />
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8. PROTON LAUNCH SYSTEM ENHANCEMENTS<br />
The Proton <strong>launch</strong> system is supported by the Russian space industry leaders Khrunichev State Research and<br />
Production Space Center and Russian Space Complex Energia. Based on the long, successful history of the Proton<br />
<strong>launch</strong> system, ILS, Khrunichev, and Energia have continued to assess and conceptually study hardware and <strong>launch</strong><br />
services enhancements that would improve the competitiveness and capability of the Proton <strong>launch</strong> system.<br />
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Figure 8-1: Proton Evolution Options<br />
Years of operat io n<br />
LEO capability<br />
GSO capability<br />
2000-<br />
22 .0 M T<br />
4.5 M T<br />
Proton-KM<br />
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A number of enhancement options are being studied for the Proton system including enhanced payload fairing options,<br />
avionics enhancements, and upper stage modifications to further improve capabilities of Proton heavy lift <strong>launch</strong><br />
system.<br />
8.1 HIGH VOLUME PAYLOAD FAIRINGS<br />
With the quantity of LEO, GTO, and planetary spacecraft on the Proton <strong>launch</strong> record, a number of payload fairing<br />
configurations have been developed in the past to uniquely meet the needs of the diverse Russian government<br />
programs. Most notable in recent developments has been the <strong>launch</strong> of large low earth orbit payloads such as the MIR<br />
space-station using the three stage Proton <strong>vehicle</strong>.<br />
With adoption of the Breeze M upper stage, larger payload fairings can be easily mounted to the Proton booster<br />
<strong>vehicle</strong>. Unlike the Block DM configuration <strong>vehicle</strong>, the design of the Breeze M enables the stage to be completely<br />
encapsulated inside the fairing. For the Breeze M configuration, the payload fairing is attached directly to the third<br />
stage of Proton enabling large volume/higher mass fairings to be successfully carried with the limits of the Proton<br />
design. A number of payload fairing options have been assessed for development with the commercial Proton<br />
M/Breeze M configuration. Figure 8.1-1 illustrates some of those options.<br />
With these conceptual systems, usable volume diameters as great as 4.2-meters have been envisioned. Usable<br />
cylindrical volume lengths of up to 15,000-mm have been assessed and determined feasible with the existing control<br />
authority capability of the Proton <strong>launch</strong> system.<br />
ILS and KhSC are willing to develop these concepts with the award of firm <strong>launch</strong> services contracts. These fairing<br />
options can be fielded in 30-36 months of authority to proceed and will support Proton M and Proton M/Breeze M<br />
<strong>vehicle</strong> <strong>launch</strong>es as early as mid-2001.<br />
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Figure 8.1-1: Proton Larger Payload Fairing Concepts<br />
TBS<br />
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8.2 TANDEM LAUNCH SYSTEM<br />
In addition to large volume payload fairing options, ILS and KhSC have studied and conceptually designed a tandem<br />
<strong>launch</strong> system payload carrier concept. The tandem <strong>launch</strong> system concept may offer attractive opportunities for<br />
<strong>launch</strong> of multiple payload constellations to low and intermediate altitude orbits. Figure 8.2-1 illustrates possible<br />
interfaces available with the tandem <strong>launch</strong> system concept.<br />
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Issue 1, Revision 4, March 1, 1999<br />
Figure 8.2-1: Breeze-M Launch Configuration With Tandem Launch Systems (TLS)<br />
5<br />
0<br />
30<br />
6<br />
4<br />
39<br />
45<br />
Φ 4350<br />
Φ 2490<br />
Φ 4100<br />
30<br />
20<br />
21<br />
1<br />
4<br />
1 6<br />
1<br />
6<br />
2<br />
6<br />
19<br />
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8.3 AVIONICS SYSTEM MASS UPGRADES<br />
With the introduction of the Breeze M stage, it becomes possible to remove the booster avionics system from the third<br />
stage of Proton and use a modified upper stage avionics system for control of the entire <strong>launch</strong> system from liftoff to<br />
separation of the payload in the targeted orbit. Removal of the avionics system increases typical performance to<br />
geosynchronous transfer orbits by approximately 100 kg.<br />
8.4 NEW CYROGENIC UPPER STAGE<br />
In a the course of a <strong>vehicle</strong> development contract with another entity, Khrunichev has developed a liquid oxygen,<br />
liquid hydrogen upper stage. This upper stage can be adapted to fly on the Proton booster and could significantly<br />
enhance <strong>launch</strong> system performance to high energy transfer orbits. Payload performance gains to GSO and GTO of<br />
almost 50% are possible over that available with the Breeze M upper stage.<br />
8.5 SUMMARY<br />
The mature partnership between Lockheed Martin and Russian partners Khrunichev Space Center and RSC Energia is<br />
now enabling ILS to explore the next evolutionary steps in effectively supporting Proton commercial <strong>launch</strong> services.<br />
ILS is ready to discuss, with potential customers, straight forward enhancements in Proton <strong>launch</strong> services capability<br />
to meet near term commercial <strong>launch</strong> services needs.<br />
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A. PROTON LAUNCH SYSTEM HISTORY<br />
A.1 BACKGROUND AND HISTORY<br />
Development of the Proton <strong>launch</strong> <strong>vehicle</strong> was undertaken in the early 1960s, under the direction of the Soviet<br />
academician, V. N. Chelomey. The first <strong>launch</strong> took place in July 1965. The two-stage 8K82 (D) version, last flown<br />
in 1966, was used to <strong>launch</strong> four flights of the Proton satellite series, from which the <strong>launch</strong> <strong>vehicle</strong> takes its name.<br />
The two-stage version has been superseded by the three stage Proton K also known as (D-1, SL-13) model and the<br />
four-stage Proton K/Block DM (D-l-e, SL-12) model, both of which are currently in use. An improved version of<br />
the Proton (Proton M) is now in development, incorporating changes to the first three stages, as well as the<br />
development of a new storable propellantBreeze M upper stage. Figure A.1-1 shows the major variants of the Proton<br />
<strong>launch</strong> <strong>vehicle</strong> family.<br />
Figure A.1-1: Proton Launch Vehicle Family<br />
Years of operation<br />
LEO capability<br />
GTO capability<br />
GSO capability<br />
8K82<br />
(Proton D)<br />
1965-1967<br />
12.2 MT<br />
-<br />
-<br />
Proton K<br />
(Proton D-1)<br />
1967<br />
19.76 MT<br />
-<br />
-<br />
Proton K/Block DM<br />
(Proton D-1-e)<br />
1967- Present<br />
-<br />
4.9 MT<br />
1.9 MT<br />
Proton-M<br />
Breeze-M<br />
1999- Present<br />
21.0 MT<br />
5.5 MT<br />
2.92 MT<br />
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The Block DM fourth stage of the Proton was developed independently during the 1960s as the fifth stage of the<br />
Russian manned lunar <strong>launch</strong> <strong>vehicle</strong>, the N2-L3. It was originally known in Russia as the Block D ("block" is the<br />
common translation of the Russian word denoting a rocket "stage", while "D" is the fifth letter in the Russian<br />
alphabet). The <strong>vehicle</strong> was upgraded during the 1970s to the current Block DM (modernized) version.<br />
The Proton model numbers D, D-1, D-l-e, SL-13, and SL-12 were the designations in prior use in the United States,<br />
with the D numbers having been applied by the Library of Congress and the SL numbers originating with the<br />
Department of Defense.<br />
Proton has flown more than 235 <strong>mission</strong>s, and has carried the Salyut series space stations and the Mir space station<br />
modules. It has <strong>launch</strong>ed the Ekran, Raduga, and Gorizont series of geostationary communications satellites (which<br />
provided telephone, telegraph, and television service within Russia and between member states of the Intersputnik<br />
Organization), as well as the Zond, Luna, Venera, Mars, Vega, and Phobos inter-planetary exploration spacecraft.<br />
The Proton has also <strong>launch</strong>ed the entire constellation of Glonass position location satellites. All Russian geostationary<br />
and interplanetary <strong>mission</strong>s are <strong>launch</strong>ed on Proton. Approximately 90% of all Proton <strong>launch</strong>es have been the fourstage<br />
version.<br />
The Proton <strong>launch</strong> <strong>vehicle</strong> has a long history of outstanding reliability. From its first operational <strong>launch</strong> in 1970 to the<br />
present day, Proton has averaged a 92.5% success rate. Today the Proton <strong>launch</strong> <strong>vehicle</strong> has a 92% (moving average)<br />
success rate over its last 50 <strong>launch</strong>es. The recent history of Proton's <strong>launch</strong> reliability is shown in Table<br />
A.1-2.<br />
Table A.1-2: Proton 50-Launch Moving Average<br />
Year Proton Block DM Proton (3 stage) Failures/Cause<br />
1992 8<br />
1993 8 1 booster second and third stage propulsion failure<br />
1994 13<br />
1995 6 1<br />
1996 7 1 1 Block DM engine failure, 1 Mars 96 SC control failure<br />
1997 8 1 Block DM propulsion failure<br />
1998 5 1<br />
Last 50 Proton/Block DM: 46/50 92% (Since 10 Sep 1992)<br />
A.2 PROTON FLIGHT HISTORY<br />
The Proton <strong>launch</strong> <strong>vehicle</strong> is one of the most reliable commercial <strong>launch</strong> <strong>vehicle</strong>s available today, with a 92 percent<br />
reliability record over the last 50 <strong>launch</strong>es.<br />
Summary <strong>launch</strong> data by year are shown in Table A.2-1.<br />
Page A-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.2-1: Proton Launch Record Summary (1970-1998)<br />
Year Number of<br />
Launches<br />
Number of Launches by Version Total Launches on<br />
Accrual Basis<br />
4-Stage Version 3-Stage Version<br />
Failures<br />
1970 7 4 3 7 1 Proton K/Block DM<br />
& 1 Proton K<br />
1971 7 5 2 14<br />
1972 2 1 1 16 1 Proton K<br />
1973 7 5 2 23<br />
1974 6 4 2 29<br />
1975 5 5 - 34 1 Proton K/Block DM<br />
1976 5 3 2 39<br />
1977 6 2 2 45 1 Proton K<br />
1978 8 7 1 53 3 Proton K/Block<br />
DM’s<br />
1979 6 5 1 59<br />
1980 5 5 - 64<br />
1981 7 6 1 71<br />
1982 10 9 1 81 2 Proton K/Block<br />
DM’s<br />
1983 12 11 1 93<br />
1984 13 13 _ 106<br />
1985 10 9 1 116<br />
1986 8 7 1 124 1 Proton K<br />
1987 13 11 2 137 2 Proton K/Block<br />
DM’s<br />
1988 12 13 _ 150 2 Proton K/Block<br />
DM’s<br />
1989 11 10 1 161<br />
1990 11 10 1 172 1 Proton K/Block DM<br />
1991 9 8 1 181<br />
1992 8 8 _ 189<br />
1993 7 7 _ 196 1 Proton K/Block DM<br />
1994 13 13 _ 209<br />
1995 7 6 1 216<br />
1996 8 7 1 224 2 Proton K/Block<br />
DM’s<br />
1997 8 8 - 232 1 Proton K/Block DM<br />
1998 6 5 1 238<br />
Page A-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
A.3 DETAILED FLIGHT HISTORY<br />
The Proton <strong>launch</strong> history since 1970 is shown in Table A.3-1. This historical data was verified by Khrunichev in<br />
December 1993, and subsequent <strong>launch</strong>es confirmed on a case by case basis.<br />
Table A.3-1: Proton Launch History<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
6 Feb 1970 w Cosmos Failed to orbit Command abort<br />
8 Aug 1970 w Unknown Failed to orbit No data<br />
12 Sep 1970 w Luna-16 Escape<br />
20 Oct 1970 w Zond-8 Escape<br />
10 Nov 1970 w Luna-17 Escape<br />
24 Nov 1970 w Cosmos-379 192 km x 1210 km at 51.9 deg<br />
2 Dec 1970 w Cosmos-382 2464 km x 5189 km at 51.9 deg<br />
26 Feb 1971 w Cosmos-398 201 km x 7250 km at 51.6 deg<br />
19 Apr 1971 w Salyut-1 200 km x 210 km at 51.6 deg<br />
10 May 1971 w Cosmos-419 145 km x 159 km at 51.5 deg<br />
19 May 1971 w Mars-2 Escape<br />
28 May 1971 w Mars-3 Escape<br />
2 Sep 1971 w Luna-18 Escape<br />
28 Sep 1971 w Luna-19 Escape<br />
14 Feb 1972 w Luna-20 Escape<br />
29 Jul 1972 w Salyut Failed to orbit<br />
8 Jan 1973 w Luna-21 Escape<br />
3 Apr 1973 w Salyut-2 207 km x 248 km at 51.6 deg No data<br />
11 May 1973 w Cosmos-557 214 km x 243 km at 51.6 deg<br />
21 Jul 1973 w Mars-4 Escape<br />
25 Jul 1973 w Mars-5 Escape<br />
5 Aug 1973 w Mars-6 Escape<br />
9 Aug 1973 w Mars-7 Escape<br />
26 Mar 1974 w Cosmos-637 LEO<br />
29 May 1974 w Luna-22 Escape<br />
24 Jun 1974 w Salyut-3 LEO<br />
29 Jul 1974 w Molniya-1S Elliptical orbit<br />
28 Oct 1974 w Luna-23 Escape<br />
26 Dec 1974 w Salyut-4 LEO<br />
6 Jun 1975 w Venera-9 Earth escape<br />
14 Jun 1975 w Venera-10 Earth escape<br />
8 Oct 1975 w Cosmos-775 LEO<br />
16 Oct 1975 w Luna Escape<br />
22 Dec 1975 w Raduga-1 GSO<br />
22 Jun 1976 w Salyut-5 LEO<br />
9 Aug 1976 w Luna-24 Escape<br />
11 Sep 1976 w Raduga-2 GSO<br />
26 Oct 1976 w Ekran-1 GSO<br />
15 Dec 1976 w Cosmos-881 and 882 LEO<br />
Page A-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.3-1a: Proton Launch History (Continued)<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
17 Jul 1977 w Cosmos-929 301 km x 308 km at 51.5 deg<br />
23 Jul 1977 w Raduga-3 GSO<br />
4 Aug 1977 F Cosmos Failed to orbit<br />
20 Sep 1977 w Ekran-2 GSO<br />
29 Sep 1977 w Salyut-6 380 km x 391 km at 51.6 deg<br />
14 Oct 1977 F Unknown Failed to orbit<br />
30 Mar 1978 Cosmos-997 and 998 230 km x 200 km at 51.6 deg<br />
27 May 1978 F Ekran Failed to orbit First stage failure<br />
18 Jul 1978 w Raduga-4 GSO<br />
17 Aug 1978 F Ekran Failed to orbit Second stage failure<br />
9 Sep 1978 w Venera-l l Escape<br />
14 Sep 1978 w Venera-12 Escape<br />
17 Oct 1978 F Ekran Failed to orbit Second stage failure<br />
19 Dec 1978 w Gorizont-1 20,600 km x 50,960 km at 14.3 deg Failure?<br />
21 Feb 1979 w Ekran-3 GSO<br />
25 Apr 1979 w Raduga-5 GSO<br />
22 May 1979 w Cosmos- 1100 and 1101 193 km x 223 km at 51.6 deg<br />
5 Jul 1979 w Gorizont-2 GSO<br />
3 Oct 1979 w Ekran-4 GSO<br />
28 Dec 1979 w Gorizont-3 GSO<br />
2 Feb 1980 w Raduga-6 GSO<br />
14 Jun 1980 w Gorizont-4 GSO<br />
15 Jul 1980 w Ekran-5 GSO<br />
5 Oct 1980 w Raduga-7 GSO<br />
26 Dec 1980 w Ekran-6 GSO<br />
18 Mar 1981 w Raduga-8 GSO<br />
25 Apr 1981 w Cosmos-1267 240 km x 278 km at 51.5 deg<br />
26 Jun 1981 w Ekran-7 GSO<br />
30 Jul 1981 w Raduga-9 GSO<br />
9 Oct 1981 w Raduga-10 GSO<br />
30 Oct 1981 w Venera-13 Escape<br />
4 Nov 1981 w Venera-14 Escape<br />
5 Feb 1982 w Ekran-8 GSO<br />
15 Mar 1982 w Gorizont-5 GSO<br />
19 Apr 1982 w Salyut-7 473 km x 474 km at 51.6 deg<br />
17 May 1982 w Cosmos-1366 GSO<br />
23 Jul 1982 F Ekran Failed to orbit First stage failure<br />
16 Sep 1982 w Ekran-9 GSO<br />
12 Oct 1982 w Cosmos-1413 and 1415 19,000 km x 19,000 km at 64.7 deg<br />
20 Oct 1982 w Gorizont-6 GSO<br />
26 Nov 1982 w Raduga-11 GSO<br />
24 Dec 1982 F Raduga Failed to orbit Second stage failure<br />
Page A-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.3-1b: Proton Launch History (Continued)<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
2 Mar 1983 w Cosmos-1443 324 km x 327 km at 51.6 deg<br />
12 Mar 1983 w Ekran-10 GSO<br />
23 Mar 1983 w Astron-1 1,950 km x 201,100 km at 51.09 deg<br />
8 Apr 1983 w Raduga-12 GSO<br />
2 Jun 1983 w Venera-15 Escape<br />
6 Jun 1983 w Venera-16 Escape<br />
1 Jul 1983 w Gorizont-7 GSO<br />
10 Aug 1983 w Cosmos-1490 and 1492 19,000 km x 19,000 km at 64.8 deg<br />
25 Aug 1983 w Raduga-13 GSO<br />
29 Sep 1983 w Ekran-II GSO<br />
30 Nov 1983 w Gorizont-8 GSO<br />
29 Dec 1983 w Cosmos-1519 and 1521 19,000 km x 19,000 km at 64.8 deg<br />
15 Feb 1984 w Raduga-14 GSO<br />
2 Mar 1984 w Cosmos-1540 GSO<br />
16 Mar 1984 w Ekran-12 GSO<br />
29 Mar 1984 w Cosmos-1546 GSO<br />
22 Apr 1984 w Gorizont-9 GSO<br />
19 May 1984 w Cosmos-1554 and 1556 19,000 km x 19,000 km at 64.8 deg<br />
22 Jun 1984 w Raduga-15 GSO<br />
1 Aug 1984 Gorizont-10 GSO<br />
24 Aug 1984 w Ekran-13 GSO<br />
4 Sep 1984 w Cosmos-1593 and 1595 19,000 km x 19,000 km at 64.8 deg<br />
28 Sep 1984 w Cosmos-1603 836 km x 864 km at 71 deg<br />
15 Dec 1984 w Vega-1 Escape<br />
21 Dec 1984 w Vega-2 Escape<br />
18 Jan 1985 w Gorizont-ll GSO<br />
21 Feb 1985 w Cosmos-1629 GSO<br />
22 Mar 1985 w Ekran-14 GSO<br />
17 May 1985 w Cosmos-1650 and 1652 19,000 km x 19,000 km at 64.8 deg<br />
30 May 1985 w Cosmos-1656 800 km x 860 km at 71.1 deg<br />
8 Aug 1985 w Raduga-16 GSO<br />
27 Sep 1985 w Cosmos-1686 291 km x 312 km at 51.6 deg<br />
25 Oct 1985 w Cosmos-1700 GSO<br />
15 Nov 1985 w Raduga-17 GSO<br />
24 Dec 1985 w Cosmos-1710 and 1712 19,000 km x 19,000 km at 64.8 deg<br />
17 Jan 1986 w Raduga- 18 GSO<br />
19 Feb 1986 w Mir 335 km x 358 km at 51.6 deg<br />
4 Apr 1986 w Cosmos-1738 GSO<br />
24 May 1986 w Ekran-15 GSO<br />
10 Jun 1986 Gorizont-12 GSO<br />
Page A-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.3-1c: Proton Launch History (Continued)<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
16 Sep 1986 w Cosmos-1778 and 1780 19,000 km x 19,000 km at 64.8 deg<br />
25 Oct 1986 w Raduga-I9 GSO<br />
18 Nov 1986 w Gorizont-13 GSO<br />
30 Jan 1987 F Cosmos-1817 188 km x 210 km at 51.6 deg Fourth stage failed to<br />
ignite<br />
19 Mar 1987 w Raduga-20 GSO<br />
31 Mar 1987 w Kvant-1 298 km x 344 km at 51.6 deg<br />
24 Apr 1987 F Cosmos- 1838 to 1840 200 km x 17,000 km at 64.9 deg Fourth stage early<br />
shutdown<br />
11 May 1987 w Gorizont-14 GSO<br />
25 Jul 1987 Cosmos-1870 237 km x 249 km at 71.9 deg<br />
3 Sep 1987 w Ekran-16 GSO<br />
16 Sep 1987 w Cosmos-1883 and 1885 19,000 km x 19,000 km at 64.8 deg<br />
1 Oct 1987 w Cosmos-1888 GSO<br />
28 Oct 1987 w Cosmos-1894 GSO<br />
26 Nov 1987 w Cosmos-1897 GSO<br />
10 Dec 1987 w Raduga-21 GSO<br />
27 Dec 1987 w Ekran-17 GSO<br />
17 Feb 1988 F Cosmos-1917P1919 162 km x 170 km at 64.8 deg Fourth stage did not ignite<br />
31 Mar 1988 w Gorizont-15 GSO<br />
26 Apr 1988 w Cosmos-1940 GSO<br />
6 May 1988 w Ekran-18 GSO<br />
21 May 1988 w Cosmos 1946-1948 19,000 km x19,000 km at 64.9 deg<br />
7 Jul 1988 w Phobos-1 Escape<br />
12 Jul 1988 w Phobos-2 Escape<br />
1 Aug 1988 w Cosmos-1961 GSO<br />
18 Aug 1988 w Gorizont-16 GSO<br />
16 Sep 1988 w Cosmos-1970P1972 19,000 km x 19,000 km at 64.8 deg<br />
20 Oct 1988 w Raduga-22 GSO<br />
10 Dec 1988 w Ekran-I9 GSO<br />
10 Jan 1989 w Cosmos-1987P1989 19,000 km x 19,000 km at 64.9 deg<br />
26 Jan 1989 w Gorizont-17 GSO<br />
14 Apr 1989 w Raduga-23 GSO<br />
31 May 1989 w Cosmos-2022P2024 19,000 km x 19,000 km at 64.8 deg<br />
21 Jun 1989 w Raduga-l-1 GSO<br />
5 Jul 1989 w Gorizont-18 GSO<br />
28 Sep 1989 w Gorizont- 19 GSO<br />
26 Nov 1989 w Kvant-2 215 km x 321 km at 51.6 deg<br />
1 Dec 1989 w Granat 1957 km x 201,700 km at 52.1 deg<br />
15 Dec 1989 Raduga-24 GSO<br />
27 Dec 1989 w Cosmos-2054 Unknown<br />
Page A-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.3-1d: Proton Launch History (Continued)<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
15 Feb 1990 w Raduga-25 GSO<br />
19 May 1990 w Cosmos-2079P81 19,000 km x19,000 km at 65 deg<br />
31 May 1990 w Kristall 383 km x 481 km at 51.6 deg<br />
20Jun 1990 w Gorizont-20 GSO<br />
18 Jul 1990 w Cosmos-2085 GSO<br />
9 Aug 1990 F Unknown Did not achieve orbit<br />
3 Nov 1990 w Gorizont-21 GSO<br />
23 Nov 1990 w Gorizont-22 GSO<br />
8 Dec 1990 w Cosmos-2109P11 19,000 km x 19,000 km at 64.8 deg<br />
20 Dec 1990 w Raduga-26 GSO<br />
27 Dec 1990 w Raduga-26 GSO<br />
14 Feb 1991 w Cosmos-2133 GSO<br />
28 Feb 1991 w Raduga-27 GSO<br />
31 Mar 1991 w Almaz-1 268 km x 281 km at 72.7 deg<br />
4 Apr 1991 w Cosmos-2139P41 19,000 km x 19,000 km at 64.9 deg<br />
1 Jul 1991 w Gorizont-23 GSO<br />
13 Sep 1991 w Cosmos-2155 GSO<br />
23 Oct 1991 w Gorizont-24 GSO<br />
22 Nov 1991 w Cosmos-2172 GSO<br />
19 Dec 1991 w Raduga-28 GSO<br />
29 Jan 1992 w Cosmos-2177P79 19,000 km x 19,000 km at 64.8 deg<br />
2 Apr 1992 w Gorizon-25 GSO<br />
14Ju11992 w Gorizont-26 GSO<br />
30 Jul 1992 w Cosmos-2204-06 19,000 km x 19,000 km at 64.8 deg<br />
10 Sep 1992 w Cosmos-2209 GSO<br />
30 Oct 1992 w Ekran-20 GSO<br />
27Nov 1992 w Gorizont-27 GSO<br />
17 Dec 1992 w Cosmos-2224 GSO<br />
17 Feb 1993 w Cosmos-223?P3? 19,000 km x 19,000 km at 64.8 deg<br />
17 Mar 1993 w Raduga-29 GSO<br />
27 May 1993 F Gorizont Did not achieve orbit 2 nd and 3 rd stage<br />
propulsion failure<br />
30 Sep 1993 w Gorizont GSO<br />
28 Oct 1993 w Gorizont GSO<br />
18 Nov 1993 w Gorizont GSO<br />
23 Dec 1993 w Gorizont GSO<br />
20 Jan 1994 w GALS GSO<br />
5 Feb 1994 w Raduga - 30 GSO<br />
18 Feb 1994 w Raduga- 31 GSO<br />
11 Apr 1994 w Glonass l9,000 km x l9,000 km at64.8°<br />
20 May 1994 w Gorizant GSO<br />
Page A-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table A.3-1e: Proton Launch History (Continued)<br />
Date Proton Variant Payload Orbit Type Comments<br />
4-stage 3-stage<br />
7 Jul 1994 w Cosmos GSO<br />
11 Aug 1994 w Glonass 19,000 km x 19,000 km at 64.8°<br />
21 Sep 1994 w Cosmos - 2291 GSO<br />
13 Oct 1994 w Express GSO<br />
31 Oct 1994 w Electro GSO<br />
20 Nov 1994 w Glonass 19,000 km x 19,999 km at 64.8°<br />
16 Dec 1994 w Luch GSO<br />
28 Dec 1994 w F Raduga - 32 GSO<br />
7 Mar 1995 w Glonass 19,000 km x 19,000 km at 64.8°<br />
20 May 1995 w Spektr 335 km x 358 km at 51.6°<br />
24 Jul 1995 w Glonass 19,000 km x 19,000 km at 64.8°<br />
31 Aug 1995 w Gazer GSO<br />
11 Oct 1995 w Looch - 1 GSO<br />
17 Nov 1995 w GALS GSO<br />
14 Dec l995 w F Glonass 19,140 km x l9,l00 km at 64.8°<br />
25 Jan 1996 w Gorizant GSO<br />
19 Feb 1996 F Raduga GSO Block DM propulsion<br />
failure<br />
9 Apr 1996 w Astra 1F Hi-GTO Commercial<br />
23 Apr 1996 w Priroda 214 km x 328 km at 51.6 deg<br />
25 May 1996 w Gorizant GSO<br />
6 Sep 1996 w Inmarsat 3 F2 GSO Commercial<br />
26 Sep 1996 w Express GSO<br />
16 Nov 1996 F Mars 96 Did not achieve escape trajectory Failure of Mars 96 control<br />
system to initiate Block<br />
D2 engine ignition<br />
24 May 1997 w Telstar-5 Hi-GTO Commercial<br />
6 June 1997 w Arak GSO<br />
18 June 1997 w Iridium LEO Commercial<br />
14 Aug 1997 w Cosmos-2345 GSO<br />
28 Aug 1997 w PanAmSat-5 Hi-GTO Commercial<br />
15 Sep 1997 w Iridium LEO Commercial<br />
3 Dec 1997 w Astra-1G Hi-GTO Commercial<br />
25 Dec 1997 F AsiaSat-3 GTO Block DM Engine Failure<br />
7 Apr 1998 w Iridium LEO Commercial<br />
29 Apr 1998 w Cosmos-2350 GSO<br />
8 May 1998 w Echostar-IV Hi-GTO Commercial<br />
30 Aug 1998 w Astra 2A Hi-GTO Commercial<br />
04 Nov 1998 w PanAmSat-8 Hi-GTO Commercial<br />
20 Nov 1998 w Zarya (FGB) LEO RSA/NASA<br />
Page A-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Notes:<br />
a) The SL-12 <strong>launch</strong> <strong>vehicle</strong> is also designated D-l-e and is the four-stage version of the Proton.<br />
b) The SL-13 <strong>launch</strong> <strong>vehicle</strong> is also designated D-1 and is the three-stage version of the Proton.<br />
c) The stated orbital parameters are approximate and included for reference only.<br />
A.4 FAILURES CAUSES AND CORRECTIVE ACTION<br />
Data provided below was provided by Khrunichev Space Center.<br />
a) 1970: After 128.3 seconds of flight, 1 st stage engine cutoff due to false alarm from the <strong>launch</strong> <strong>vehicle</strong> safety system<br />
activated by the engine pressure gage. Manufacturing defect. Additional check of gages introduced at point of<br />
installation.<br />
b) 1972: After 181.9 seconds of flight, 2 nd stage automated stabilization system failure due to a relay short circuit in<br />
the "pitch" and "yawing" channels caused by elastic deformation of the device housing (which operates in vacuum).<br />
Design defect. Design of instruments upgraded and additional testing undertaken.<br />
c) 1975: Failure of 4 th stage oxidizer booster pump. Manufacturing/design defect. Cryogen-helium condensate<br />
freezing. Booster pump blowing introduced.<br />
d) 1977: After 40.13 seconds of flight, spontaneous cutoff of 1 st stage steering motor, loss of stability and engine<br />
cutoff at 53.68 seconds into the flight safety system command. Steering most failure due to spool-an-sleeve pair<br />
manufacturing defect (faulty liner) which caused penetration of hard particles under liner rim and resulted in<br />
spool-and-sleeve seizure.<br />
e) 1978: After 87 seconds of flight, loss of stability commenced due to error of 1st stage second combustion chamber<br />
steering gear. High temperature impact on cables due to heptyl leak into second block engine compartment. Leak<br />
likely developed at heptyl feed coupling to gas generator. Coupling upgraded.<br />
f) 1978: Flight terminated after 259.1 seconds due to loss of <strong>launch</strong> <strong>vehicle</strong> stability. Automatic stabilization system<br />
electric circuit failure in rear compartment of 2nd stage caused by hot gases leaking from second engine gas inlet<br />
due to faulty sealing of pressure gage. Gage attaching point upgraded.<br />
g) 1978: After 235.62 seconds of flight, 2 nd stage engine shutoff and loss of stability caused by a turbine part igniting<br />
in turbo pump gas tract followed by gas inlet destruction and hot air ejection into 2 nd rear section. Engine design<br />
upgraded.<br />
h) 1982: At 45.15 seconds into the flight, major malfunctioning of 1 st stage engine fifth chamber. Flight terminated<br />
by <strong>launch</strong> <strong>vehicle</strong> safety system command. Failure caused by steering motor malfunctioning: first stage of<br />
hydraulic booster got out of balance coupled with booster dynamic excitation at resonance frequencies. Hydraulic<br />
booster design redefined.<br />
i) 1982: 2 nd stage engine failure caused by high-frequency vibrations. Engine design upgraded.<br />
j) 1986: Control system failure due to brief relay contact separation caused by engine vibration. Upgrading included<br />
introduction of self-latching action capability for program power distributor shaft.<br />
k) 1987: 4 th stage control system failure due to component (relay) defect. Manufacturing defect. Remedial program<br />
introduced at supplier's factory. Inspection made more stringent.<br />
l) 1987: 4 th stage control system failure due to control system instrument defect. Manufacturing defect. Device<br />
manufactured at the time of transfer from developer's pilot production to a factory for full scale production.<br />
Remedial program introduced at relevant factory. No recurring failures recorded.<br />
Page A-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
m) 1988: 3 rd stage engine failure caused by destruction of fuel line leading to mixer. Unique manufacturing defect.<br />
Inventory rechecked.<br />
n) 1988: 4 th stage engine failure due to temperature rise in combustion chamber caused by penetration of foreign<br />
particles from the fuel tank. Manufacturing defect. Remedial program introduced at point of manufacture to<br />
prevent penetration of foreign particles into tanks. No recurring failures recorded.<br />
o) 1990: 2 nd stage engine shutoff due to termination of oxidizer supply. Fuel line clogged by a piece of textile<br />
(wiping rag). Remedial program introduced to prevent wiping rags from being left inside engine and <strong>launch</strong><br />
<strong>vehicle</strong>.<br />
p) 1993: 2 nd and 3 rd stage engine failures. Multiple engine combustion chamber burn-through caused by propellant<br />
contaminants. Remedial program introduced to modify propellant specifications and testing procedures. All<br />
<strong>launch</strong> site propellant storage, transfer, and handling equipment purged and cleaned.<br />
q) 1996: Block DM 4th stage second burn ignition failure. Remedial program involved corrective actions to prevent<br />
two possible causes. The first involved introduction of redundant lockers, revised installation procedures, and<br />
increased factory inspections to prevent a loosening of a tube joint causing a leak that would prevent engine<br />
ignition. The second involved additional contamination control procedures to further precule particulate<br />
contamination of the hypergolic start system.<br />
r) 1996: Block D2 4 th stage engine failure during second burn due to malfunction of Mars 96 spacecraft control<br />
system, and associated improper engine command sequences. Unique configuration of spacecraft and 4 th stage.<br />
Remedial program includes stringent adherence to established integration and test procedures.<br />
s) 1997: Block DM 4 th stage engine failure resulting from improperly coated turbopump seal. Remedial program<br />
includes removal of unnecessary (for < 4 burn <strong>mission</strong>s) coating.<br />
Page A-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
B1. QUALITY SYSTEM<br />
TBS<br />
Page B1-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Introduction<br />
This section describes the data required from the spacecraft customer to determine the compatibility of the spacecraft<br />
with the Proton <strong>launch</strong> <strong>vehicle</strong>. Providing this data in full constitutes providing an Interface Requirements Document<br />
which is a contractual document provided by the Customer at the beginning of a <strong>mission</strong> integration cycle. For<br />
preliminary feasibility studies, a smaller subset of this information can be provided as indicated in the table.<br />
The requested information is provided in the sequence the data appears in the Interface Control Document in order to<br />
simplify the creation of this document once a contract is signed.<br />
C1. GENERAL INFORMATION<br />
Item Parameter Units Feas. Study<br />
Spacecraft Name<br />
Manufacturer<br />
Isometric view, <strong>launch</strong> configuration<br />
Isometric view, on-orbit configuration<br />
Estimated <strong>launch</strong> date<br />
Page C1-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C2. INPUT DOCUMENTS<br />
Item Parameter Units Feas. Study<br />
Spacecraft Dynamic<br />
Model<br />
Spacecraft Thermal<br />
Model<br />
Spacecraft Physical<br />
Model<br />
Per specification LKES-9704-0207<br />
Per specification LKET 9704-0206<br />
Per specification LKEB 9801-039<br />
Launch Operations Plan Reference Section 5.3.5 of this Mission Planner’s Guide<br />
Environmental Test<br />
Plan<br />
SC Interface Control<br />
Drawing<br />
Reference Section 5.3.3 of this Mission Planner’s Guide<br />
Drawing based on physical model above which shows:<br />
Locations of all critical surfaces (surfaces closer than 50 mm from<br />
useable volume envelope). Show on drawing and in table in SC<br />
coordinates x,y and z.<br />
mm<br />
Major SC dimensions while in <strong>launch</strong> configuration mm<br />
Complete description of any points OUTSIDE of the useable volume<br />
Complete dimensions of interface ring with structural definition up to<br />
125 mm above interface plane including stiffness characteristics<br />
Umbilical connector definition including position relative to<br />
separation plane<br />
Pneumatic connector interfaces (if applicable)<br />
Access envelopes for accessing SC and description of how access is to<br />
be made<br />
Drawing to show handling fixtures and how attached to SC (hoists,<br />
slings, crossbars)<br />
Page C2-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3. INTERFACES<br />
C3.1 MECHANICAL INTERFACES<br />
Item Parameter Units Feas. Study<br />
Adapter system Specify which standard adapter system. Reference Appendix D of this<br />
Mission Planner’s Guide.<br />
# pushoff springs (if known). Reference Appendix D of this Mission<br />
Planner’s Guide.<br />
Specific requirements for mechanical interfaces not provided by above<br />
standard adapter system<br />
Coordinate system Drawing showing spacecraft coordinate system<br />
SC Logo Provide design<br />
SC interface flange Cross section properties<br />
I xx<br />
Drawing showing desired orientation of spacecraft coordinate system<br />
relative to adapter coordinate system<br />
Description of constraints on spacecraft orientation<br />
I yy<br />
L s = 25 mm<br />
S<br />
Scribe mark location<br />
SC stiffness Minimum fundamental lateral and axial mode frequencies (must be<br />
greater than 10 and 25 Hz respectively<br />
SC interface loads Confirm SC lifting device and structure can lift SC+adapter mass =<br />
200 kg<br />
SC mass properties Fill in tables in attached Table C3.1-1 and C3.1-2.<br />
mm 4<br />
mm 4<br />
mm<br />
mm 2<br />
Hz<br />
Yes or no<br />
Fairing access doors Location required for access SC<br />
coordinates<br />
Method of access required<br />
Time when access required<br />
Fitcheck/ shocktest Confirm fitcheck and shock test requirements<br />
SC Pendulum model Provide pendulum model of SC during powered flight per<br />
Figure C3.1-1<br />
SC Slosh model Provide slosh model of SC during ballistic flight and at separation by<br />
providing parameters in Figure C3.1-2.<br />
Propellant tank Provide general propellant tank geometry per Figure C3.1-3<br />
Page C3-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3.2 ELECTRICAL INTERFACES<br />
Item Parameter Units Feas. Study<br />
Electrical connector Confirm type of connectors<br />
LVIJ1<br />
LVIJ2<br />
Confirm location and keying<br />
LVIJ1<br />
LVIJ2<br />
Confirm quantities to be supplied<br />
LVIJ1<br />
LVIJ2<br />
LVIP1<br />
LVIP2<br />
Provide pin locations of SC separation jumper loops<br />
LVIJ1 pins:<br />
LVIJ2 pins:<br />
Current flow 20 seconds or less prior to separation. mA (must<br />
be less<br />
than 100)<br />
Dry loop commands Dry loop commands required? Yes/no<br />
Characteristics if yes:<br />
# commands<br />
Desired pin locations<br />
LVIJ1 pins:<br />
LVIJ2 pins:<br />
Current<br />
Voltage<br />
Pulse duration<br />
Time during flight<br />
mA<br />
V<br />
ms<br />
sec prior<br />
to<br />
separation<br />
SC telemetry processing LV processing of SC telemetry required? Yes/no<br />
Characteristics of if yes:<br />
Desired pin locations<br />
LVIJ1 pins:<br />
Current through<br />
umbilicals at liftoff<br />
LVIJ2 pins:<br />
Voltage<br />
Current<br />
Data rate<br />
V<br />
mA<br />
Hz<br />
Current flow 5 minutes or less prior to liftoff. mA (must<br />
be less<br />
than 100)<br />
SC RF characteristics Fill in table in attached Table C3.2-1.<br />
SC omni antenna Provide location of omni antennas to be used on pad, including:<br />
Antenna pattern showing antenna origin in SC coordinates, -3db<br />
beamwidth<br />
Umbilical wiring Fill out Table C3.2-2 with pin assignments and desired line<br />
characteristics<br />
Page C3-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3.3 ENVIRONMENTAL INTERFACES<br />
Item Parameter Units Feas. Study<br />
Thermal requirements Provide any particular ground thermal requirements<br />
Provide any particular thermal requirements during ascent or parking<br />
orbit<br />
Provide any particular thermal requirements during transfer orbit<br />
including:<br />
sun angle vs. time<br />
Provide any particular thermal requirements during final injection<br />
orbit<br />
Venting analysis Provide archimedes volume of SC in <strong>launch</strong> configuration (for<br />
venting analysis)<br />
SC Testing Provide confirmation of compliance with Planner’s Guide test<br />
requirements for acoustic, sine, static testing and for the<br />
fitcheck/shocktest<br />
EMC Provide confirmation of compliance with EM Susceptibility curve in<br />
LV in Planner’s Guide<br />
Provide confirmation of acceptability of LV radiated e<strong>mission</strong>s curve<br />
in Planner’s Guide<br />
Humidity Provide humidity requirements for ground transportation<br />
Provide humidity requirements for processing facility<br />
Page C3-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3.4 FLIGHT DESIGN<br />
Item Parameter Units Feas. Study<br />
Parking Orbit Define thermal conditioning maneuvers required<br />
Define sun angle constraints<br />
Transfer Orbit Define thermal conditioning maneuvers required<br />
Define sun angle constraints<br />
Injection Orbit Target orbit for performance calculation:<br />
Injection mass<br />
Perigee<br />
Apogee<br />
Inclination<br />
Argument of Perigee<br />
Longitude of Ascending Node<br />
Launch Window Define constraints on <strong>launch</strong> window<br />
Provide target <strong>launch</strong> date<br />
Provide <strong>launch</strong> window at perigee passage for one year covering the<br />
contractual <strong>launch</strong> date. Include open and close times in GMT for<br />
each day.<br />
Separation Define type of separation 3 axis<br />
stabilized,<br />
spinning<br />
or<br />
transverse<br />
spin<br />
Define desired separation attitude with a diagram showing the<br />
pointing vector in SC coordinates and the pointing vector relative to<br />
relative right ascension and declination as defined in the Planner’s<br />
Guide.<br />
Define desired spin rate about each SC axis.<br />
Define desired spin rate tolerance about each SC axis<br />
kg<br />
km<br />
km<br />
deg.<br />
deg.<br />
deg.<br />
deg.<br />
deg.<br />
Define desired separation velocity m/sec.<br />
Page C3-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3.5 OPERATIONS<br />
Item Parameter Units Feas. Study<br />
EGSE Fill out table in Table C3.5-1 for all EGSE to be used at Launch Site<br />
Fluid and gases Fill out table in Table C3.5-2 for quantities and types of fluids and<br />
gases<br />
Contamination Control Provide any special requirements<br />
Campaign Support Provide list of support required in each area (if nonstandard):<br />
Area 92A-50<br />
Hall 102<br />
Hall 101<br />
Hall 103A<br />
Control Rm<br />
Hall 103<br />
Offices<br />
Area 31<br />
Rm 100C<br />
Rm 100B<br />
Rm 100A<br />
Rm 119<br />
Offices<br />
Launch Complex 81<br />
Bunker<br />
Vault<br />
Pad<br />
Bldg 92-1<br />
Page C3-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
C3.6 OPERATIONS (Continued)<br />
Item Parameter Units Feas. Study<br />
Transportation Provide description of all items including propellant to be shipped<br />
including:<br />
Item name<br />
Quantity<br />
Weight<br />
Tiedown method<br />
Storage requirements<br />
Handling Provide description of items which require physical handling at the<br />
<strong>launch</strong> site including:<br />
Equipment name/location<br />
Dimensions<br />
Weight<br />
Handling method<br />
Communications Define number and type of international lines required. Up to 3 64<br />
kbps lines can be provided.<br />
Ground Electrical<br />
Interfaces<br />
If multiplexer provided, provide characteristics of and desired<br />
location.<br />
Provide locations of all intersite data trans<strong>mission</strong>s including data type<br />
and rate and interface (analog modem, V.35 or RS232 interface)<br />
Provide requirements for hardline data trans<strong>mission</strong> between Area 81<br />
Vault and Bunker, Bldg 44 Rm 58 and Bldg 40D Rm 120 and/or<br />
between Bldg 40D Rm 301 and Room 120.<br />
Provide block diagrams of desired umbilical interfaces between SC<br />
and EGSE while being processed and while on the pad<br />
Feedthroughs Provide description of feedthroughs required between Control Rm of<br />
92A-50 and Hall 103A or between Rm 58 of Bldg 44 and Hall 1,<br />
including:<br />
Feedthrough designation<br />
Cable designation<br />
Cable connector dia.<br />
Cable dia.<br />
kg<br />
m<br />
kg<br />
mm<br />
mm<br />
Page C3-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.1-1: SC Mass properties<br />
SC Mass Properties are shown with a normal distribution. Table C3.1-1c provides the dry spacecraft mass properties to<br />
be used for analysis of separation dynamics taking into account fluid sloshing effects.<br />
Approximately TBD kg of helium gas pressurant is included in the full-up spacecraft mass.<br />
Table C3.1-1a: SC Mass Properties Near 0g<br />
Mass (kg) Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to spacecraft<br />
c.g.(kg-M 2 )<br />
CGX CGY CGZ IXX IYY IZZ PXY PXZ PYZ<br />
Nominal TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- + TBD + TBD + TBD + TBD TBD% TBD% TBD% TBD% TBD% TBD%<br />
Table C3.1-1b: SC Mass Properties Near 1g<br />
Mass (kg) Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to spacecraft<br />
c.g.(kg-M 2 )<br />
CGX CGY CGZ IXX IYY IZZ PXY PXZ PYZ<br />
Nominal TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- + TBD + TBD + TBD + TBD TBD% TBD% TBD% TBD% TBD% TBD%<br />
Note:<br />
a) Maximum to minimum inertia ratio is greater than or equal to 1.02<br />
b) Z coordinate relative to separation plane<br />
c) Maximum required tolerance on the final weight before <strong>launch</strong> =+0/- TBD kg. and will be based on the SC manufacturers<br />
final mass properties report<br />
d) Above data based on the SC manufacturers Mass Properties Report dated TBD.<br />
Page C3-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.1-1c: SC Mass Properties(Dry Spacecraft)<br />
Mass (kg) Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to spacecraft<br />
c.g.(kg-M 2 )<br />
CGX CGY CGZ IXX IYY IZZ IXY IXZ IYZ<br />
Nominal TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- + TBD + TBD +TBD + TBD TBD % TBD % TBD % TBD% TBD% TBD%<br />
Notes:<br />
a) Z coordinate relative to separation plane<br />
b) Above data based on the SC manufacturers Mass Properties Report dated TBD.<br />
Page C3-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.1-2: Description Of Liquid Masses<br />
These tables provide the mass properties for the individual tanks for the nominal propellant load of TBD% fill fraction<br />
(full tanks) being assumed, in a near 0 g field in a 1 g field and during transportation.<br />
The associated tank geometry is shown in Figure C3.1-3.<br />
a) near 0 g (0.125 g)<br />
Mass<br />
(kg)<br />
Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to propellant c.g.<br />
(kg-M 2 )<br />
CGX CGY CGZ* IXX IYY IZZ IXY IXZ IYZ<br />
Ox TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD% TBD % TBD % TBD % TBD % TBD %<br />
Fuel TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD% TBD % TBD % TBD % TBD % TBD %<br />
b) 1g<br />
Mass<br />
(kg)<br />
Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to propellant c.g.<br />
(kg-M 2 )<br />
CGX CGY CGZ* IXX IYY IZZ IXY IXZ IYZ<br />
Ox TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD %<br />
Fuel TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD %<br />
Note: Z coordinate relative to separation plane<br />
c) 1g (During Transportation)<br />
Mass<br />
(kg)<br />
Center of Gravity location<br />
(spacecraft coordinates, mm)<br />
Moments of inertia relative to propellant c.g.<br />
(kg-M 2 )<br />
CGX CGY CGZ* IXX IYY IZZ IXY IXZ IYZ<br />
Ox TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD %<br />
Fuel TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD<br />
+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD %<br />
Note: Z coordinate relative to separation plane<br />
Page C3-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure C3.1-1: SC Pendulum Model<br />
1. 0g<br />
2. 1g<br />
L<br />
m s<br />
a s<br />
x s<br />
Pendulum Model Parameters:<br />
Tank m s, Kg L, m x s, m a s, m , % m 0, Kg x 0, m<br />
F<br />
O<br />
Tank m s, Kg L, m x s, m a s, m , % m 0, Kg x 0, m<br />
F<br />
O<br />
Filling Parameters: (t = 20 O C)<br />
Tank Mass, Kg Level, m Filling, % Density, Kg/m 3 Kinematic viscosity, m 2 /s<br />
F<br />
O<br />
a s - Pivot point of the pendulum from the tank bottom, m<br />
L - Length of the pendulum, m;<br />
x s - Slosh mass location from the tank bottom, m;<br />
m s - Mass of the pendulum, Kg;<br />
- Damping, %.<br />
m 0 - Mass of fixed liquid, kg;<br />
x 0 - Coordinate of mass m 0 from the tank bottom, m.<br />
Page C3-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure C3.1-2: SC Slosh Properties During Ballistic Flight and at Separation<br />
dω<br />
I g + M g C ×<br />
dt<br />
2<br />
dω d l<br />
- M g C × + M g 2<br />
dt dt<br />
2<br />
d ai<br />
dω<br />
= - a<br />
2<br />
i<br />
dt dt<br />
2<br />
d l<br />
A i = +<br />
2<br />
dt<br />
g<br />
dt<br />
s<br />
∑<br />
2<br />
d l<br />
T - ×<br />
2<br />
dt<br />
ω ( I ω ) ( × F )<br />
g<br />
= F - ×<br />
a<br />
× - 2 ω ×<br />
dt<br />
dω<br />
d i<br />
× r + ×<br />
oi<br />
M = M + m ( 1 − G )<br />
i<br />
M g C = s s r M + mi ( 1 − Gi<br />
)<br />
I = I - m ( 1 − G )<br />
g<br />
s<br />
( r )<br />
x<br />
i<br />
S = S y<br />
oi ⎢ i ⎥<br />
⎣ z<br />
i ⎦<br />
i<br />
i<br />
i<br />
∑<br />
ω ( × C)<br />
i<br />
R oi CT ⋅i<br />
∑<br />
ω M g - F CT ⋅i<br />
i<br />
- ω × ( ω × a i ) - i A + F / mi<br />
ω ( ω × r )<br />
∑ r oi<br />
∑<br />
i<br />
i i ( r ) oi<br />
⎡<br />
⎤<br />
2(<br />
1−<br />
Ki<br />
)<br />
Gi =<br />
1+<br />
( 1−<br />
K )<br />
1<br />
ri=<br />
1 3 1−<br />
K 1<br />
d<br />
i<br />
−<br />
i ÷ 4 R<br />
i<br />
i<br />
oi<br />
S S(<br />
r )<br />
0 z ⎡ i<br />
− y<br />
i ⎤<br />
= − z ⎢ i<br />
0 x<br />
i ⎥<br />
⎣ y<br />
i<br />
− x<br />
i<br />
0 ⎦<br />
oi<br />
MINIMUM<br />
G i CT ⋅i<br />
ai<br />
R = r +<br />
oi oi<br />
G<br />
F CT ⋅ i =<br />
• β i ( a)<br />
β i ( )<br />
3H<br />
i<br />
( )( ) a ∈[<br />
−0.<br />
5;<br />
+ 0.<br />
5]<br />
β i(<br />
a)<br />
= −d0i<br />
d0i<br />
( Ri<br />
− ri<br />
)<br />
β ( a)<br />
i<br />
i<br />
⎧<br />
0<br />
⎨<br />
⎩<br />
− k pr d i − k demV<br />
− k pi shearV<br />
ti<br />
1 − K<br />
K<br />
a i K i<br />
= −<br />
1 − K<br />
⎡1⎤<br />
+ aH<br />
⎢ ⎥<br />
⎢<br />
0<br />
⎥<br />
⎢⎣<br />
0⎥⎦<br />
a i<br />
i<br />
⎡<br />
⎢1<br />
⎢⎣<br />
=<br />
i<br />
−3<br />
⎜<br />
⎛<br />
⎝<br />
i<br />
i<br />
d i •<br />
dt<br />
da<br />
i<br />
V Pi = • d i I V ti<br />
= −V<br />
Pi<br />
i • i<br />
dt<br />
k shear<br />
d<br />
da<br />
d<br />
1 − K<br />
= 170229N<br />
/ m k 16256 N ( m / sec)<br />
shear = k 354. 3 N ( m / sec)<br />
shear =<br />
⎛<br />
⎜<br />
⎟<br />
⎞⎜<br />
i ⎠⎜<br />
⎜<br />
⎝<br />
3H<br />
1 +<br />
4R<br />
r < R<br />
i<br />
i<br />
i<br />
r ≥ R<br />
⎞<br />
⎟<br />
i<br />
⎟<br />
⎟<br />
i ⎠<br />
⎤<br />
⎥<br />
⎥⎦<br />
i<br />
Page C3-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
NOMENCLATURE<br />
ω<br />
the spacecraft angular rate vector;<br />
l the vector determining the position of the coordinate origin fixed in the spacecraft body in the inertial<br />
coordinate system;<br />
C<br />
T<br />
F-<br />
a i<br />
V pi,V ti<br />
M s, I s<br />
R i<br />
r s<br />
m i<br />
r oi<br />
R oi<br />
A i<br />
K i<br />
H i<br />
the vector determining the spacecraft mass center position;<br />
the sum of external moments with respect to the selected origin of coordinates;<br />
the sum of external forces acting on the spacecraft;<br />
position vector of fluid center of mass of i-th tank about center of this tank;<br />
dai the radial and tangential components of velocity of the liquid mass center in the i - tank;<br />
dt<br />
the mass and matrix of inertia of the dry spacecraft (without liquid components);<br />
the i - tank radius;<br />
the vector determining the dry spacecraft mass center position in the i - tank with respect to the origin of<br />
coordinates;<br />
the liquid mass in the i - tank;<br />
the vector drawn from the origin of coordinates to the i - tank center;<br />
the vector determining the liquid mass center position in the i - tank with respect to the origin of coordinates;<br />
the vector of inertial acceleration of the i - tank;<br />
the i - tank filling factor<br />
K i=V filling<br />
V total<br />
the i - tank cylinder length<br />
H i=0 corresponds to spherical tank<br />
where: V filling is the volume of poured fuel component;<br />
V total is the total volume of the tank.<br />
Page C3-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure C3.1-3: Propellant Tank Geometry Required Data<br />
TBD<br />
rTBD<br />
TBD<br />
-X<br />
SC/Upper Stage Separation Plane<br />
sc<br />
TBD<br />
TB<br />
D<br />
TBD<br />
Liquid Level<br />
Page C3-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.2-1: SC RF Characteristics<br />
Transmitter 1 Transmitter 2 Receiver 1 Receiver 2<br />
Description TM Transmitter 1 TM Transmitter 2 Command Receiver<br />
1A<br />
Carrier Frequency (Ghz) TBD TBD TBD TBD<br />
3 db Bandwidth (Mhz) TBD TBD TBD TBD<br />
20 db Bandwidth (Mhz) [TBD] [TBD] TBS TBS<br />
60 db Bandwidth (Mhz) TBD TBD TBD TBD<br />
Modulation type and<br />
characteristics<br />
TBD TBD TBD TBD<br />
Intermediate frequency (Mhz) TBD TBD<br />
Local Oscillator frequencies<br />
(Ghz)<br />
Transmit antenna output power<br />
EIRP (dbW)<br />
Max (on-orbit setting)<br />
Max (normal on pad setting)<br />
Nom<br />
Min<br />
Receive antenna receive flux<br />
density (dbW/M 2 )<br />
Max<br />
Nom<br />
Min<br />
Antenna description, polarization,<br />
pattern<br />
Antenna location<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD TBD<br />
TBD<br />
TBD<br />
TBD<br />
Command Receiver<br />
1B<br />
TBD<br />
TBD<br />
TBD<br />
TBD(c) TBD(c) TBD(c) TBD(c)<br />
Operating on pad? TBD (e) TBD (e) TBD TBD<br />
Operating in flight? TBD TBD TBD TBD<br />
Ground equipment output power<br />
(dbm)<br />
Max<br />
Nom<br />
Min<br />
Ground equipment receive power<br />
(dbm)<br />
Max (see note f)<br />
Nom<br />
Min<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
TBD<br />
Page C3-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Notes: [Sample]<br />
a) TBD RF command links and TBD RF telemetry links are required. The SC Checkout Station will have 1 physical command<br />
interface and 1 physical telemetry interface.<br />
b) Transmit and receive frequencies at the RF Console are the same as those indicated in the above table.<br />
c) Vertical polarization (E-Field parallel to spacecraft Y axis), H: horizontal polarization (E-Field parallel to spacecraft X axis).<br />
d) SC-TLM amplifiers are TBD.<br />
e) SC Telemetry Transmitters are on and amplifiers are off during all on-pad operations.<br />
f) Ground equipment can accommodate TBD dBm power levels without damage.<br />
g) SC GSE RF interface impedance shall be 50 ohm<br />
h) Uninterrupted operation of RF devices shall not exceed 8 hours, with a 30 minute break before the next 8 hour session.<br />
i) The SC manufacturer shall provide to Khrunichev the measured coefficient values for TT &C signals via the RF window<br />
obtained during the RF channel checkup in the integration facility following the Ascent Unit encapsulation.<br />
j) Prior to installation of the LV+Ascent Unit on the pad and following the delivery of the STE to the bunker, the SC<br />
manufacturer shall verify continuity between command RF link and STE and issue to Khrunichev the Certificate of Launch<br />
Pad Readiness to accommodate the LV and Ascent Unit. The spectrum analyzer to be provided by the SC manufacturer shall<br />
be adapted to 220 V 50 Hz.<br />
k) After installation of LV+Ascent Unit and prior to the roll-up of service tower, the SC manufacturer, in conjunction with<br />
Khrunichev, shall checkup the RF link between the Ascent Unit and STE. Such checkup shall be performed 20 minutes after<br />
the mating of the LV aft section. Loral to confirm functionality of the RF link within 45 minutes.<br />
l) At L-6 months, the SC manufacturer shall provide to Khrunichev two connectors for installation by Khrunichev on the<br />
existing RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructions on cable dressing<br />
and cable performances. (If required)<br />
Page C3-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.2-2a: LVIJ1 Umbilical Pin Assignments<br />
Connector<br />
Pin<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
7<br />
8<br />
9<br />
10<br />
11<br />
12<br />
13<br />
14<br />
15<br />
16<br />
17<br />
18<br />
19<br />
20<br />
21<br />
22<br />
23<br />
24<br />
25<br />
26<br />
27<br />
28<br />
29<br />
30<br />
31<br />
32<br />
33<br />
34<br />
35<br />
36<br />
37<br />
38<br />
39<br />
40<br />
41<br />
42<br />
43<br />
44<br />
45<br />
Function Max.<br />
Voltage<br />
(V)*<br />
Max.<br />
Current<br />
(A)<br />
Max.<br />
Voltage at<br />
Liftoff (V)<br />
Max.<br />
Current at<br />
Liftoff (A)<br />
Line<br />
Resistance<br />
(ohms) (4)<br />
Time of<br />
Usage<br />
Page C3-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Connector<br />
Pin<br />
Notes:<br />
46<br />
47<br />
48<br />
49<br />
50<br />
51<br />
52<br />
53<br />
54<br />
55<br />
56<br />
57<br />
58<br />
59<br />
60<br />
61<br />
Function Max.<br />
Voltage<br />
(V)*<br />
Max.<br />
Current<br />
(A)<br />
Max.<br />
Voltage at<br />
Liftoff (V)<br />
Max.<br />
Current at<br />
Liftoff (A)<br />
Line<br />
Resistance<br />
(ohms) (4)<br />
Time of<br />
Usage<br />
a) For the bus power lines, the maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.<br />
*Maximum voltage measured at the M&C interface at the vault<br />
b) SC CONTRACTOR equipment provides protection against exceeding 100 V at spacecraft umbilical interface. It also provides<br />
continuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power through<br />
the umbilicals will be automatically shutoff within 0.2 sec if max current is exceeded by 50%.<br />
c) Separation jumpers are configured as follows:<br />
1) Pins TBD and TBD jumpered on LV side<br />
2) Pins TBD and TBD jumpered on LV side<br />
3) Pins TBD and TBD jumpered on SC side<br />
d) Indicated resistance values are from LVIP1/P2 IFD connection to the KhSC/SC CONTRACTOR EGSE interface in the<br />
vault room (or on the Mobile Service Tower for designated circuits).<br />
e) Some circuits (e.g. battery charging circuits) may be terminated at the LV Mobile Service Tower if necessary<br />
Page C3-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.2.2b: LVIJ2 Umbilical Pin Assignments<br />
Connector<br />
Pin<br />
1<br />
2<br />
3<br />
4<br />
5<br />
6<br />
7<br />
8<br />
9<br />
10<br />
11<br />
12<br />
13<br />
14<br />
15<br />
16<br />
17<br />
18<br />
19<br />
20<br />
21<br />
22<br />
23<br />
24<br />
25<br />
26<br />
27<br />
28<br />
29<br />
30<br />
31<br />
32<br />
33<br />
34<br />
35<br />
36<br />
37<br />
38<br />
39<br />
40<br />
41<br />
42<br />
43<br />
44<br />
45<br />
Function Max.<br />
Voltage<br />
(V)*<br />
Max.<br />
Current<br />
(A)<br />
Max.<br />
Voltage at<br />
Liftoff (V)<br />
Max.<br />
Current at<br />
Liftoff (A)<br />
Line<br />
Resistance<br />
(ohms) (4)<br />
Time of<br />
Usage<br />
Page C3-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Connector<br />
Pin<br />
Notes:<br />
46<br />
47<br />
48<br />
49<br />
50<br />
51<br />
52<br />
53<br />
54<br />
55<br />
56<br />
57<br />
58<br />
59<br />
60<br />
61<br />
Function Max.<br />
Voltage<br />
(V)*<br />
Max.<br />
Current<br />
(A)<br />
Max.<br />
Voltage at<br />
Liftoff (V)<br />
Max.<br />
Current at<br />
Liftoff (A)<br />
Line<br />
Resistance<br />
(ohms) (4)<br />
Time of<br />
Usage<br />
a) For the bus power lines, the maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.<br />
* Maximum voltage measured at the M&C interface in the vault<br />
b) SC Contractor equipment provides protection against exceeding 100 V at spacecraft umbilical interface. It also provides<br />
continuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power through<br />
the umbilicals will be automatically shutoff within 0.2 sec if max current is exceeded by 50%.<br />
c) Separation jumpers are configured as follows:<br />
1) Pins TBD and TBD jumpered on LV side<br />
2) Pins TBD and TBD jumpered on LV side<br />
3) Pins TBD and TBD jumpered on SC side<br />
d) Indicated resistance values are from LVIP1/P2 IFD connection to the KhSC/SC CONTRACTOR EGSE interface in the<br />
vault room (or on the Mobile Service Tower for designated circuits).<br />
e) Some circuits (e.g. battery charging circuits) may be terminated at the LV Mobile Service Tower if necessary.<br />
Page C3-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.5-1: EGSE Description<br />
Connectors<br />
Equipment Power Source Power Required Heat Output Equipment Plug side Facility side<br />
BLDG. 92A-50<br />
Hall 101<br />
Hall 103<br />
Hall 103A<br />
Control Room<br />
Office Areas<br />
Page C3-22
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.5.1-1: SC Contractor Electrical Ground Support Equipment (Continued)<br />
Connectors<br />
Equipment Power Source Power Required Heat Output Equipment Plug side Facility side<br />
AREA 31<br />
Room 100A<br />
Room 100B<br />
Room 119<br />
Room 120<br />
Room 121<br />
Offices<br />
Bldg 44 Hall 1<br />
Page C3-23
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.5.1-1: SC Contractor Electrical Ground Support Equipment (Continued)<br />
Connectors<br />
Equipment Power Source Power Required Heat Output Equipment Plug side Facility side<br />
Bldg 44 Rm 58<br />
BLDG. 92-1<br />
LAUNCH<br />
COMPLEX 81<br />
Bunker (Rm. 250)<br />
Bunker (Rm 251)<br />
Bunker (Rm 244)<br />
Vault (Rm. 64 or 76))<br />
Page C3-24
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Table C3.5-2: Fluids/Gases Requirements<br />
Name Conditions Supplied By Location of Use<br />
Compressed Air for Breathing<br />
(SCAPE)<br />
See Launch Campaign Guide for max available KhSC/ILS Fueling Hall<br />
distilled water See Launch Campaign Guide for max available. KhSC/ILS Processing Hall<br />
Fueling Hall<br />
demineralized water See Launch Campaign Guide for max available KhSC/ILS Decontamination<br />
Area<br />
Nitrogen GOST-92-93-74,<br />
technical grade 1<br />
Nitrogen GOST-92-93-74,<br />
technical grade 1<br />
See Launch Campaign Guide for max available. KhSC/ILS Decontamination<br />
Area<br />
See Launch Campaign Guide for max available KhSC/ILS Fueling Hall<br />
Liquid Nitrogen (LN2) TBD liters KhSC/ILS Fueling Hall<br />
He (Ghe) per spec Mil-p-27407<br />
Type 1, Grade A<br />
TBD K-bottles high pres (400 bar) TBD Kbottles<br />
low pres (135 bar)<br />
MMH TBD Cylinders - Tot weight ea TBD kg max,<br />
TBD kg max prop weight<br />
SC contractor Fueling Hall<br />
SC contractor Fueling Hall<br />
Nitrogen TBD SC contractor Fueling Hall<br />
Nitrogen Tetroxide TBD Cylinders - Tot weight ea TBD kg max,<br />
TBD kg max prop weight<br />
SC contractor Fueling Hall<br />
Shop Air See Launch Campaign Guide for max available KhSC/ILS Fueling Hall<br />
Ethyl Alcohol See Launch Campaign Guide for max available. KhSC/ILS Fueling Hall<br />
Service Water See Launch Campaign Guide for max available KhSC/ILS Fueling Hall<br />
Grade “Extra” or Highest Grade<br />
GOST 18300-87 Ethyl Alcohol<br />
See Launch Campaign Guide for max available. KhSC/ILS Fueling Hall<br />
Freon TBD SC contractor<br />
Argon TBD SC contractor<br />
IPA TBD SC contractor<br />
Page C3-25
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D1. 1194AX-500 Adapter System<br />
D1.1 INTRODUCTION<br />
The 1194AX-500 adapter system is comprised of the 1194 AX clamp band system, separation springs, a payload<br />
adapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interface<br />
characteristics, structural capability, usable volume, and accelerometer installation.<br />
D1.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY<br />
The 1194AX-500 adapter is a one piece conic frustum of monocoque construction with a height of 500 mm and<br />
fabricated from aluminum alloy. Mechanical drawings for the 1194AX-500 adapter are presented in Section D1.7 of<br />
this Appendix.<br />
D1.2.1 Interface Ring Characteristics<br />
The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a load<br />
path between the spacecraft and adapter during ground operations and flight. The outboard features of the combined<br />
cross section are designed to interface with the separation system clamp band. The cross section and material property<br />
characteristics for the spacecraft and adapter interface ring are presented in Table D1.2.1-1 and Figure<br />
D1.2.1-1. The dimensions of the spacecraft interface are presented in Section D1.7 of this Appendix.<br />
Table D1.2.1-1: Spacecraft and Adapter Interface Ring Characteristics<br />
Ring Characteristics Spacecraft Ring Adapter Ring<br />
Height Of Effective Cross Section (L) 25 mm 25 mm<br />
Cross Section Area (A) 460 mm 2 ± 15% 558 mm 2<br />
Moment Of Inertia (Ixx) 52000 mm 4 ± 15% 72000 mm 4<br />
Moment Of Inertia (Iyy) 13400 mm 4 ± 15% 30000 mm 4<br />
Young Modulus (E) 69 x 10 3 MPa 70 x 10 3 MPa<br />
Page D1-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.2.1-1: Spacecraft and Adapter Interface Ring Cross Section<br />
Spacecraft Ring<br />
y<br />
y<br />
Adapter Ring<br />
x<br />
x<br />
x<br />
x<br />
y<br />
y<br />
L<br />
Separation Plane<br />
Separation Plane<br />
Page D1-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Scribe marks on the spacecraft interface ring verify proper alignment of the spacecraft relative to the <strong>launch</strong> <strong>vehicle</strong>.<br />
The attributes and location of the scribe mark for the spacecraft and adapter ring are presented in Section D1.7 of this<br />
Appendix.<br />
D1.2.2 Structural Capability<br />
The structural capability of each adapter system is based on the allowable line load obtained from testing. The<br />
relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1194AX-500 adapter system is<br />
presented in Figure D1.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffness<br />
characteristics presented in Table D1.2-1, the geometry presented in Section D1.7 of this Appendix and the quasistatic<br />
design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculated<br />
by classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reduce<br />
the allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as point<br />
loads through the spacecraft structure close to the interface.<br />
The structural capability presented should only be used as a <strong>guide</strong>line for assessment of interface structural<br />
compatibility. Coupled loads analysis performed early in the <strong>mission</strong> integration will verify margins for structural<br />
loading of the adapter and separation system.<br />
Page D1-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC)<br />
Longitudinal Offset of SC C.G. from<br />
Separation Plane<br />
3.50<br />
3.00<br />
2.50<br />
2.00<br />
1.50<br />
1.00<br />
0.50<br />
0.00<br />
26.6kN Clamp Band Tension<br />
40kN Clamp Band Tension<br />
2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600<br />
SC Mass (kg)<br />
Preliminary<br />
SC Mass<br />
Allowable C.G. Offset (m)<br />
(kg) 40kN Tension 26.6kN Tension<br />
2000 3.07 2.30<br />
2200 2.79 2.09<br />
2400 2.56 1.92<br />
2600 2.37 1.78<br />
2800 2.20 1.65<br />
3000 2.06 1.55<br />
3200 1.93 1.45<br />
3400 1.82 1.37<br />
3600 1.72 1.29<br />
3800 1.63 1.23<br />
4000 1.55 1.17<br />
4200 1.48 1.11<br />
4400 1.42 1.07 Adapter system capability based on the following data:<br />
4600 1.36 1.02 Allowable Limit Line Loads<br />
4800 1.30 0.98 Tension (26.6 kN clamp band tension) Nt = 91 N/mm<br />
5000 1.25 0.94 Tension (40 kN clamp band tension) Nt = 122 N/mm<br />
5200 1.19 0.89 Compression Nc = 156 N/mm<br />
5400 1.14 0.84 Quasi Static Loads<br />
5600 1.08 0.79 Per Figure 3.4.1.2-1 of this Mission Planner's Guide<br />
Page D1-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D1.3 USABLE VOLUME<br />
The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-500 adapter system is<br />
presented in Figures D1.3-1a through D1.3-1e. The spacecraft static envelope (maximum dimensions of unloaded<br />
spacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the<br />
useable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements<br />
due to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DM<br />
may protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements will<br />
not exceed 50 mm. This must be verified by coupled loads analysis.<br />
Page D1-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter<br />
Page D1-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 1 of 4)<br />
Page D1-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 2 of 4)<br />
Page D1-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 3 of 4)<br />
Page D1-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 4 of 4)<br />
Page D1-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D1.4 SEPARATION SYSTEM<br />
The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp band<br />
are preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preload<br />
is 40.0 kN for the1194AX adapter system. The separation system is released when these bolts are severed by bolt<br />
cutters. Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the<br />
retention device moves the clamp band away from the interface ring. The retention device also secures the band to the<br />
adapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and<br />
adapter.<br />
The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planner’s Guide.<br />
To ensure proper separation velocity, a matched set of separation springs are provided. Each separation springs has an<br />
initial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring required<br />
to provide the desired separation velocity and separation rates. Table D1.4-1 defines the maximum and minimum<br />
spring stroke range with the associated spring force characteristics. Spring sets can include any number between two<br />
and twelve.<br />
Table D1.4-1: Separation Spring Characteristics<br />
Stroke<br />
(mm)<br />
Initial Force<br />
(N)<br />
Final Force<br />
Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20<br />
Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20<br />
The separation event is affected by interface hardware that impart force during separation. This hardware consists of<br />
electrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electrical<br />
disconnects and grounding connectors are provided as shown in Section D1.7 of this Appendix. This symmetrical<br />
arrangement is provided to minimize overturning moments at separation. The force profile for each electrical<br />
disconnects is shown in Figure D1.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resists<br />
separation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D1.4-2.<br />
(N)<br />
Page D1-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.4-1: Electrical Disconnect Force Profile for a Single 37 Pin Electrical Connector<br />
Force (N)<br />
200<br />
150<br />
100<br />
50<br />
0<br />
-50<br />
-100<br />
-150<br />
0 1 2 3 4 5 6 7 8 9 10<br />
Force Displacement<br />
(N) (mm)<br />
178 0<br />
156 3.4<br />
-115 3.4<br />
-115 6.7<br />
0 6.7<br />
Note: A positive force assists separation.<br />
Displacement (mm)<br />
Page D1-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D1.4-2: Purge Connector Force Diagram (Typical)<br />
Force (N)<br />
100<br />
80<br />
60<br />
40<br />
20<br />
0<br />
0.00 0.10 0.20 0.30 0.40 0.50 0.60<br />
Axial Travel (mm)<br />
Page D1-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D1.5 ELECTRICAL INTERFACE<br />
In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. These<br />
electrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment and<br />
the <strong>launch</strong> <strong>vehicle</strong>. The standard adapter system includes two diametrically opposed electrical rise off disconnects.<br />
There are two standard configurations for the electrical interface. Details for each of the two standard electrical<br />
interfaces configurations are presented in Section D1.7 of this Appendix. Connectors conforming to MIL-C-81703<br />
have been used for both adapter system configurations. Part numbers for the standard 37 pin connectors are presented<br />
in Table D1.5-1.<br />
Table D1.5-1: Standard Electrical Connectors<br />
Connector ID No. of Pins Deutsch Part No.<br />
Spacecraft Side LVIP1 37 MS3446E37-50P<br />
LVIP2 37 MS3446E37-50P<br />
Launch Vehicle Side LVIJ1 37 MS3464E37-50S<br />
LVIJ2 37 MS3464E37-50S<br />
For each electrical connector, two pins and a loop back on the spacecraft side of the interface are required for <strong>launch</strong><br />
<strong>vehicle</strong> separation indicators. Loop backs on the <strong>launch</strong> <strong>vehicle</strong> side of the interface for indication of separation for the<br />
spacecraft can be provided as required.<br />
Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electrical<br />
connectors and GSE. This information includes available wire types, shielding, voltage requirements, current<br />
requirement, and resistance requirements.<br />
The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuity<br />
across the separation interface. Electrical continuity across the separation interface is provided by two diametrically<br />
opposed electrical grounding connectors as presented in Section D1.7 of this Appendix.<br />
D1.6 INSTRUMENTATION<br />
Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. The<br />
standard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented to<br />
monitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D1.7 of<br />
this Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels for<br />
the Proton <strong>mission</strong> are presented in Section 4.2.1.7 of the Mission Planner’s Guide.<br />
Page D1-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS<br />
Page D1-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 2 of 9)<br />
TBS<br />
Page D1-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 3 of 9)<br />
TBS<br />
Page D1-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 4 of 9)<br />
TBS<br />
Page D1-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 5 of 9)<br />
TBS<br />
Page D1-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 6 of 9)<br />
TBS<br />
Page D1-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 7 of 9)<br />
TBS<br />
Page D1-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 8 of 9)<br />
TBS<br />
Page D1-22
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-500 Adapter Mechanical Drawings (Continued 9 of 9)<br />
TBS<br />
Page D1-23
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D2. 1194AX-625 ADAPTER SYSTEM<br />
D2.1 INTRODUCTION<br />
The 1194AX-625 adapter system is comprised of the 1194 AX clamp band system, separation springs, a payload<br />
adapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interface<br />
characteristics, structural capability, usable volume, and accelerometer installation.<br />
D2.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY<br />
The 1194AX-625 adapter is a one piece conic frustum of monocoque construction 625 mm in height and fabricated<br />
from aluminum alloy. Mechanical drawings for the 1194AX-625 adapter are presented in Section D2.7 of this<br />
Appendix.<br />
D2.2.1 Interface Ring Characteristics<br />
The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a load<br />
path between the spacecraft and adapter during ground operations and flight. The outboard features of the combined<br />
cross section are designed to interface with the separation system clamp band. The cross section and material property<br />
characteristics for the spacecraft and adapter interface ring are presented in Table D2.2.1-1 and Figure<br />
D2.2.1-1. The dimensions of the spacecraft interface are presented in Section D2.7 of this Appendix.<br />
Table D2.2.1-1: Spacecraft and Adapter Interface Ring Characteristics<br />
Ring Characteristics Spacecraft Ring Adapter Ring<br />
Height Of Effective Cross Section(L) 25 mm 25 mm<br />
Cross Section Area (A) 481 mm 2 ± 15% 558 mm 2<br />
Moment Of Inertia (Ixx) 56900 mm 4 ± 15% 72000 mm 4<br />
Moment Of Inertia (Iyy) 13400 mm 4 ± 15% 30000 mm 4<br />
Young Modulus (E) 69 x 10 3 MPa 70 x 10 3 MPa<br />
Page D2-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.2.1-1: Spacecraft and Adapter Interface Ring Cross Section<br />
Spacecraft Ring<br />
y<br />
y<br />
Adapter Ring<br />
x<br />
x<br />
x<br />
x<br />
y<br />
y<br />
L<br />
Separation Plane<br />
Separation Plane<br />
Page D2-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Scribe marks on the spacecraft interface ring verify proper alignment of the spacecraft relative to the <strong>launch</strong> <strong>vehicle</strong>.<br />
The attributes and location of the scribe mark for the spacecraft and adapter ring are presented in Section D2.7 of this<br />
Appendix.<br />
D2.2.2 Structural Capability<br />
The structural capability of each adapter system is based on the allowable line load obtained from testing. The<br />
relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the 1194AX-625 adapter system is<br />
presented in Figure D2.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffness<br />
characteristics presented in Table D2.2-1, the geometry presented in Section D2.7 of this Appendix and the quasistatic<br />
design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculated<br />
by classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reduce<br />
the allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as point<br />
loads through the spacecraft structure close to the interface.<br />
The structural capability presented should only be used as a <strong>guide</strong>line for assessment of interface structural<br />
compatibility. Coupled loads analysis performed early in the <strong>mission</strong> integration will verify margins for structural<br />
loading of the adapter and separation system.<br />
Page D2-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC)<br />
Longitudinal Offset of SC C.G. from<br />
Separation Plane<br />
3.50<br />
3.00<br />
2.50<br />
2.00<br />
1.50<br />
1.00<br />
0.50<br />
0.00<br />
26.6kN Clamp Band Tension<br />
40kN Clamp Band Tension<br />
2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600<br />
SC Mass (kg)<br />
Preliminary<br />
SC Mass<br />
Allowable C.G. Offset (m)<br />
(kg) 40kN Tension 26.6kN Tension<br />
2000 3.07 2.30<br />
2200 2.79 2.09<br />
2400 2.56 1.92<br />
2600 2.37 1.78<br />
2800 2.20 1.65<br />
3000 2.06 1.55<br />
3200 1.93 1.45<br />
3400 1.82 1.37<br />
3600 1.72 1.29<br />
3800 1.63 1.23<br />
4000 1.55 1.17<br />
4200 1.48 1.11<br />
4400 1.42 1.07 Adapter system capability based on the following data:<br />
4600 1.36 1.02 Allowable Limit Line Loads<br />
4800 1.30 0.98 Tension (26.6 kN clamp band tension) Nt = 91 N/mm<br />
5000 1.25 0.94 Tension (40 kN clamp band tension) Nt = 122 N/mm<br />
5200 1.19 0.89 Compression Nc = 156 N/mm<br />
5400 1.14 0.84 Quasi Static Loads<br />
5600 1.08 0.79 Per Figure 3.4.1.2-1 of this Mission Planner's Guide<br />
Page D2-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D2.3 USABLE VOLUME<br />
The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-625 adapter system is<br />
presented in Figures D2.3-1a through D2.3-3e for the Block DM and Breeze M versions of the LV. The spacecraft<br />
static envelope (maximum dimensions of unloaded spacecraft, including manufacturing tolerances and expansion of<br />
thermal blankets) must not protrude beyond the useable volume, except where it is mutually agreed upon by ILS and<br />
Khrunichev. Spacecraft dynamic displacements due to ground or flight loads and deviations caused by an imperfect<br />
installation of the spacecraft on the Fourth Stage may protrude beyond the boundaries of this useable volume. It is<br />
assumed that spacecraft dynamic displacements will not exceed 50 mm. This must be verified by coupled loads<br />
analysis.<br />
Page D2-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter<br />
Page D2-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4)<br />
Page D2-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4)<br />
Page D2-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4)<br />
Page D2-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4)<br />
Page D2-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-2a: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter<br />
Page D2-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-2b: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC)<br />
Page D2-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-2c: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC)<br />
Page D2-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-2d: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC)<br />
Page D2-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-2e: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC)<br />
Page D2-15
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-3a: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter<br />
Page D2-16
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-3b: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC)<br />
Page D2-17
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-3c: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC)<br />
Page D2-18
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-3d: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC)<br />
Page D2-19
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.3-3e: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC)<br />
Page D2-20
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D2.4 SEPARATION SYSTEM<br />
The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp band<br />
are preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preload<br />
is 26.6 or 40.0 kN for the 1194AX adapter system. The separation system is released when these bolts are severed by<br />
bolt cutters. Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the<br />
retention device moves the clamp band away from the interface ring. The retention device also secures the band to the<br />
adapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and<br />
adapter.<br />
The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planners Guide.<br />
To ensure proper separation velocity, a matched set of separation springs are provided. Each separation springs has an<br />
initial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring required<br />
to provide the desired separation velocity and separation rates. Table D2.4-1 defines the maximum and minimum<br />
spring stroke range with the associated spring force characteristics. Spring sets can include any number between two<br />
and twelve.<br />
Table D2.4-1: Separation Spring Characteristics<br />
Stroke<br />
(mm)<br />
Initial Force<br />
(N)<br />
Final Force<br />
Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20<br />
Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20<br />
The separation event is affected by interface hardware that impart force during separation. This hardware consists of<br />
electrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electrical<br />
disconnects and grounding connectors are provided as shown in Section D2.7 of this Appendix. This symmetrical<br />
arrangement is provided to minimize overturning moments at separation. The force profile for each electrical<br />
disconnects is shown in Figure D2.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resists<br />
separation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D2.4-2.<br />
(N)<br />
Page D2-21
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.4-1: Electrical Disconnect Force Profile for a Single 37 Pin Electrical Connector<br />
Force (N)<br />
200<br />
150<br />
100<br />
50<br />
0<br />
-50<br />
-100<br />
-150<br />
0 1 2 3 4 5 6 7 8 9 10<br />
Force Displacement<br />
(N) (mm)<br />
178 0<br />
156 3.4<br />
-115 3.4<br />
-115 6.7<br />
0 6.7<br />
Note: A positive force assists separation.<br />
Displacement (mm)<br />
Page D2-22
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D2.4-2: Purge Connector Force Diagram (Typical)<br />
Force (N)<br />
100<br />
80<br />
60<br />
40<br />
20<br />
0<br />
0.00 0.10 0.20 0.30 0.40 0.50 0.60<br />
Axial Travel (mm)<br />
Page D2-23
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D2.5 ELECTRICAL INTERFACE<br />
In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. These<br />
electrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment and<br />
the <strong>launch</strong> <strong>vehicle</strong>. The standard adapter system includes two diametrically opposed electrical rise off disconnects.<br />
There are two standard configurations for the electrical interface. Details for each of the two standard electrical<br />
interfaces configurations are presented in Section D2.7 of this Appendix. Connectors conforming to MIL-C-81703<br />
have been used for both adapter system configurations. Part numbers for the standard 37 pin connectors are presented<br />
in Table D2.5-1.<br />
Table D2.5-1: Standard Electrical Connectors<br />
Connector ID No. of Pins Deutsch Part No.<br />
Spacecraft Side LVIP1 37 MS3446E37-50P<br />
LVIP2 37 MS3446E37-50P<br />
Launch Vehicle Side LVIJ1 37 MS3464E37-50S<br />
LVIJ2 37 MS3464E37-50S<br />
For each electrical connector, two pins and a loop back on the spacecraft side of the interface are required for <strong>launch</strong><br />
<strong>vehicle</strong> separation indicators. Loop backs on the <strong>launch</strong> <strong>vehicle</strong> side of the interface for indication of separation for the<br />
spacecraft can be provided as required.<br />
Refer to Section 4.2.2 of the Mission Planners Guide for information on the electrical wiring between the electrical<br />
connectors and GSE. This information includes available wire types, shielding, voltage requirements, current<br />
requirement, and resistance requirements.<br />
The requirement for maximum resistance across the separation interface is of 10 milliohms to ensure electrical<br />
continuity across the separation interface. Electrical continuity across the separation interface is provided by two<br />
diametrically opposed electrical grounding connector as presented in Section D2.7 of this Appendix.<br />
D2.6 INSTRUMENTATION<br />
Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. The<br />
standard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented to<br />
monitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D2.7 of<br />
this Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels for<br />
the Proton <strong>mission</strong> are presented in Section 4.2.1.7 of the Mission Planners Guide.<br />
Page D2-24
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS<br />
Page D2-25
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 2 of 9)<br />
TBS<br />
Page D2-26
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 3 of 9)<br />
TBS<br />
Page D2-27
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 4 of 9)<br />
TBS<br />
Page D2-28
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 5 of 9)<br />
TBS<br />
Page D2-29
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 6 of 9)<br />
TBS<br />
Page D2-30
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 7 of 9)<br />
TBS<br />
Page D2-31
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 8 of 9)<br />
TBS<br />
Page D2-32
Proton Mission Planners Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1194AX-625 Adapter Mechanical Drawings(Sheet 9 of 9)<br />
TBS<br />
Page D2-33
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D3. 1666V-1000 ADAPTER SYSTEM<br />
D3.1 INTRODUCTION<br />
The 1666V-1000 adapter system is comprised of the 1666 V clamp band system, separation springs, a payload adapter<br />
and electrical rise off disconnects. This appendix defines the mechanical and electrical interface characteristics,<br />
structural capability, usable volume, and accelerometer installation.<br />
D3.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY<br />
The 1666V-1000 adapter is a two piece conic frustum of monocoque construction 1000 mm in height and fabricated<br />
from aluminum alloy. Mechanical drawings for the 1666V-1000 adapter are presented in Section D3.7 of this<br />
Appendix.<br />
D3.2.1 Interface Ring Characteristics<br />
The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a load<br />
path between the spacecraft and adapter during ground operations and flight. The outboard features of the combined<br />
cross section are designed to interface with the separation system clamp band. The cross section and material property<br />
characteristics for the spacecraft and adapter interface ring are presented in Table D3.2.1-1 and Figure<br />
D3.2.1-1. The dimensions of the spacecraft interface are presented in Section D3.7 of this Appendix.<br />
Table D3.2.1-1: Spacecraft and Adapter Interface Ring Characteristics<br />
Ring Characteristics Spacecraft Ring Adapter Ring<br />
Height Of Effective Cross Section (L) 25 mm 25 mm<br />
Cross Section Area (A) 460 mm 2 ± 15% 392 mm 2<br />
Moment Of Inertia (Ixx) 52000 mm 4 ± 15% 46700 mm 4<br />
Moment Of Inertia (Iyy) 13400 mm 4 ± 15% 19100 mm 4<br />
Young Modulus (E) 69 x 10 3 MPa 70 x 10 3 MPa<br />
Page D3-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.2.1-1: Spacecraft and Adapter Interface Ring Cross Section<br />
Spacecraft Ring<br />
y<br />
y<br />
Adapter Ring<br />
x<br />
x<br />
x<br />
x<br />
y<br />
y<br />
L<br />
Separation Plane<br />
Separation Plane<br />
Page D3-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Scribe marks on the spacecraft interface ring verify proper alignment of the spacecraft relative to the <strong>launch</strong> <strong>vehicle</strong>.<br />
The attributes and location of the scribe mark for the spacecraft and adapter ring are presented in Section D3.7 of this<br />
Appendix.<br />
D3.2.2 Structural Capability<br />
The structural capability of each adapter system is based on the allowable line load obtained from testing. The<br />
relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666V-1000 adapter system is<br />
presented in Figure D3.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffness<br />
characteristics presented in Table D3.2-1, the geometry presented in Section D4.7 of this Appendix and the quasistatic<br />
design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculated<br />
by classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reduce<br />
the allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as point<br />
loads through the spacecraft structure close to the interface.<br />
The structural capability presented should only be used as a <strong>guide</strong>line for assessment of interface structural<br />
compatibility. Coupled loads analysis performed early in the <strong>mission</strong> integration will verify margins for structural<br />
loading of the adapter and separation system.<br />
Page D3-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.2.2-1: Capability of 1666V Adapter System - SC Mass vs Longitudinal C.G. (TBC)<br />
Longitudinal Offset of SC C.G. from<br />
Separation Plane<br />
3.50<br />
3.00<br />
2.50<br />
2.00<br />
1.50<br />
1.00<br />
0.50<br />
0.00<br />
2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600<br />
SC Mass (kg)<br />
Preliminary<br />
Allowable C.G.<br />
SC Mass Offset (m)<br />
(kg) 30.0kN Tension<br />
2000 3.39<br />
2200 3.09<br />
2400 2.83<br />
2600 2.62<br />
2800 2.44<br />
3000 2.28<br />
3200 2.14<br />
3400 2.02<br />
3600 1.89<br />
3800 1.77<br />
4000 1.66 Note: A positive force assists separation.<br />
4200 1.56<br />
4400 1.47<br />
4600 1.39 Adapter system capability based on the following data:<br />
4800 1.32 Allowable Limit Line Loads<br />
5000 1.25 Tension Nt = 69 N/mm<br />
5200 1.18 Compression N c = 86 N/mm<br />
5400 1.12 Quasi Static Loads<br />
5600 1.07 Per Figure 3.4.1.2-1 of this Mission Planner's Guide<br />
Page D3-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D3.3 USABLE VOLUME<br />
The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666V-1000 adapter system is<br />
presented in Figures D3.3-1a through D3.3-1e. The spacecraft static envelope (maximum dimensions of unloaded<br />
spacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the<br />
useable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements<br />
due to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DM<br />
may protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements will<br />
not exceed 50 mm. This must be verified by coupled loads analysis.<br />
Page D3-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter<br />
Page D3-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 1 of 4)<br />
Page D3-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 2 of 4)<br />
Page D3-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 3 of 4)<br />
Page D3-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 4 of 4)<br />
Page D3-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D3.4 SEPARATION SYSTEM<br />
The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp band<br />
are preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preload<br />
is 30.0 kN for the1666 V adapter system. The separation system is released when these bolts are severed by bolt cutters.<br />
Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the retention<br />
device moves the clamp band away from the interface ring. The retention device also secures the band to the adapter<br />
and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter.<br />
The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planner’s Guide.<br />
To ensure proper separation velocity, a matched set of separation springs are provided. Each separation springs has an<br />
initial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring required<br />
to provide the desired separation velocity and separation rates. Table D3.4-1 defines the maximum and minimum<br />
spring stroke range with the associated spring force characteristics. Spring sets can include any number between two<br />
and twelve.<br />
Table D3.4-1: Separation Spring Characteristics<br />
Stroke<br />
(mm)<br />
Initial Force<br />
(N)<br />
Final Force<br />
Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20<br />
Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20<br />
The separation event is affected by interface hardware that impart force during separation. This hardware consists of<br />
electrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects and grounding<br />
connectors are provided pairs as shown in Section D3.7 of this Appendix. This symmetrical arrangement is provided to<br />
minimize overturning moments at separation. The force profile for each electrical disconnects is shown in Figure<br />
D3.4-1.<br />
(N)<br />
Page D3-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D3.4-1: Electrical Disconnect Force Profile for a Single 61 Pin Electrical Connector<br />
Force (N)<br />
400<br />
300<br />
200<br />
100<br />
0<br />
-100<br />
-200<br />
0 1 2 3 4 5 6 7 8 9 10<br />
Force Displacement<br />
(N) (mm)<br />
300 0<br />
281 3.4<br />
-156 3.4<br />
-156 6.7<br />
0 6.7<br />
Note: A positive force assists separation.<br />
Displacement (mm)<br />
Page D3-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D3.5 ELECTRICAL INTERFACE<br />
In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. These<br />
electrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment and<br />
the <strong>launch</strong> <strong>vehicle</strong>. The standard adapter system includes two diametrically opposed electrical rise off disconnects.<br />
There are two standard configurations for the electrical interface. Details for each of the two standard electrical<br />
interfaces configurations are presented in Section D3.7 of this Appendix. Connectors conforming to MIL-C-81703 are<br />
used for both adapter system configurations. Part numbers for the standard 61 pin connectors are presented in Table<br />
D3.5-1.<br />
Table D3.5-1: Standard Electrical Connectors<br />
Connector ID No. of Pins Deutsch Part No.<br />
Spacecraft Side LVIJ1 61 MS3424E61-50S<br />
LVIJ2 61 MS3424E61-50S<br />
Launch Vehicle Side LVIP1 61 MS3446E61-50P<br />
LVIP2 61 MS3446E61-50P<br />
For each electrical connector, two pins and a loop back on the spacecraft side of the interface are required for <strong>launch</strong><br />
<strong>vehicle</strong> separation indicators. Loop backs on the <strong>launch</strong> <strong>vehicle</strong> side of the interface for indication of separation for the<br />
spacecraft can be provided as required.<br />
Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electrical<br />
connectors and GSE. This information includes available wire types, shielding, voltage requirements, current<br />
requirement and resistance requirements.<br />
The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuity<br />
across the separation interface. Electrical continuity across the separation interface is provided by conductive coatings<br />
on both the spacecraft and adapter interface flanges.<br />
D3.6 INSTRUMENTATION<br />
Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. The<br />
standard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented to<br />
monitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D3.7 of<br />
this Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels for<br />
the Proton <strong>mission</strong> are presented in Section 4.2.1.7 of the Mission Planner’s Guide.<br />
Page D3-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS<br />
Page D3-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 2 of 9)<br />
TBS<br />
Page D3-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 3 of 9)<br />
TBS<br />
Page D3-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 4 of 9)<br />
TBS<br />
Page D3-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 5 of 9)<br />
TBS<br />
Page D3-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 6 of 9)<br />
TBS<br />
Page D3-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 7 of 9)<br />
TBS<br />
Page D3-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 8 of 9)<br />
TBS<br />
Page D3-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666V-1000 Adapter Mechanical Drawings(Sheet 9 of 9)<br />
TBS<br />
Page D3-22
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D4. 1666A-1150 ADAPTER SYSTEM<br />
D4.1 INTRODUCTION<br />
The 1666A-1150 adapter system is comprised of the 1666 A clamp band system, separation springs, a payload adapter<br />
and electrical rise off disconnects. This appendix defines the mechanical and electrical interface characteristics,<br />
structural capability, usable volume, and accelerometer installation.<br />
D4.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY<br />
The 1666A-1150adapter is a two piece conic frustum of monocoque construction 1150 mm in height and fabricated<br />
from aluminum alloy. Mechanical drawings for the 1666A-1150adapter are presented in Section D4.7 of this<br />
Appendix.<br />
D4.2.1 Interface Ring Characteristics<br />
The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a load<br />
path between the spacecraft and adapter during ground operations and flight. The outboard features of the combined<br />
cross section are designed to interface with the separation system clamp band. The cross section and material property<br />
characteristics for the spacecraft and adapter interface ring are presented in Table D4.2.1-1 and Figure D4.2.1-1. The<br />
dimensions of the spacecraft interface are presented in Section D4.7 of this Appendix.<br />
Table D4.2.1-1: Spacecraft and Adapter Interface Ring Characteristics<br />
Ring Characteristics Spacecraft Ring Adapter Ring<br />
Height Of Effective Cross Section (L) 25 mm 25 mm<br />
Cross Section Area (A) 460 mm 2 ± 15% 344 mm 2<br />
Moment Of Inertia (Ixx) 52000 mm 4 ± 15% 33800 mm 4<br />
Moment Of Inertia (Iyy) 13400 mm 4 ± 15% 18700 mm 4<br />
Young Modulus (E) 69 x 10 3 MPa 70 x 10 3 MPa<br />
Page D4-1
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.2.1-1: Spacecraft and Adapter Interface Ring Cross Section<br />
Spacecraft Ring<br />
y<br />
y<br />
Adapter Ring<br />
x<br />
x<br />
x<br />
x<br />
y<br />
y<br />
L<br />
Separation Plane<br />
Separation Plane<br />
Page D4-2
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Scribe marks on the spacecraft interface ring verify proper alignment of the spacecraft relative to the <strong>launch</strong> <strong>vehicle</strong>.<br />
The attributes and location of the scribe mark for the spacecraft and adapter ring are presented in Section D4.7 of this<br />
Appendix.<br />
D4.2.2 Structural Capability<br />
The structural capability of each adapter system is based on the allowable line load obtained from testing. The<br />
relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666A-1150 adapter system is<br />
presented in Figure D4.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffness<br />
characteristics presented in Table D4.2-1, the geometry presented in Section D1.7 of this Appendix and the quasistatic<br />
design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculated<br />
by classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reduce<br />
the allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as point<br />
loads through the spacecraft structure close to the interface.<br />
The structural capability presented should only be used as a <strong>guide</strong>line for assessment of interface structural<br />
compatibility. Coupled loads analysis performed early in the <strong>mission</strong> integration will verify margins for structural<br />
loading of the adapter and separation system.<br />
Page D4-3
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.2.2-1: Capability of 1666A Adapter System - SC Mass vs Longitudinal C.G. (TBC)<br />
Longitudinal Offset of SC C.G. from<br />
Separation Plane<br />
3.50<br />
3.00<br />
2.50<br />
2.00<br />
1.50<br />
1.00<br />
0.50<br />
0.00<br />
2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600<br />
SC Mass (kg)<br />
Preliminary<br />
Allowable C.G.<br />
SC Mass Offset (m)<br />
(kg) 30.0kN Tension<br />
2000 3.39<br />
2200 3.09<br />
2400 2.83<br />
2600 2.62<br />
2800 2.44<br />
3000 2.28<br />
3200 2.14<br />
3400 2.02<br />
3600 1.89<br />
3800 1.77<br />
4000 1.66 Note: A positive force assists separation.<br />
4200 1.56<br />
4400 1.47<br />
4600 1.39 Adapter system capability based on the following data:<br />
4800 1.32 Allowable Limit Line Loads<br />
5000 1.25 Tension Nt = 69 N/mm<br />
5200 1.18 Compression N c = 86 N/mm<br />
5400 1.12 Quasi Static Loads<br />
5600 1.07 Per Figure 3.4.1.2-1 of this Mission Planner's Guide<br />
Page D4-4
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D4.3 USABLE VOLUME<br />
The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666A-1150adapter system is<br />
presented in Figures D4.3-1a through D4.3-1e. The spacecraft static envelope (maximum dimensions of unloaded<br />
spacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the<br />
useable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements<br />
due to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DM<br />
may protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements will<br />
not exceed 50 mm. This must be verified by coupled loads analysis.<br />
Page D4-5
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter<br />
Page D4-6
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 1 of 4)<br />
Page D4-7
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 2 of 4)<br />
Page D4-8
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 3 of 4)<br />
Page D4-9
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 4 of 4)<br />
Page D4-10
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D4.4 SEPARATION SYSTEM<br />
The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp band<br />
are preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preload<br />
is 29.5 kN for the1666 A adapter system. The separation system is released when these bolts are severed by bolt cutters.<br />
Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the retention<br />
device moves the clamp band away from the interface ring. The retention device also secures the band to the adapter<br />
and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter.<br />
The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planner’s Guide.<br />
To ensure proper separation velocity, a matched set of separation springs are provided. Each separation springs has an<br />
initial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring required<br />
to provide the desired separation velocity and separation rates. Table D4.4-1 defines the maximum and minimum<br />
spring stroke range with the associated spring force characteristics. Spring sets can include any number between two<br />
and twelve.<br />
Table D4.4-1: Separation Spring Characteristics<br />
Stroke<br />
(mm)<br />
Initial Force<br />
(N)<br />
Final Force<br />
Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20<br />
Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20<br />
The separation event is affected by interface hardware that impart force during separation. This hardware consists of<br />
electrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects connectors are<br />
provided pairs as shown in Section D4.7 of this Appendix. This symmetrical arrangement is provided to minimize<br />
overturning moments at separation. The force profile for each electrical disconnect is shown in Figure D4.4-1.<br />
(N)<br />
Page D4-11
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
Figure D4.4-1: Electrical Disconnect Force Profile for a Single 61 Pin Electrical Connector<br />
Force (N)<br />
400<br />
300<br />
200<br />
100<br />
0<br />
-100<br />
-200<br />
0 1 2 3 4 5 6 7 8 9 10<br />
Force Displacement<br />
(N) (mm)<br />
300 0<br />
281 3.4<br />
-156 3.4<br />
-156 6.7<br />
0 6.7<br />
Note: A positive force assists separation.<br />
Displacement (mm)<br />
Page D4-12
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D4.5 ELECTRICAL INTERFACE<br />
In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. These<br />
electrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment and<br />
the <strong>launch</strong> <strong>vehicle</strong>. The standard adapter system includes two diametrically opposed electrical rise off disconnects.<br />
There are two standard configurations for the electrical interface. Details for each of the two standard electrical<br />
interfaces configurations are presented in Section D4.7 of this Appendix. Connectors conforming to MIL-C-81703<br />
have been used for both adapter system configurations. Part numbers for the standard 61 pin connectors are presented<br />
in Table D4.5-1.<br />
Table D4.5-1: Standard Electrical Connectors<br />
Connector ID No. of Pins Deutsch Part No.<br />
Spacecraft Side LVIJ1 61 MS3424E61-50S<br />
LVIJ2 61 MS3424E61-50S<br />
Launch Vehicle Side LVIP1 61 MS3446E61-50P<br />
LVIP2 61 MS3446E61-50P<br />
For each electrical connector, two pins and a loop back on the spacecraft side of the interface are required for <strong>launch</strong><br />
<strong>vehicle</strong> separation indicators. Loop backs on the <strong>launch</strong> <strong>vehicle</strong> side of the interface for indication of separation for the<br />
spacecraft can be provided as required.<br />
Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electrical<br />
connectors and GSE. This information includes available wire types, shielding, voltage requirements, current<br />
requirement and resistance requirements.<br />
The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuity<br />
across the separation interface. Electrical continuity across the separation interface is provided by conductive coatings<br />
on the spacecraft and adapter interface flanges.<br />
D4.6 INSTRUMENTATION<br />
Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. The<br />
standard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented to<br />
monitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D4.7 of<br />
this Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels for<br />
the Proton <strong>mission</strong> are presented in Section 4.2.1.7 of the Mission Planner’s Guide.<br />
Page D4-13
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS<br />
Page D4-14
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 2 of 9)<br />
TBS<br />
Page D4-15
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 3 of 9)<br />
TBS<br />
Page D4-16
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 4 of 9)<br />
TBS<br />
Page D4-17
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 5 of 9)<br />
TBS<br />
Page D4-18
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 6 of 9)<br />
TBS<br />
Page D4-19
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 7 of 9)<br />
TBS<br />
Page D4-20
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 8 of 9)<br />
TBS<br />
Page D4-21
Proton Mission Planner’s Guide, LKEB-9812-1990<br />
Issue 1, Revision 4, March 1, 1999<br />
1666A-1150 Adapter Mechanical Drawings(Sheet 9 of 9)<br />
TBS<br />
Page D4-22