Proton Rev.4

Revision 4 March 1999

PROTON LAUNCH VEHICLEMISSION PLANNER’S GUIDE

REVISION NOTICEThis document supersedes the Proton Launch Vehicle User’sGuide - Revision 3, Issue 1 dated February 1997

DISCLOSURE OF DATA LEGENDThese technical data are exempt from the licensing requirementof International Traffic in Arms Regulation (ITAR) under ITARPart 125.4(b)(13) and are cleared for public release.LKEB-9812-1990 1999 International Launch Services

International Launch Services101 West Broadway, Suite 2000San Diego, California 92101 USA

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Revision 4 March 1999

PROTON LAUNCH VEHICLEMISSION PLANNER’S GUIDE

Anatoli I. KiselevGeneral DirectorKhrunichev State Research andProduction Space Center

Wilbur C. TraftonPresidentInternational Launch Services

Anatoli K. NedaivodaGeneral DesignerSalyut Design BureauKhrunichev State Research andProduction Space Center

Eric F. LaursenChief EngineerInternational Launch Services

International Launch Services101 West Broadway, Suite 2000San Diego, California 92101 USA

Proton Mission Planner’s Guide, LKEB-9812-1990Issue 1, Revision 4, March 1, 1999

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PREFACE

The Proton Mission Planner’s Guide is intended to provide information to potential Customers and spacecraftsuppliers, concerning spacecraft design criteria, Proton launch capability, available mission analysis and customengineering support, documentation availability and requirements, and program planning. It is intended to serve as anaid to the planning of future missions but should not be construed as a contractual commitment.

The units of measurement referred to in this document are based on the International System of Units (SI), withEnglish units given in parenthesis and all identified dimensions shown should be considered as approximate. In theevent that one or more dimensions are critical to a specific payload intregration or processing operation, the SCCustomer should obtain accurate dimensions from International Launch Services (ILS).

This Guide will be updated or revised periodically. All comments and suggestions for additional information arehereby solicited and will be greatly appreciated.

Those Customers wishing to receive revisions and updates to this manual, or who wish to submit comments orsuggestions, are asked to kindly contact:

International Launch Services101 West Broadway, Suite 2000

San Diego, California 92101 USATelephone: (619) 645-6400Facsimile: (619) 645-6500

http:// www.lmco.com/ILS/


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REVISION HISTORY

Revision Date Revision No. Change Description Approval

15 December 1993 1, Issue 1 Eric LaursenChief Engineer, LKEI

December 1995 2, Issue 1 Eric LaursenChief Engineer, LKEIProton Division, ILS

February 1997 3, Issue 1 Section 1• Updated Integration Schedule• Minor typographical corrections

Section 2• Proton M fairing dimension update• Launch history update and corrections• Failure/Corrective Action update

Section 3• Addition of Proton K/Block DM

performance with use of standard kerosene

Section 4• Ground Ops Instrumentation measurement

capabilities update• Updated Proton LV radiated emmissions• Updated flight instrumentations capabilities• Updated flight loads environments• Updated flight acoustics

Section 6• Updated Proton/BlockDM usable fairing

envelopes with standard adapters

Section 7• Updated Mission Integration schedule• Updated analysis, meetings and

documentation schedules

Eric LaursenChief Engineer, LKEIProton Division, ILS

March 1999 4, Issue 1 • Complete rewrite/update of document toreflect flight measured environments,interfaces and performance

• Addition of Proton M/Breeze M vehicledata

• Discussion of Baikonur payload processingand Launch operations facilities

Eric LaursenProton Chief Engineer,ILS

Rich WatermanManager MissionDevelopment, ILS


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FOREWORD

International Launch Services (ILS), is pleased to offer one of the most capable commercial launch vehicles, and themost comprehensive launch services, available today. The Proton’s services are now available to worldwide Customersat a most competitive price.

ILS is the exclusive marketing agent for commercial sales of the Proton launch vehicle worldwide, and is supported inits operations by full access to the incomparable technological expertise of its parent companies; Lockheed MartinCorporation (LMC), Khrunichev State Research and Production Space Center (KhSC), and Russian Space ComplexEnergia ( Energia). ILS provides customers with a single point of contact for all mission analyses, custom engineering,and launch support tasks involved in using the Proton launch vehicle. Both individually and collectively, the membersof the ILS team are committed to providing the most cost-effective launch services available in the world-from initialprogram planning to successful spacecraft launch.

This document provides performance, environments and interfaces for Proton vehicles that have already been qualifiedto manage payloads up to 4.8 mt in mass. For SC exceeding 4.8 mt, parameters referred to in this document should beconsidered as PRELIMINARY and will be further defined/refined considering the geometry, mass properties andother physical characteristics of a particular spacecraft.


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TABLE OF CONTENTS

PREFACE ................................................................................................................................................... II

REVISION HISTORY.................................................................................................................................III

FOREWORD ...............................................................................................................................................IV

TABLE OF CONTENTS................................................................................................................................ V

LIST OF TABLES........................................................................................................................................XI

LIST OF FIGURES......................................................................................................................................XI

ABBREVIATIONS AND ACRONYMS........................................................................................................ XIX

1. PROTON LAUNCH SERVICES.............................................................................................................. 1-1

1.1 CONSTITUENT ILS COMPANIES........................................................................................................... 1-21.2 ILS CONSTITUENT COMPANY EXPERTISE......................................................................................... 1-31.3 ADVANTAGES OF USING THE PROTON LAUNCH VEHICLE .......................................................... 1-41.4 VEHICLE DESCRIPTION ......................................................................................................................... 1-51.5 GENERAL DESCRIPTION OF THE PROTON FAMILY ....................................................................... 1-71.6 PROTON THREE - STAGE BOOSTER..................................................................................................... 1-8

1.6.1 Proton First Stage ................................................................................................................................. 1-101.6.2 Proton Second Stage.............................................................................................................................. 1-101.6.3 Proton Third Stage................................................................................................................................ 1-101.6.4 Proton Flight Control System ............................................................................................................... 1-11

1.7 BLOCK DM FOURTH STAGE ................................................................................................................ 1-111.8 BREEZE M FOURTH STAGE................................................................................................................. 1-151.9 PAYLOAD FAIRINGS .............................................................................................................................. 1-181.10 BAIKONUR INFRASTRUCTURE AND SERVICES ........................................................................... 1-23

1.10.1 Baikonur Infrastructure ....................................................................................................................... 1-231.10.2 Proton Launch Campaign.................................................................................................................... 1-25

1.11 PLANNED ENHANCEMENTS ............................................................................................................. 1-271.12 PROTON PRODUCTION AND OPERATIONS-SLIDES.................................................................... 1-27

2. VEHICLE PERFORMANCE .................................................................................................................. 2-1

2.1 OVERVIEW .................................................................................................................................................. 2-12.2 PROTON LAUNCH SYSTEM CAPABILITIES......................................................................................... 2-1

2.2.1 The Baikonur Launch Site....................................................................................................................... 2-22.2.2 Launch Availability ................................................................................................................................. 2-22.2.3 Payload Fairings and Adapters ................................................................................................................ 2-32.2.4 Upper Stage Capabilities ......................................................................................................................... 2-3

2.3 PROTON ASCENT PROFILE .................................................................................................................... 2-42.3.1 Proton Booster Ascent ............................................................................................................................. 2-42.3.2 Block DM Trajectory and Sequence ........................................................................................................ 2-8


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2.3.3 Breeze M Trajectory Sequence ...............................................................................................................2-122.3.4 Collision and Contamination Avoidance Maneuver............................................................................... 2-13

2.4 PERFORMANCE GROUNDRULES....................................................................................................... 2-142.4.1 Payload Systems Mass Definition.......................................................................................................... 2-142.4.2 Payload Fairings.................................................................................................................................... 2-152.4.3 Mission Analysis Groundrules ............................................................................................................... 2-152.4.4 Performance Confidence Levels ............................................................................................................ 2-15

2.5 DIRECT INJECTION LEO MISSIONS................................................................................................... 2-152.6 GEOSYNCHRONOUS TRANSFER MISSIONS..................................................................................... 2-16

2.6.1 Launch to Geosynchronous Transfer Orbit ............................................................................................ 2-162.7 ORBIT INJECTION ACCURACY.............................................................................................................2-232.8 SPACECRAFT ORIENTATION AND SEPARATION.............................................................................2-242.9 LAUNCH VEHICLE TELEMETRY DATA ............................................................................................. 2-252.10 MISSION OPTIMIZATION/ PERFORMANCE ENHANCEMENTS..................................................2-29

2.10.1 Nonstandard Mission Designs ..............................................................................................................2-292.10.2 Subsynchronous Transfer ......................................................................................................................2-292.10.3 Super Synchronous Transfer .................................................................................................................2-29

3. SPACECRAFT ENVIRONMENTS .......................................................................................................... 3-1

3.1 THERMAL/HUMIDITY ............................................................................................................................ 3-13.1.1 Ground Thermal Environment................................................................................................................ 3-13.1.2 Ascent ..................................................................................................................................................... 3-63.1.3 Orbit ....................................................................................................................................................... 3-73.1.4 Humidity ................................................................................................................................................ 3-73.1.5 Air Impingement Velocity ....................................................................................................................... 3-7

3.2 CONTAMINATION ENVIRONMENT...................................................................................................... 3-83.2.1 Ground Contamination Control.............................................................................................................. 3-83.2.2 In Flight Contamination Control ............................................................................................................ 3-8

3.3 PRESSURE .................................................................................................................................................. 3-93.3.1 Payload Compartment Venting ............................................................................................................... 3-9

3.4 MECHANICAL LOADS............................................................................................................................ 3-103.4.1 Quasi-Static Loads................................................................................................................................ 3-103.4.2 Sine and Random Vibration Loads ........................................................................................................ 3-143.4.3 Acoustic Loads ...................................................................................................................................... 3-193.4.4 Shock Loads.......................................................................................................................................... 3-213.4.5 Environmental Test Requirements ........................................................................................................ 3-23

3.5 ELECTROMAGNETIC COMPATIBILITY ............................................................................................ 3-243.5.1 EMI Safety Margin (EMISM) ............................................................................................................. 3-243.5.2 Radiated Emissions ............................................................................................................................... 3-243.5.3 RF Transmitter/Receiver Systems EMC............................................................................................... 3-29

4. SPACECRAFT INTERFACES................................................................................................................. 4-1

4.1 MECHANICAL INTERFACES .................................................................................................................. 4-14.1.1 Structural Interfaces ................................................................................................................................ 4-14.1.2 General SC Structural and Load Requirements ....................................................................................... 4-14.1.3 Fairing Interfaces .................................................................................................................................... 4-24.1.4 GN2/Dry Air Purge Option .................................................................................................................. 4-23


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4.2 ELECTRICAL INTERFACES .................................................................................................................. 4-244.2.1 Airborne Interfaces................................................................................................................................ 4-244.2.2 Launch Pad EGSE Interfaces................................................................................................................ 4-404.2.3 Telemetry/Command Links.................................................................................................................. 4-414.2.4 Electrical Grounding............................................................................................................................. 4-474.2.5 Electrical Bonding ................................................................................................................................ 4-474.2.6 SC/LV Lightning Protection ................................................................................................................. 4-474.2.7 Static Discharge .................................................................................................................................... 4-47

4.3 FITCHECK OF MECHANICAL/ELECTRICAL INTERFACES........................................................... 4-47

5. MISSION INTEGRATION AND MANAGEMENT .................................................................................. 5-1

5.1 MANAGEMENT PROVISIONS................................................................................................................. 5-15.1.1 Key Personnel ......................................................................................................................................... 5-15.1.2 Interface Control Document (ICD) ........................................................................................................ 5-15.1.3 Schedule Monitoring............................................................................................................................... 5-15.1.4 Documentation Control and Delivery ..................................................................................................... 5-45.1.5 Meetings and Reviews ............................................................................................................................. 5-45.1.6 DTRA Oversight ..................................................................................................................................... 5-85.1.7 Quarterly Report ..................................................................................................................................... 5-85.1.8 Quality Provisions ................................................................................................................................... 5-85.1.9 Launch License And Permits................................................................................................................... 5-8

5.2 ILS DELIVERABLES.................................................................................................................................. 5-85.2.1 ICD Development................................................................................................................................... 5-95.2.2 Preliminary and Critical Design.............................................................................................................. 5-95.2.3 Spacecraft Testing - Fitcheck/ Shock Test Plan, and Report ................................................................. 5-115.2.4 Safety .................................................................................................................................................... 5-115.2.5 Launch Campaign and Launch.............................................................................................................. 5-115.2.6 Management and Reports...................................................................................................................... 5-11

5.3 CUSTOMER DELIVERABLES................................................................................................................ 5-115.3.1 ICD Development ................................................................................................................................ 5-135.3.2 Preliminary and Critical Design............................................................................................................ 5-135.3.3 Spacecraft Testing ................................................................................................................................. 5-135.3.4 Required Safety Data and Certificates................................................................................................... 5-135.3.5 Launch Campaign and Launch.............................................................................................................. 5-14

5.4 SPECIFIC CUSTOMER RESPONSIBILITIES....................................................................................... 5-145.4.1 Campaign Duration .............................................................................................................................. 5-145.4.2 Spacecraft And Associated Ground Equipment ..................................................................................... 5-145.4.3 Final Spacecraft Data............................................................................................................................ 5-155.4.4 Spacecraft Readiness ............................................................................................................................. 5-155.4.5 Removal of SC Support Equipment ...................................................................................................... 5-155.4.6 Evaluation Of Launch Vehicle And Associated Services ......................................................................... 5-155.4.7 Spacecraft Propellants ........................................................................................................................... 5-155.4.8 Connectors ............................................................................................................................................ 5-15

5.5 ILS SERVICES AND MATERIAL SPECIFICALLY EXCLUDED ........................................................ 5-16

6. SPACECRAFT AND LAUNCH FACILITIES ........................................................................................... 6-1

6.1 FACILITIES OVERVIEW ........................................................................................................................... 6-1


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6.1.1 Yubeleini Airport..................................................................................................................................... 6-16.1.2 Building 92A-50 ..................................................................................................................................... 6-36.1.3 Area 31 Facilities .................................................................................................................................... 6-36.1.4 Building 92-1 and the Proton Launch Zone (Area 81) ............................................................................ 6-46.1.5 Hotels ..................................................................................................................................................... 6-4

6.2 SPACECRAFT PROCESSING FACILITIES............................................................................................. 6-46.2.1 Facility 92A-50....................................................................................................................................... 6-56.2.2 Area 31, Buildings 40/40D General Description ................................................................................... 6-146.2.3 Area 31, Bldg 44 - General Description ................................................................................................ 6-206.2.4 Building 254 (TBD).............................................................................................................................. 6-236.2.5 Area 92, Building 92-1 - General Description....................................................................................... 6-23

6.3 LAUNCH COMPLEX FACILITIES......................................................................................................... 6-246.3.1 Area 81, Launch Pad 23 - General Description ..................................................................................... 6-246.3.2 Facility Layout & Area Designations..................................................................................................... 6-26

7. LAUNCH CAMPAIGN ......................................................................................................................................7-1

7.1 ORGANIZATIONAL RESPONSIBILITIES.............................................................................................. 7-17.1.1 Khrunichev.............................................................................................................................................. 7-17.1.2 Strategic Rocket Forces ........................................................................................................................... 7-17.1.3 Energia (Block DM launches only) ......................................................................................................... 7-17.1.4 ILS.......................................................................................................................................................... 7-27.1.5 Spacecraft Customer ............................................................................................................................... 7-2

7.2 CAMPAIGN ORGANIZATION ................................................................................................................. 7-27.2.1 Contractual and Planning Organization................................................................................................... 7-27.2.2 Organization During Combined Operations ............................................................................................ 7-37.2.3 Planning Meetings................................................................................................................................... 7-4

7.3 COUNTDOWN ORGANIZATION ............................................................................................................ 7-57.4 ABORT CAPABILITY ................................................................................................................................. 7-67.5 LAUNCH OPERATIONS OVERVIEW....................................................................................................... 7-7

7.5.1 Launch Vehicle Processing .................................................................................................................... 7-127.5.2 Spacecraft Preparations Through Arrival ............................................................................................... 7-127.5.3 Area 31 (Buildings 40, 40D, 44) - Spacecraft Testing, Fueling, and Ascent Unit integration ................ 7-127.5.4 Area 92 (Building 92A-50) - Spacecraft Testing, Fueling, and Ascent Unit Integration ........................ 7-14

7.6 LAUNCH VEHICLE INTEGRATION (BUILDING 92-1) THRU LAUNCH PAD OPERATIONS.... 7-157.7 LAUNCH PAD OPERATIONS................................................................................................................. 7-16

8. PROTON LAUNCH SYSTEM ENHANCEMENTS.................................................................................. 8-1

8.1 HIGH VOLUME PAYLOAD FAIRINGS................................................................................................... 8-38.2 TANDEM LAUNCH SYSTEM ................................................................................................................... 8-58.3 AVIONICS SYSTEM MASS UPGRADES.................................................................................................. 8-78.4 NEW CYROGENIC UPPER STAGE ......................................................................................................... 8-78.5 SUMMARY.................................................................................................................................................. 8-7

A. PROTON LAUNCH SYSTEM HISTORY................................................................................................A-1

A.1 BACKGROUND AND HISTORY..............................................................................................................A-1A.2 PROTON FLIGHT HISTORY ...................................................................................................................A-2A.3 DETAILED FLIGHT HISTORY ...............................................................................................................A-4A.4 FAILURES CAUSES AND CORRECTIVE ACTION.............................................................................A-10


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B1. QUALITY SYSTEM ...........................................................................................................................B1-1

C1. GENERAL INFORMATION...............................................................................................................C1-1

C2. INPUT DOCUMENTS....................................................................................................................... C2-2

C3. INTERFACES.................................................................................................................................... C3-3

C3.1 MECHANICAL INTERFACES .............................................................................................................C3-3C3.2 ELECTRICAL INTERFACES................................................................................................................C3-4C3.3 ENVIRONMENTAL INTERFACES......................................................................................................C3-5C3.4 FLIGHT DESIGN ..................................................................................................................................C3-6C3.5 OPERATIONS.........................................................................................................................................C3-7C3.6 OPERATIONS (CONTINUED).............................................................................................................C3-8

D1. 1194AX-500 ADAPTER SYSTEM...................................................................................................... D1-1

D1.1 INTRODUCTION................................................................................................................................. D1-1D1.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY ................................................. D1-1

D1.2.1 Interface Ring Characteristics............................................................................................................ D1-1D1.2.2 Structural Capability ......................................................................................................................... D1-3

D1.3 USABLE VOLUME ............................................................................................................................... D1-5D1.4 SEPARATION SYSTEM ..................................................................................................................... D1-11D1.5 ELECTRICAL INTERFACE .............................................................................................................. D1-14D1.6 INSTRUMENTATION ....................................................................................................................... D1-14D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS..................................................................... D1-15

D2. 1194AX-625 ADAPTER SYSTEM .......................................................................................................D2-1

D2.1 INTRODUCTION ................................................................................................................................. D2-1D2.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY.................................................. D2-1


D2.3 USABLE VOLUME ............................................................................................................................... D2-5D2.4 SEPARATION SYSTEM ..................................................................................................................... D2-21D2.5 ELECTRICAL INTERFACE .............................................................................................................. D2-24D2.6 INSTRUMENTATION ....................................................................................................................... D2-24D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS..................................................................... D2-25

D3. 1666V-1000 ADAPTER SYSTEM .......................................................................................................D3-1



D3.3 USABLE VOLUME ............................................................................................................................... D3-5D3.4 SEPARATION SYSTEM ..................................................................................................................... D3-11D3.5 ELECTRICAL INTERFACE .............................................................................................................. D3-13D3.6 INSTRUMENTATION ....................................................................................................................... D3-13D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS...................................................................... D3-14


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D4. 1666A-1150 ADAPTER SYSTEM ...................................................................................................... D4-1



D4.3 USABLE VOLUME ............................................................................................................................... D4-5D4.4 SEPARATION SYSTEM ..................................................................................................................... D4-11D4.5 ELECTRICAL INTERFACE .............................................................................................................. D4-13D4.6 INSTRUMENTATION ....................................................................................................................... D4-13D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS ..................................................................... D4-14


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LIST OF FIGURES

FIGURE 1-1: ILS LAUNCH SERVICES CHARTER ...................................................................................................... 1-1FIGURE 1.1-1: ILS CORPORATE PARENTAGE.......................................................................................................... 1-2FIGURE 1.1-2: ILS/LKE TASK PARTITIONS............................................................................................................ 1-3FIGURE 1.2-1: LAUNCH EXPERIENCE ..................................................................................................................... 1-4FIGURE 1.3-1: BENEFITS TO THE SPACECRAFT DESIGNER AND OWNER................................................................ 1-5FIGURE 1.4-1: PROTON LAUNCH VEHICLES FLIGHT PROVEN HARDWARE.............................................................. 1-6FIGURE 1.6-1: PROTON K AND M MAJOR HARDWARE ELEMENTS ......................................................................... 1-9FIGURE 1.7-1: PROTON K/BLOCK DM MAJOR HARDWARE ELEMENTS................................................................ 1-12FIGURE 1.7-2: PROTON DM UPPER STAGE WITHIN ITS AERODYNAMIC SHROUDS .............................................. 1-13FIGURE 1.7-3: BLOCK DM AS IT APPEARS IN FLIGHT.......................................................................................... 1-14FIGURE 1.8-1: PROTON M/BREEZE M MAJOR HARDWARE ELEMENTS ................................................................ 1-16FIGURE 1.8-2: BREEZE M (WITH AUXILIARY PROPELLANT TANK) ....................................................................... 1-17FIGURE 1.8-3: THE BREEZE M AND BREEZE M CORE AS THEY APPEAR IN FLIGHT............................................ 1-17FIGURE 1.9-1: PROTON K AND M PAYLOAD FAIRINGS ......................................................................................... 1-19FIGURE 1.9-2: BLOCK DM PAYLOAD FAIRING FOR SINGLE SPACECRAFT ............................................................ 1-20FIGURE 1.9-3: BREEZE-M PAYLOAD FAIRING (STANDARD) FOR SINGLE SPACECRAFT....................................... 1-21FIGURE 1.9-4: BREEZE M PAYLOAD FAIRING (LONG) ..........................................................................................1-22FIGURE 1.10.1-1: OVERALL LAYOUT OF THE BAIKONUR COSMODROME .............................................................. 1-23FIGURE 1.10.1.2-1: BUILDING 92A-50 SPACE VEHICLE PROCESSING FACILITY ................................................... 1-24FIGURE 1.10.2-1: PROTON LAUNCH CAMPAIGN FOR 92A-50 ............................................................................... 1-26FIGURE S-1: TANK COMPONENT FABRICATION AT KHRUNICHEV USES AUTOMATED MACHINING CENTERS........ S-1FIGURE S-2: FIRST STAGE PROPELLANT TANK AUTOMATED WELD-UP................................................................. S-1FIGURE S-3: FIRST, SECOND, AND THIRD STAGE SUB-ASSEMBLIES AWAITING INTEGRATION.............................. S-2FIGURE S-4: PROTON INTERSTAGE COMPONENT FABRICATION ............................................................................. S-2FIGURE S-5: FIRST STAGE BUILD UP ON PROTON FIXTURE................................................................................... S-3FIGURE S-6: INTERSTAGE JOINING SECOND AND THIRD STAGES........................................................................... S-3FIGURE S-7: RD-253 HIGH-PRESSURE ENGINE ON THE FIRST STAGE EXTERNAL FUEL TANK............................S-5FIGURE S-8: PROTON FINAL ASSEMBLY HALL AT KHRUNICHEV ............................................................................S-5FIGURE S-9: ASSEMBLED FIRST STAGE SHOWING HOLD DOWN POINTS AND AFT END SERVICES

CONNECTORS .................................................................................................................................................. S-6FIGURE S-10: END-TO END TEST OF ASSEMBLED PROTON AT KHRUNICHEV FACTORY ....................................... S-6FIGURE S-11: THE BLOCK DM UNDERGOING FINAL ASSEMBLY AND TESTING AT THE ENERGIA PLANT IN

KOROLEV, NEAR MOSCOW............................................................................................................................... S-7FIGURE S-12: THE COMPLETED BLOCK DM STAGE, BEFORE ATTACHMENT OF THE EXTERNAL SHROUD ........... S-7FIGURE S-13: FINISHED BLOCK DM STAGES IN THEIR AERODYNAMIC SHROUDS, AWAITING SHIPMENT TO

BAIKONUR ....................................................................................................................................................... S-8FIGURE S-14: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING EXTERNAL VIEW .......................................... S-8FIGURE S-15: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING INTERNAL VIEW........................................... S-9FIGURE S-16: BREEZE M STAGE STRUCTURAL COMPONENTS IN MANUFACTURE ................................................ S-9FIGURE S-17: BREEZE M CORE STAGE MANUFACTURING .................................................................................. S-10FIGURE S-18: BREEZE M AVIONICS BAY ASSEMBLY ............................................................................................. S-10FIGURE S-19: BREEZE M STAGE FINAL ASSEMBLY .............................................................................................. S-11FIGURE S-20: SPACECRAFT ARRIVAL AT BAIKONUR.............................................................................................. S-11


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FIGURE S-21: SPACECRAFT PROCESSING IN AREA 92 (SVPF).............................................................................. S-12FIGURE S-22: SPACECRAFT PROCESSING AND ENCAPSULATION IN AREA 31........................................................ S-13FIGURE S-23: THE COMPLETE PROTON LAUNCH VEHICLE BEING LIFTED ONTO ITS

TRANSPORTER/ERECTOR .............................................................................................................................. S-13FIGURE S-24: LAUNCH VEHICLE INTERFACES AT A PROTON LAUNCH PAD, BAIKONUR....................................... S-14FIGURE S-25: PROTON ERECTION AT THE PAD ..................................................................................................... S-14FIGURE S-26: MOBILE SERVICE TOWER (MST).................................................................................................... S-15FIGURE S-27: LIFT-OFF OF THE PROTON K/BLOCK DM CARRYING A COMMERCIAL COMMUNICATIONS SATELLITE

...................................................................................................................................................................... S-16FIGURE 2.2.1-1: THE BAIKONUR LAUNCH SITE, SHOWING AVAILABLE DIRECT INJECTION INCLINATIONS ........... 2-2FIGURE 2.3.1-1: TYPICAL PROTON BOOSTER ASCENT ............................................................................................ 2-6FIGURE 2.3.1-2: TYPICAL PROTON BOOSTER ASCENT GROUND TRACK AND VACUUM IMPACT POINTS................. 2-6FIGURE 2.3.1-3: TYPICAL GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO THE

SUPPORT ORBIT............................................................................................................................................... 2-7FIGURE 2.3.1-4: TYPICAL PROTON LOWER ASCENT ALTITUDE, INERTIAL VELOCITY, ACCELERATION, AND

DYNAMIC PRESSURE ....................................................................................................................................... 2-7FIGURE 2.3.2-1: PROTON/BLOCK DM UPPER ASCENT.......................................................................................... 2-9FIGURE 2.3.2-2: BLOCK DM TWO-IMPULSE TRANSFER TO GEOSYNCHRONOUS TRANSFER ORBIT ..................... 2-10FIGURE 2.3.2-3: PROTON K/BLOCK DM UPPER ASCENT GROUND TRACK TO GSO ........................................... 2-11FIGURE 2.3.2-4: GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO A GSO ................................ 2-11FIGURE 2.3.3-1: TYPICAL BREEZE M FLIGHT PROFILE TO GEOSYNCHRONOUS TRANSFER ORBIT ...................... 2-13FIGURE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS

TRANSFER ORBITS ......................................................................................................................................... 2-17FIGURE 2.6.1-2: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS

TRANSFER ORBITS ......................................................................................................................................... 2-19FIGURE 2.8-1: SC SEPARATION ATTITUDE ........................................................................................................... 2-25FIGURE 3.1.1.1-1A: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS

TIMELINE (BLOCK DM) .................................................................................................................................. 3-4FIGURE 3.1.1.1-1B: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS

TIMELINE (BREEZE M) ................................................................................................................................... 3-5FIGURE 3.1.1.2-1: SUPPLEMENTAL FAIRING AIR CONDITIONING SCHEMATIC (REPRESENTATIVE; DETAILED DESIGN

CONDUCTED PER CUSTOMER REQUEST) .......................................................................................................... 3-6FIGURE 3.3.1-1: TYPICAL VENTING PROFILE DURING ASCENT ............................................................................. 3-9FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS .................................................................................. 3-11FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS (CONTINUED)........................................................... 3-12FIGURE 3.4.1.2-2: FLIGHT LIMIT ACCELERATIONS AT THE SC INTERFACE.......................................................... 3-13FIGURE 3.4.2-1: EQUIVALENT SINE LEVELS AT SPACECRAFT INTERFACE - FLIGHT ENVIRONMENT .................... 3-15FIGURE 3.4.2-2: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC IN CONTAINER

AND SC ATTACHED TO ASCENT UNIT .......................................................................................................... 3-16FIGURE 3.4.2-3: TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER, TRANSPORTATION OF SC IN

KHSC CONTAINER, TRANSPORTATION OF ASCENT UNIT .............................................................................. 3-17FIGURE 3.4.2-4: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC AND ASCENT

UNIT ATTACHED TO L/V................................................................................................................................ 3-18FIGURE 3.4.2-4: TRANSPORTATION OF INTEGRATED PROTON LV........................................................................ 3-19FIGURE 3.4.3-1: MAX EXPECTED ACOUSTIC ENVIRONMENT (THIRD OCTAVE).................................................... 3-20


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FIGURE 3.4.4-1: PYROSHOCK SPECTRUM AT ADAPTER/PAYLOAD INTERFACE......................................................3-22FIGURE 3.5.2-1A: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS

(PROTON K/BLOCK DM) (TBC)................................................................................................................... 3-27FIGURE 3.5.2-1B: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS

(PROTON M/BREEZE M) (TBC) ....................................................................................................................3-28FIGURE 3.5.2-2: LAUNCH VEHICLE AND LAUNCH PAD RADIATED SUSCEPTIBILITY LIMITS................................. 3-29FIGURE 4.1.1-1: LV COORDINATE SYSTEM ............................................................................................................ 4-1FIGURE 4.1.3.1-1A: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 3) .................. 4-4FIGURE 4.1.3.1-1B: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 3) .................. 4-5FIGURE 4.1.3.1-1C: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 3).................. 4-6FIGURE 4.1.3.1-2A: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 1 OF 3) ........................................................................................................................... 4-7FIGURE 4.1.3.1-2B: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 2 OF 3) ........................................................................................................................... 4-8FIGURE 4.1.3.1-2C: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 3 OF 3) ........................................................................................................................... 4-9FIGURE 4.1.3.1-3A: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT

(SHEET 1 OF 3)............................................................................................................................................... 4-10FIGURE 4.1.3.1-3B: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT

(SHEET 2 OF 3)............................................................................................................................................... 4-11FIGURE 4.1.3.1-3C: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT

(SHEET 3 OF 3)............................................................................................................................................... 4-12FIGURE 4.1.3.1-4A: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 1 OF 3) ......................................................................................................................... 4-13FIGURE 4.1.3.1-4B: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 2 OF 3) ......................................................................................................................... 4-14FIGURE 4.1.3.1-4C: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME

DIMENSIONS (SHEET 3 OF 3) ......................................................................................................................... 4-15FIGURE 4.1.3.1-5A: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 4) ....... 4-16FIGURE 4.1.3.1-5B: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 4) ....... 4-17FIGURE 4.1.3.1-5C: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 4)....... 4-18FIGURE 4.1.3.1-5D: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 4 OF 4)....... 4-19FIGURE 4.1.3.1-6A: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS

(SHEET 1 OF 3)............................................................................................................................................... 4-20FIGURE 4.1.3.1-6B: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS

(SHEET 2 OF 3)............................................................................................................................................... 4-21FIGURE 4.1.3.1-6C: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS

(SHEET 3 OF 3)................................................................................................................................................4-22FIGURE 4.2.1.6-1: DRY LOOP FUNCTIONAL SCHEMATIC ..................................................................................... 4-25FIGURE 4.2.1.7-1A: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BLOCK DM FAIRING (TYPICAL) ........ 4-32FIGURE 4.2.1.7-1B: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BREEZE M FAIRING (TYPICAL) ......... 4-33FIGURE 4.2.1.7-2: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1666 ADAPTER SYSTEM (TYPICAL)....... 4-34FIGURE 4.2.1.7-3: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1194 ADAPTER SYSTEM (TYPICAL) ...... 4-35FIGURE 4.2.1.7-4: INSTRUMENTATION DURING TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER......... 4-36FIGURE 4.2.1.7-5: INSTRUMENTATION DURING TRANSPORTATION OF SC IN KHSC CONTAINER....................... 4-36FIGURE 4.2.1.7-6: INSTRUMENTATION DURING TRANSPORTATION OF ASCENT UNIT ......................................... 4-37


Page xiv

FIGURE 4.2.1.7-7: INSTRUMENTATION DURING INTEGRATION OF ASCENT UNIT TO LV..................................... 4-37FIGURE 4.2.1.7-8: INSTRUMENTATION DURING TRANSPORTATION OF INTEGRATED PROTON LV....................... 4-38FIGURE 4.2.1.7-9: INSTRUMENTATION DURING ON-PAD OPERATIONS .............................................................. 4-39FIGURE 4.2.2-1: SPACECRAFT TO LAUNCH VEHICLE AND GROUND SYSTEMS ELECTRICAL INTERFACES............. 4-40FIGURE 4.2.3-1: SC TO BUNKER RF/ELECTRICAL INTERFACE BLOCK DIAGRAM ............................................... 4-46FIGURE 5.1.3-1A: BASELINE INTEGRATION SCHEDULE (NON-RECURRING PROGRAM)........................................ 5-2FIGURE 5.1.3-1B: BASELINE INTEGRATION SCHEDULE (RECURRING PROGRAM) ................................................. 5-3FIGURE 6.1-1: BAIKONUR FACILITIES MAP............................................................................................................ 6-2FIGURE 6.2.1.1-1: BUILDING 92A-50 GENERAL ARRANGEMENT .......................................................................... 6-7FIGURE 6.2.1.4-1: BUILDING 92A-50 SPACECRAFT PROCESSING AND FUELING AREA ....................................... 6-10FIGURE 6.2.2-1: SPACECRAFT PROCESSING ZONE (AREA 31) .............................................................................. 6-15FIGURE 6.2.2.2-1: BUILDING 40, HALL 100 (COMMON HALL) LAYOUT............................................................... 6-16FIGURE 6.2.2.3-1: BUILDING 40D, FIRST FLOOR LAYOUT................................................................................... 6-17FIGURE 6.2.2.4-1: BUILDING 40D, SECOND FLOOR LAYOUT............................................................................... 6-19FIGURE 6.2.2.4-2: BUILDING 40D, THIRD FLOOR LAYOUT .................................................................................. 6-19FIGURE 6.2.3.1-1: BUILDING 44 (FILLING HALL) FLOOR PLAN .......................................................................... 6-21FIGURE 6.5.2.1-1: BUILDING 92-1....................................................................................................................... 6-24FIGURE 6.3.1-1: PROTON LAUNCH ZONE, AREA 81.............................................................................................. 6-25FIGURE 6.3.2.2-1: PROTON MOBILE SERVICE TOWER .......................................................................................... 6-27FIGURE 7.2.1-1: ORGANIZATION DURING LAUNCH CAMPAIGN ............................................................................ 7-3FIGURE 7.3-1: COUNTDOWN ORGANIZATION ........................................................................................................ 7-5FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50................................. 7-8FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50 (CONTINUED) ......... 7-9FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 .................................................. 7-10FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 (CONTINUED)........................... 7-11FIGURE 7.7-1: TYPICAL LAUNCH PAD OPERATIONS TIMELINE ............................................................................ 7-17FIGURE 8-1: PROTON EVOLUTION OPTIONS........................................................................................................... 8-2FIGURE 8.1-1: PROTON LARGER PAYLOAD FAIRING CONCEPTS ............................................................................ 8-4FIGURE 8.2-1: BREEZE-M LAUNCH CONFIGURATION WITH TANDEM LAUNCH SYSTEMS (TLS).......................... 8-6FIGURE A.1-1: PROTON LAUNCH VEHICLE FAMILY ...............................................................................................A-1FIGURE C3.1-1: SC PENDULUM MODEL ...........................................................................................................C3-12FIGURE C3.1-2: SC SLOSH PROPERTIES DURING BALLISTIC FLIGHT AND AT SEPARATION ...............................C3-13FIGURE C3.1-3: PROPELLANT TANK GEOMETRY REQUIRED DATA ...................................................................C3-15FIGURE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D1-2FIGURE D1.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D1-4FIGURE D1.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM

ADAPTER ...................................................................................................................................................... D1-6FIGURE D1.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM

ADAPTER (SHEET 1 OF 4).............................................................................................................................. D1-7FIGURE D1.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM

ADAPTER (SHEET 2 OF 4).............................................................................................................................. D1-8FIGURE D1.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM

ADAPTER (SHEET 3 OF 4).............................................................................................................................. D1-9FIGURE D1.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 500MM

ADAPTER (SHEET 4 OF 4)............................................................................................................................ D1-10


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FIGURE D1.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 37 PIN ELECTRICAL

CONNECTOR ............................................................................................................................................... D1-12FIGURE D1.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D1-13FIGURE D2.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D2-2FIGURE D2.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D2-4FIGURE D2.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM

ADAPTER ...................................................................................................................................................... D2-6FIGURE D2.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM

ADAPTER (SHEET 1 OF 4).............................................................................................................................. D2-7FIGURE D2.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM

ADAPTER (SHEET 2 OF 4).............................................................................................................................. D2-8FIGURE D2.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM

ADAPTER (SHEET 3 OF 4).............................................................................................................................. D2-9FIGURE D2.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1194AX X 625MM

ADAPTER (SHEET 4 OF 4)............................................................................................................................ D2-10FIGURE D2.3-2A: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER ....................................................................................................................... D2-11FIGURE D2.3-2B: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 1 OF 4) (TBC) ................................................................................... D2-12FIGURE D2.3-2C: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 2 OF 4) (TBC) ................................................................................... D2-13FIGURE D2.3-2D: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 3 OF 4) (TBC) ................................................................................... D2-14FIGURE D2.3-2E: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 4 OF 4) (TBC) ................................................................................... D2-15FIGURE D2.3-3A: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER ....................................................................................................................... D2-16FIGURE D2.3-3B: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 1 OF 4) (TBC) ................................................................................... D2-17FIGURE D2.3-3C: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 2 OF 4) (TBC) ................................................................................... D2-18FIGURE D2.3-3D: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 3 OF 4) (TBC) ................................................................................... D2-19FIGURE D2.3-3E: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH

1194AX X 625MM ADAPTER (SHEET 4 OF 4) (TBC) ................................................................................... D2-20FIGURE D2.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 37 PIN ELECTRICAL

CONNECTOR ................................................................................................................................................D2-22FIGURE D2.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D2-23FIGURE D3.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D3-2FIGURE D3.2.2-1: CAPABILITY OF 1666V ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D3-4FIGURE D3.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM

ADAPTER ...................................................................................................................................................... D3-6FIGURE D3.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM

ADAPTER (SHEET 1 OF 4).............................................................................................................................. D3-7FIGURE D3.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM

ADAPTER (SHEET 2 OF 4).............................................................................................................................. D3-8


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FIGURE D3.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM

ADAPTER (SHEET 3 OF 4).............................................................................................................................. D3-9FIGURE D3.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1666V X 1000MM

ADAPTER (SHEET 4 OF 4)............................................................................................................................ D3-10FIGURE D3.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 61 PIN ELECTRICAL

CONNECTOR ............................................................................................................................................... D3-12FIGURE D4.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D4-2FIGURE D4.2.2-1: CAPABILITY OF 1666A ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D4-4FIGURE D4.3-1A: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150

ADAPTER ...................................................................................................................................................... D4-6FIGURE D4.3-1B: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150

ADAPTER (SHEET 1 OF 4).............................................................................................................................. D4-7FIGURE D4.3-1C: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150

ADAPTER (SHEET 2 OF 4).............................................................................................................................. D4-8FIGURE D4.3-1D: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150

ADAPTER (SHEET 3 OF 4).............................................................................................................................. D4-9FIGURE D4.3-1E: USABLE VOLUME - PROTON/BLOCK DM COMMERCIAL FAIRING WITH 1166A X 1150

ADAPTER (SHEET 4 OF 4)............................................................................................................................ D4-10FIGURE D4.4-1: ELECTRICAL DISCONNECT FORCE PROFILE FOR A SINGLE 61 PIN ELECTRICAL

CONNECTOR ............................................................................................................................................... D4-12


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LIST OF TABLES

TABLE 1.6-1: PROTON K AND PROTON M CHARACTERISTICS COMPARISON........................................................... 1-8TABLE 2.2-1: SUMMARY PROTON PERFORMANCE (PSM) TO REFERENCE ORBITS ................................................ 2-1TABLE 2.2.2-1: ALLOWABLE LAUNCH ENVIRONMENT CONSTRAINTS ..................................................................... 2-3TABLE 2.3.1-1: TYPICAL BOOSTER ASCENT EVENT TIMES ....................................................................................... 2-5TABLE 2.3.2-1: TYPICAL BLOCK DM ATTITUDE MANEUVERS FOR GEOSYNCHRONOUS MISSION (90O EAST

LONGITUDE) ..................................................................................................................................................2-12TABLE 2.4.1-1: LAUNCH VEHICLE MISSION PECULIAR HARDWARE ..................................................................... 2-14TABLE 2.5-1: PROTON BOOSTER PERFORMANCE TO LOW EARTH ORBITS (DIRECT INJECTION, NO

UPPER STAGE)................................................................................................................................................ 2-15TABLE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER

ORBITS............................................................................................................................................................2-18TABLE 2.6.1-2: PROTON K/BLOCK DM THREE-BURN MISSION PERFORMANCE TO REPRESENTATIVE

GEOSYNCHRONOUS TRANSFER ORBITS ..........................................................................................................2-18TABLE 2.6.1-3: PROTON K/BLOCK DM PARAMETRIC GEOSYNCHRONOUS TRANSFER PERFORMANCE DATE

(2 BURN MISSION) .........................................................................................................................................2-20TABLE 2.6.1-4: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER ORBITS

.......................................................................................................................................................................2-21TABLE 2.6.1-5: PROTON M/BREEZE M PARAMETRIC GEOSYNCHRONOUS TRANSFER PERFORMANCE DATA

(CONFIGURATION 3).......................................................................................................................................2-22TABLE 2.7-1: BLOCK DM UPPER STAGE ORBIT INJECTION ACCURACIES .............................................................2-23TABLE 2.7-2: BREEZE M UPPER STAGE ORBIT INJECTION ACCURACIES ..............................................................2-23TABLE 2.8-1: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION I................................................................2-24TABLE 2.8-2: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION II ...............................................................2-24TABLE 2.8-3: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION III..............................................................2-24TABLE 2.9-1: FORMAT I - PRELIMINARY STATE VECTOR DATA PROVIDED FOLLOWING UPPER STAGE

1ST BURN.........................................................................................................................................................2-26TABLE 2.9-2: FORMAT II - TRANSFER ORBIT PARAMETERS FOLLOWING UPPER STAGE 1ST BURN........................2-26TABLE 2.9-3: FORMAT III - PRELIMINARY VECTOR DATA AT SEPARATION EPOCH ...............................................2-27TABLE 2.9-4: FORMAT IV - VECTOR DATA AT SEPARATION EPOCH.......................................................................2-27TABLE 2.9-5: FORMAT V - STATE VECTOR DATA AT SEPARATION EPOCH .............................................................2-28TABLE 2.9-6: SPACECRAFT SUPPLIED SEPARATION DATA .....................................................................................2-28TABLE 3.1.1-1: 3σ AMBIENT TEMPERATURES AT THE BAIKONUR COSMODROME................................................... 3-1TABLE 3.1.1-2: SPACECRAFT THERMAL ENVIRONMENT......................................................................................... 3-2TABLE 3.1.4-1: GROUND HUMIDITY ENVIRONMENT ............................................................................................. 3-7TABLE 3.2.1-1: GROUND CONTAMINATION ENVIRONMENT ................................................................................... 3-8TABLE 3.4.1.1-1: GROUND LIMIT QUASISTATIC LOAD ENVIRONMENT-TRANSPORTATION AND HANDLING

OPERATIONS .................................................................................................................................................. 3-10TABLE 3.4.4-1: SHOCK LOADS DURING TRANSPORTATION IN SC CONTAINER..................................................... 3-21TABLE 3.4.5.1-1: ACOUSTIC TEST REQUIREMENTS .............................................................................................. 3-23TABLE 3.4.5.2-1: SINE TEST REQUIREMENTS ....................................................................................................... 3-23TABLE 3.5.2-1: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON K/BLOCK DM (TBC)............................ 3-25TABLE 3.5.2-2: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON M (TBC)............................................... 3-25TABLE 3.5.2-3: LAUNCH VEHICLE RF CHARACTERISTICS FOR BREEZE M (TBC) ............................................... 3-26TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS ........................ 4-27


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TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS (CONTINUED) ..4-28TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (TYPICAL).................. 4-29TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-30TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-31TABLE 4.2.3-1A: C-BAND RF LINK CHARACTERISTICS......................................................................................... 4-42TABLE 4.2.3-1B: KU-BAND RF LINK 1 CHARACTERISTICS .................................................................................. 4-43TABLE 4.2.3-1C: K-BAND RF LINK 2 CHARACTERISTICS..................................................................................... 4-44TABLE 4.2.3-1D: KU-BAND RF LINK 3 CHARACTERISTICS .................................................................................. 4-45TABLE 5.1.5.1-1A: BASELINE MEETING SCHEDULE FOR NON-RECURRING PROGRAM ......................................... 5-5TABLE 5.1.5.1-1B: BASELINE MEETING SCHEDULE FOR RECURRING PROGRAM .................................................. 5-6TABLE 5.2-1: ILS DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM................... 5-9TABLE 5.2.2-1: DESIGN REVIEW ANALYSES .......................................................................................................... 5-10TABLE 5.3-1: CUSTOMER DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM..... 5-12TABLE A.1-2: PROTON 50-LAUNCH MOVING AVERAGE .........................................................................................A-2TABLE A.2-1: PROTON LAUNCH RECORD SUMMARY (1970-1998).........................................................................A-3TABLE A.3-1: PROTON LAUNCH HISTORY ..............................................................................................................A-4TABLE A.3-1A: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-5TABLE A.3-1B: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-6TABLE A.3-1C: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-7TABLE A.3-1D: PROTON LAUNCH HISTORY (CONTINUED) ....................................................................................A-8TABLE A.3-1E: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-9TABLE C3.1-1: SC MASS PROPERTIES ..................................................................................................................C3-9TABLE C3.1-1A: SC MASS PROPERTIES NEAR 0G.................................................................................................C3-9TABLE C3.1-1B: SC MASS PROPERTIES NEAR 1G.................................................................................................C3-9TABLE C3.1-1C: SC MASS PROPERTIES(DRY SPACECRAFT) ..............................................................................C3-10TABLE C3.1-2: DESCRIPTION OF LIQUID MASSES..............................................................................................C3-11TABLE C3.2-1: SC RF CHARACTERISTICS ..........................................................................................................C3-16TABLE C3.2-2A: LVIJ1 UMBILICAL PIN ASSIGNMENTS......................................................................................C3-18TABLE C3.2.2B: LVIJ2 UMBILICAL PIN ASSIGNMENTS.......................................................................................C3-20TABLE C3.5-1: EGSE DESCRIPTION..................................................................................................................C3-22TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-23TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-24TABLE C3.5-2: FLUIDS/GASES REQUIREMENTS.................................................................................................C3-25TABLE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D1-1TABLE D1.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D1-11TABLE D1.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D1-14TABLE D2.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D2-1TABLE D2.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D2-21TABLE D2.5-1: STANDARD ELECTRICAL CONNECTORS..................................................................................... D2-24TABLE D3.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D3-1TABLE D3.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D3-11TABLE D3.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D3-13TABLE D4.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D4-1TABLE D4.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D4-11TABLE D4.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D4-13


Page xix

ABBREVIATIONS AND ACRONYMS

A CLA Coupled Loads Analysis

A Ampere CLS Commercial Launch Services

Ac Alternating Current cm Centimeter

A/C Air Conditioning CRES Corrosion-Resistant Steel

ADJ Attach, Disconnect, and Jettison CSO Complex Safety Officer

AGE Aerospace Ground Equipment CT Command Transmitter

ASCII American Standard Code for InformationInterchange

CV Computervision

ASTR Air System, Thermal Regulation CW Continuous Wave

atm Atmospheres CWA Controlled Work Area

ATS Advanced Technology Satellite CX Complex

AU Ascent Unit DAWG American Wire Gage DAS Data Acquisition System

B DAT Digital Auto Tape

Batt Battery dB Decibel(s)

BOD Board of Directors DB Design Bureau

BOL Beginning-of-Life dBm Decibel(s) Relative to 1 Milliwatt

bpi Bit(s) per Inch dBW Decibel(s) Relative to 1 Watt

BPSK Binary Phase Shift Key dc direct current

Btu British Thermal Unit DEC Digital Equipment Corporation

C DLF Design Load Factor

OC Degree(s) Celsius DOF Degree(s) of Freedom

CAB Customer Awareness Board DSN Deep Space Network

CAD Computer-Aided Design Dstr Destructor

CCAM Collison and Contamination AvoidanceManuever

DTSA Defense Technology Security Administration

CCSCS Central Control Station of CommunicationSystem

DUF Dynamic Uncertainty Factor

CCTV Closed-Circuit Television ECDR Critical Design Review e Eccentricity

CERT Composite Electrical Readiness Test ECA Environmentally Controlled Area

cg Center of Gravity ECS Environmentally Controlled System

CIB Change Integration Board EED Electro-Explosive Device

CIS Commonwealth of Independent States EHF Extreme High Frequency


Page xx

EIA Electronics Industry Association ft Foot; Feet

ELAN Electronic Library for Analysis ofNonconformances

FTP Filt Transfer Protocol

EM Electromagnetic FTS Flight Termination System

EMC Electromagnetic Compatibility GEMI Electromagnetic Interference g Gravity or Gram

EMK Extended Mission Kit GCS Guidance Commanded Shutdown

EOL End-of-Life GEO Geosynchronous Orbit

EPF Extended Payload Fairing GG Gas Generator

EPT External Propellant Tank (Breeze M) GHe Gaseous Helium

ERB Engineering Review Board GMM Geometric Mathematical Model

ESABASE Thermal Control Geometric Math ModelingCode

G&N Guidance and Navigation

EUTELSAT European Telecommunications SatelliteOrganization

GN&C Guidance, Navigation, and Control

EVCF Eastern Vehicle Checkout Facility GN2 Gaseous Nitrogen

EWR Eastern/Western Range Regulation GOWG Ground Operations Working Group

ε Emissivity GSE Ground Support Equipment

F GSO Geostationary Orbit

OF Degree(s) Fahrenheit GTO Geosynchronous Transfer Orbit

FAB Final Assembly Building GTR Gantry Test Rack

FAST Flight Analogy Software Test GTS Ground Telemetry Station

FCDC Flexible Confined Detonating Cord GTV Ground Transport Vehicle

FCS Flight Control Subsystem HFDLC Final Design Loads Cycle HAIR High Accuracy Instrumentation Radar

FLSC Flexible Linear-Shaped Charge HEO High-Energy (High-Eccentricity) Orbit

FM Flight Model Frequency Modulation HEPA High-Efficiency Particulate Air

FMAR Final Mission Analysis HiGTO High Energy Geoynchrenous Transfer Orbit

FMEA Failure Modes and Effects Analysis HP Hewlett-Packard

FMH Free Molecular Heating HPF Hazardous Processing Facility

FOD Foreign Object Damage hr Hour(s)

FOM Figure of Merit Hz Hertz

FOTS Fiber-Optics Transmission System IFPA Flight Plan Approval ICD Interface Control Document

FPR Flight Performance Reserve ICWG Interface Control Working Group

FSO Flight Safety Officer I/F Interface


Page xxi

IFR Inflight Retargeting LHe Liquid Helium

IGES Initial Graphics Exchange Specification LKE Lockheed Khrunichev Energia

IIP Instantaneous Impact Point LKEI Lockheed Khrunichev Energia International

ILC Initial Launch Capability LM Lockheed Martin

ILS International Launch Services LMC Lockheed Martin Corporation

IMS Inertial Measurement Subsystem LMCLS Lockheed Martin Commercial LaunchServices

in. Inch(es) LN2 Liquid Nitrogen

INTELSAT International Telecommunications SatelliteOrganization

LO2 Liquid Oxygen

INU Inertial Navigation Unit LOB Launch Operations Building

IRD Interface Requirements Document LR Load Ratio

ISA Interstage Adapter LSC Linear Shaped Charge

ISD Interface Scheduling Document LSTR Liquid System, Thermal Regulation

ISP Specific Impulse LV Launch Vehicle

ITA Intregrated Thermal Analysis LVMP Launch Vehicle Mission Peculiar

ITAR International Traffic in Arms Regulations MK m Meter

kg Kilogram(s) mA Milliamps

KhSC Khrunichev State Research and ProductionSpace Center

MDRD Module-Level Design RequirementsDocument

kHz Kilohertz MECO Main Engine Cutoff

km Kilometer(s) MES Main Engine Start

kN Kilonewton(s) MGSE Mechanical Ground Support Equipment

kPa Kilopascal(s) MHz Megahertz

kV Kilovolt(s) MICD Mechanical Interface Control Drawing

L MIL-STD Military Standard

LAE Liquid Apogee Engine min Minute(s)

LAN Longitude of Acending Node MLV Medium Launch Vehicle

lb Pound(s) mm Millimeter(s)

lbf Pound(s)-Force MMH Monomethyl Hydrazine

lbm Pound Mass MON-3 Mixed Oxides of Nitrogen

LC Launch Conductor MOTR Multiple Object Tracking Radar

LCC Launch Control Center MOU Memorandum of Understanding

LEO Low-Earch Orbit MPDR Mission-Peculiar Design Review

LH2 Liquid Hydrogen MR Mixture Ratio


Page xxii

MRB Material Review Board PMRR Mission Readiness Review Pa Pascal

MRS Minimum Residual Shutdown PA Public Address

MS&PA Mission Success and Product Assurance PCOS Power Changeover Switch

MSPSP Mission System Prelaunch Safety Package PDLC Preliminary Design Loads Cycle

MST Mobile Service Tower PDR Preliminary Design Review

MT Metric Ton PDRD Program-Level Design RequirementsDocument

m/s Meters Per Second PFJ Payload Fairing Jettison

µV Microvolt(s) PFM Protoflight Model

mV Millivolt(s) PHA Preliminary Hazard Analysis

MWG Management Working Group PHSF Payload Hazardous Servicing Facility

N PLCP Propellant Leak Contingency Plan

N Newton(s) PLCU Propellant Loading Control Unit

NCAR National Center for Atmospheric Research PLF Payload Fairing

N2H4 Hydrazine PMPCB Parts, Materials, and Processes Control Board

NIOSH National Institute of Occupational Safety andHealth

PMR Preliminary Material Review

NM Not Measured POD Program Office Directive

NM Newton Meter PPF Payload Processing Facility

nmi Nautical Mile(s) ppm Parts Per Million

N2O4 Nitrogen Tetroxide psi/psf Pound(s) per Square Inch/ Pound(s) perSquare Foot

ns Nanosecond(s) psig Pound(s) per Square Inch, Gage

NVR Nonvolatile Residue PSM Payload Systems Mass

O PSS Payload Separation System

O2 Oxygen PST Product Support Team

OASPL Overall Sound Pressure Level PSW Payload Systems Weight

OCC/CSS

Operator Control Console/Computer Subsystem PSWC Payload Systems Weight Capability

Oct Octave(s) PTC Payload Test Conductor

OIS Orbit Insertion Stage Payload Transport Canister

ORD Operational Requirements Document PU Propellant Utilization

OTM Output Transformation Matrix PVA Perigee Velocity Augmentation

Ω Ohm(s) PVC Polyvinyl Chloride

ωP Argument of Perigee P&W Pratt & Whitney


Page xxiii

Q SIU Servo-Inverter Unit

q Dynamic Pressure SLC Space Launch ComplexSpacecraft Launch Conductor

QA Quality Assurance SOZ SOZ Unit or Auxiliary Control System

QIC Quarter-Inch Cartridge SPA Spacecraft Processing Area

R SPRB Space Program Reliability Board

RAAN Right Ascension of Ascending Node SQEP Software Quality Evaluation Plan

RCS Reaction Control System SQP Sequential Quadratic Programming

R&D Research and Development SRB Solid Rocket Booster

RDX Research Department Explosive SRF Strategic Rocket Forces

RF Radio Frequency SRM Solid Rocket Motor

RM Room SRPSC State Research and Production Space Center

Roi Return on Investment SRR System Requirements Review

RP-1 Rocket Propellant 1 (Kerosene) Sta Station

RSA Russian Space Agency STC Satellite Test Center

RSC Rocket Space Complex STDN Spaceflight Tracking and Data Network

RSC Range Safety Console STM Structural Test Model

RSF Russian Space Force STS Space Transportation System

RTS Remote Tracking Station SVPF Space Vehicle Processing Facility

S Ts Second(s) tar Tape Archive (File Format)

S/A Safe and Arm TBD To Be Determined

SAEF Spacecraft Assembly and Encapsulation Facility TBS To Be Supplied

SASU Safe/Arm and Securing Unit T&C Telemetry & Command

SC Spacecraft TC Telecommand

SCAPE Self-Contained Atmospheric ProtectiveEnsemble

Test Conductor

SCA Spacecraft Adapter TCD Terminal Countdown Demonstration

SDP Software Documentation Plan TCO Thrust Cutoff

SDRC Structural Dynamics Research Corporation TDRSS Tracking and Data Rela Satellite System

sec Seconds TIM Technical Interchange Meeting

SEPP Systems Effectiveness Program Plan Tlm Telemetry

SFC Spacecraft Facility Controller TLS Tandem Launch System

SFTS Secure Flight Termination System TMM Thermal Mathematical Model

SHA System Hazard Analysis TRM Tension Release Mechanism

SHU Space Head Unit TSB Technical Support Building

SIL Systems Integration Laboratory TVC Thrust Vector Control


Page xxiv

TVCF Transportable Vehicle Checkout Facility

TWG Technical Working Group

3-D Three-Dimensional

UUDMH Unsymmetrical Dimethylhydrazine

UHF Ultra-High Frequency

UPS Uninterruptable Power Systems

U.S. United States

US Upper Stage-Proton Fourth Stage

UT Umbilical Tower

VV Volt(s) or Velocity

Vac Volt(s) Alternating Current

Vdc Volt(s) Direct Current

VDD Version Description Document

VHF Very High Frequency

VSWR Voltage Standing-Wave Ratio

VTF Vertical Test Facility

W

W Watt(s)

WB West Bay

WDR Wet Dress Rehearsal

W/M2 Watts Per Square Meter

WSR Weather Service Radar

X

Xdcr Transducer


Page 1-1

1. PROTON LAUNCH SERVICES

International Launch Service’s (ILS) launch support services equal or exceed in comprehensiveness and quality thoseservices available from any other commercial launch service organization. These services include system integration,supply of the Proton launch vehicle, custom engineering services and mission analysis, insurance brokering, groundand air transport, Government export license assistance, launch site spacecraft integration and testing, spacecraft andlaunch vehicle integration, launch site support and security services, launch of the spacecraft, and post launch missionsupport (see Figure 1-1).

The management approach to the provision of these services has been developed to ensure efficient task completion,with essential focus on Customer satisfaction. ILS functions as a prime contractor to manage all tasks associated withthe supply of the launch vehicle and associated spacecraft launch services, including all required liaison with variousUnited States, Russian, and other government organizations and agencies, as well as accommodation of any specialCustomer requirements. ILS will support all spacecraft preparation activities, oversee the integration of the spacecraftwith the Proton, and conduct the spacecraft launch. The Customer need interact solely with ILS for full support in allaspects of the launch.

Figure 1-1: ILS Launch Services Charter

LockheedMartinCorporation

KhrunichevSRPSC

RSCEnergia

ILS

System engineering Mission planning Interface data Custom engineering Insurance brokering Export license support Air transport brokering Host services and logistics support Test planning/requirements Spacecraft Integration and Test Site GSE and consumables Launch vehicle/adapter/fairing Launch vehicle integration Payload launch Post launch data analysis


Page 1-2

1.1 CONSTITUENT ILS COMPANIES

In 1992, the Space Systems Division of Lockheed Missiles & Space Co., Inc. established the Lockheed CommercialSpace Company, Inc. (U.S.). This company, along with the Russian organizations now known as the Khrunichev StateResearch and Production Space Center and the Rocket Space Complex Energia, signed an agreement to form a jointventure, Lockheed-Khrunichev-Energia International, Inc. (LKE). This venture, which has since become the ProtonDivision of ILS, which in turn reports to the Astronautics Division of Lockheed Martin Corporation (LMC), has thecharter to provide commercial launch services based on the Proton launch vehicle to customers worldwide. TheCorporate parentage of ILS is shown in Figure 1.1-1.

Figure 1.1-1: ILS Corporate Parentage

LMC Khrunichev Energia

LMAstronautics

ILS

This joint venture has been approved by both the U.S. and Russian governments. Lockheed Martin is the United Statesleading company in all aspects of spacecraft/launch vehicle integration, launch vehicle design, development andmanufacture, and launch site operations. Khrunichev (along with its subsidiary, the Salyut Design Bureau) is theRussian designer and manufacturer of the first stage, second stage, and third stage of the Proton space launch vehicle aswell as the Breeze-M and other upper stages. Energia is the Russian designer and manufacturer of the Block DMfourth stage of the Proton, and is the largest producer of launch vehicles in Russia. ILS has access to all resources of theconstituent companies required to fulfill spacecraft launch campaign requirements.

Lockheed Martin, Khrunichev, and Energia all function as subcontractors, reporting to ILS, during the execution of aCustomer’s launch services contract. A memorandum of understanding (MOU) that delineates the responsibilities ofeach of the companies is in place. (Figure 1.1-2) Constituent senior management of each of the companies haveapproved the overall management approach for ILS, and are each represented on the ILS Board of Directors.

The personnel, hardware resources, and facilities needed to support customer launch programs are in place and readyfor immediate activation, as needed.


Page 1-3

ILS personnel are responsible for all Customer interface activities, and for the coordination of all activities of theconstituent companies so that contract objectives are met. They are also responsible for all program systemengineering, custom engineering, mission analysis, program integration and program management, liaison with allgovernment agencies, and liaison with the world’s financial and insurance markets. Khrunichev is responsible formanufacturing the Proton first, second, and third stages, manufacturing the Proton’s Breeze M fourth stage,conducting mission analysis, and providing Baikonur services. Energia is responsible for manufacturing the Proton'sBlock DM fourth stage, providing the Block DM timeline, dynamics, and other data for mission analysis, andproviding Baikonur services.

Figure 1.1-2: ILS/LKE Task Partitions

ILS/LKE

Lockheed MartinCorporation

Khrunichev(KhSC) RSC Energia

Customerliaison/reportSubcontract directionU.S. Government liaison

Launch program integrationSystem managementInsurance liaisonCustomer reportingMission analysisThird party liaisonPost launch servicesPerformance data liaisonInterface data liaisonCustom engineering servicesLaunch processing liaison

First, second, and third stage productionPreparation of stages for launchProton-induced environment design criteriaInterface servicesLaunch vehicle GSE, consumables,and launchservicesCoordination of Baikonur Cosmodrone servicesCIS government liaisonSubcontract managementBreeze M production

Block DM productionBlock DM preparation services for launchProton spacecraft interfacedesign criteriaProton-induced environment designcriteriaRussian subcontract managementInterface servicesLaunch vehicle GSE,consumables, and launch servicesSubcontract management

1.2 ILS CONSTITUENT COMPANY EXPERTISE

The joint venture of Lockheed Martin Corporation, Khrunichev and Energia offers the most capable, comprehensive,and cost-effective spacecraft launch services in the world.

ILS, with access to the technological expertise and resources of Lockheed Martin Corporation, Khrunichev andEnergia, can provide all necessary resources required to support a spacecraft launch. ILS provides spacecraft Customerswith a single point of contact for all mission engineering and launch support tasks using the Proton launch vehicle.


Page 1-4

The constituent companies of ILS have more spacecraft and launch vehicle expertise than any other organizationsproviding launch services today. (Figure 1.2-1) Lockheed Martin has built more than half of all Western satellitesflown, and has designed, built and launched the Atlas and Titan launch vehicles, as well as the Agena and Centaurupper stages, and the Polaris, Poseidon and the Trident strategic defense missiles. Khrunichev, in addition to designingand building the Proton, is responsible for subcontract management for all of the launch vehicle subsystems.Khrunichev personnel participate in all Proton integration and launch operations at the Baikonur Cosmodrome.Khrunichev has also produced a variety of missile systems for the Russian government, as well as the Salyut, Mir, andAlmaz orbital stations, and several different orbital return capsule designs. Energia has provided all of the fourth stagesfor the Proton from the inception of the program until 1999, and has also been responsible for significant subcontractmanagement. Energia is the designer and builder of the R-7 series of space launch vehicles, including the Soyuz andMolniya launchers, of which more than 1300 have been launched to date. During the 1980’s they developed both theEnergia heavy lift launch vehicle and the Buran space shuttle, in addition to significant portions of the Mir spacestation hardware. The constituent companies of ILS have the necessary expertise to successfully and efficiently supportany launch campaign.

Figure 1.2-1: Launch Experience

1960s 1970s 1980s 1990s 2000s

1950s

Lockheed MartinTitan I, II, III, , and IV launchesAtlas launchesAgena flightsPolaris/Poseidon/Trident launchesSpace Shuttle launchesMission analysis and system engineering

Khrunichev200+ Proton launchesMission analysis and system engineeringSpacecraft/launch vehicle integration

RSC EnergiaSpunik, V ostok, Voskhod, Soyuz, Molniya launchesProton fourth stage productionSpacecraft/launch vehicle integration

1960s 1970s 1980s 1990s 2000s

1950s

1.3 ADVANTAGES OF USING THE PROTON LAUNCH VEHICLE

The use of the Proton launch vehicle provides several significant advantages that result in operational and revenuegenerating benefits for the Customer (Figure 1.3-1).

ILS provides full system engineering and mission analysis services and mechanical/electrical interface coordination.The proven procedures for Proton launch vehicle operations, backed by personnel with extensive experience in theseprocedures, provide for efficient and trouble-free launch campaigns.

The considerable lift capability of the Proton, combined with the multiple restart capability of both of the fourthstages, provides the Customer with mission flexibility and maximized payload capacity to orbit. This results in uniquemission design options, including delivery of spacecraft directly to geostationary orbit by the Proton. Spacecraft apogeefuel may be dedicated to extended mission life, because inclination reduction and orbit circularization areaccomplished by the Proton’s 4th stage.


Page 1-5

The ability of ILS to meet the full range of Customer requirements guarantees each launch campaign will beconducted to the Customer’s satisfaction. This approach to satisfying key requirements is shown in Figure 1.3-1. ILS’soverriding concern in providing commercial launch services is the careful coordination of our company’s resources andcapabilities with the Customer’s detailed requirements, so as to allow ILS to tailor each launch campaign to meet theindividual Customer’s unique needs. This ensures that each campaign will proceed in an efficient manner toward asuccessful, on-time launch to the precise orbit required.

Figure 1.3-1: Benefits To The Spacecraft Designer And Owner

Full mission analysis support by all constituent companies

Access to proven launch vehicle manufacturing and support capability

Use of very capable integration and launch support infrastructure

Use of the massive lift capacity of the Proton launch vehicle

Use of the restart capability of the Proton fourth stage for final orbit insertion

1.4 VEHICLE DESCRIPTION

The Proton launch system is designed and manufactured by International Launch Services partners Khrunichev StateResearch and Production Space Center (KhSC) and Rocket Space Complex (RSC) Energia, both of Russia. Based onvehicles in production today, the Proton is one of the most capable commercial expendable launch systems in use,delivering intermediate and heavy spacecraft to a wide range of orbits from its launch base at the BaikonurCosmodrome in Kazakhstan.

The Proton vehicle used for commercial launches can be provided in either a 3-stage or 4-stage configuration to meetthe needs of Low Earth Orbit and High-Energy launch missions. The 3-stage versions of Proton, designated Proton-Kand Proton-M, are designed to lift very heavy spacecraft systems into Low Earth Orbit. The 4-stage version of Proton,using either the Block DM or the Breeze M upper stage, is designed to meet high-energy launch mission injectionrequirements such as those for the Geosynchronous Transfer Mission. Multiple payload fairing and adapter systems areavailable in order to accommodate most commercial satellite launch mission needs.

This section provides a general description of the Proton launch vehicle, summarizing the relevant capabilities of theProton-K/Block DM launch system and the Proton M/Breeze M launch system.


Page 1-6

Figure 1.4-1: Proton Launch Vehicles Flight Proven Hardware


Page 1-7

1.5 GENERAL DESCRIPTION OF THE PROTON FAMILY

The Proton is currently available as three-stage Proton K and Proton M models and as the four-stage Proton K/ BlockDM , Proton K/Breeze M, Proton M/Block DM, and Proton M/Breeze M models. A variety of supplemental orbitalpropulsion units, in a range of capabilities, can be used with either the three-stage or the four-stage Proton. Inaddition, there are multiple fairing designs presently qualified for flight, including "Standard Commercial PayloadFairings" developed specifically to meet the needs of Western customers. Tandem launch and multi-spacecraftdispenser capabilities are in development. Later sections provide more detailed information on these hardwareelements and their operation.

The lower three stages of the Proton are produced by the Khrunichev State Research and Production Space Center(KhSC) plant in Moscow. Production of the Breeze-M stage is conducted by KhSC. Production of the Block DMfourth stage is carried out by Russian Space Complex (RSC) Energia, also in Moscow. Production capacity for thecommercial Proton is approximately twelve vehicles per year.

General specifications for both versions of the Proton are similar. Overall height of the vehicle in either configurationis approximately 61 m (200 ft), while the diameter of the second and third stages, and of the first stage core tank, is 4.1m (13.5 ft). Maximum diameter of the first stage, including the outboard fuel tanks, is 7.4 m(24.3 ft). The Block DM fourth stage, when present, has an external diameter of 3.7 m (12.1 ft). The Breeze M, whenpresent, has a diameter of 4.1 m (13.5 ft). Total mass of the Proton at launch is approximately 651,500 kg (1,524,000lbm).

Maximum performance capability (Payload Systems Mass, or PSM) of the Proton in the configurations of principalinterest to Western customers, is as follows:

Proton K: Proton M:

LEO (200 km circ.): 19.76 metric tons 21.0 metric tons

Proton K/Block DM: Proton M/Breeze M:

GTO: 4.9 metric tons 5.5 metric tons

GSO: 1.9 metric tons 2.92 metric tons

These figures assume the use of the Standard Commercial Payload Fairings, and optimum event sequences. Specificperformance may differ from the performance cited, if:

a) Use is made of adapters or fairings differing in mass or other characteristics relevant to performance

b) Earlier fairing separation timing is found to be acceptable

c) Special modifications are made to the Proton or its operations


Page 1-8

1.6 PROTON THREE - STAGE BOOSTER

The Proton K and M are series-staged vehicles consisting of three stages. An isometric view of the Proton K and M,showing the relationships among the major hardware elements, is provided in Figure 1.6-1. All three stages usenitrogen tetroxide (N2O4) and unsymmetrical dimethylhydrazine (UDMH) as propellants. A summary comparison ofthe characteristics of the Proton K and Proton M launch systems are shown in Table 1.6-1.

A comparison of the characteristics of the Proton M/Breeze M with those of the Proton K/Block DM is shown inTable 1.6-1.

Table 1.6-1: Proton K And Proton M Characteristics Comparison

Launch Vehicle/Upper Stage Proton K/Block DM Proton M/Breeze M

Launch Vehicle Liftoff Mass, mt 700 700

Support orbit Payload Mass, mthcirc.= 200 km, I = 51,6O

20.7 22.0

Geostationary Orbit Payload Mass, mthcirc.= 36000 km, I = 0O

Max. 1.9-2.1 Max. 3.0

Geostationary Transfer Orbit Payload Mass, tha = 36000 km, I = 7…51,6O

3.5…6.5 5.0…7.8

Payload Bay Volume, m3 65 (Std PLF) 100 (Std PLF)

Launch Vehicle Structure Mass, mtFirst StageSecond StageThird StageUpper (fourth) Stage

31.011.754.153.13

30.611.43.7

2.37

Propulsion System Performance(Maximum Vacuum Thrust), (mt)First Stage*Second StageThird StageUpper (fourth) Stage

1070237628.5

1070237622.0

Note: *Augmented First Stage Booster Engines have been used on the Proton Launch Vehicle since 1993.


Page 1-9

Figure 1.6-1: Proton K and M Major Hardware Elements

S e c o nd s tag e

4 .1 m ∅

First st age2 0 .2 m lo ng

7.4 m d ia. ∅S t ra p-on

fu el tank

1 .7 m ∅

(6) R D -2 5 3

C o r eox idizertan k

44

.3m

lon

g

(1 ) R D - 02 1 0

Th ird s tage

4 .1 m ∅

P a y lo a dad ap te r

P a y lo a dfair ing

( 4 ) - R D - 0210

T h e thi rd st age is equ ipp ed w ith on efixed s ing le-ch am b er liquid pro p ellan tro c k et en g ine develo pin g 0 .6 M N th ru s tan d o ne l iqu id pr o pellan t c o ntro l ro c k et

eng ine, w i th f ou r g im baled n o zz les,dev elopin g 30 k N thr us t

T h e s ec on d s tag e is equ ipp ed w ith f o ur

g im baled s in gle-c h am be r l iqu idpr o pel lant ro c k et eng ines dev elo pin g atotal th rus t o f 2 .3 M N

T h e fi rs t s tag e is equ ipp ed w ith six gi m b aledsin g le-c h am ber l iqu id p ro pel lant ro c k eteng in es dev elop in g a to tal thr u s t of 9 M N atlif t-of f


Page 1-10

1.6.1 Proton First Stage

The Proton first stage consists of a central tank containing the oxidizer, surrounded by six outboard fuel tanks.Although these fuel tanks give the appearance of being strap-on boosters, they do not separate from the core tankduring first stage flight. Each fuel tank also carries one of the six RD-253 engines that provide first stage power. Totalfirst stage sea-level thrust is approximately 9.5 MN (2.1 x 106 lbf). Total first stage dry mass is approximately 31,000kg (68,343 lbm); total first stage propellant load is approximately 419,410 kg (924,641 lbm).

The Proton M booster first stage improves on the current booster with a small reduction in overall stage mass. Massoptimization results from modern manufacturing techniques and reduction in avionics system mass.

In addition to the structual enhancement, the RD-253 first stage engines are uprated. For performance purposes, ratedthrust on the engines of the first stage of Proton M is being increased by 7%. This enhancement is accomplished withonly a minor modification to the propellant flow control valves. This modification first flew as a mission-uniqueenhancement on the Proton K that delivered the MIR space station core module in 1986. Engines incorporating thischange have undergone extensive additional qualification firings since then, in order to approve them for use instandard production vehicles. Shipsets have already been produced and incorporated onto production Proton Kvehicles. To date, 65 Proton K's have been launched with the modified engines operating at 102% of rated thrust.Eight Proton K vehicles have flown with the 107% engines. There have been no flight anomalies attributed to theincreased thrust engines.

The propellant feed systems of the first and second stages have been simplified and redesigned in order to reducepropellant residuals in these stages by 50%, and a propellant purge system has been added to dump all residuals fromthe spent first and second stages before they return to the earth's surface.

While a reduction in unusable propellants results in a performance gain, the primary rationale for the increasedutilization of propellants is to reduce the environmental effects of the impact of the first and second stages in thedownrange "land" jettison zones.

1.6.2 Proton Second Stage

The second stage, of conventional cylindrical design, is powered by four RD-0210 engines; it develops a vacuum thrustof 2.3 MN, or 5.24 x 105 1bf. Total second-stage dry mass is approximately 11,715 kg (25,827 lbm); total second stagepropellant load is approximately 156,113 kg (344,170 lbm).

Modifications to Proton second stage include structural reinforcement of the forward portion of the stage, in order tocarry greater payload and aerodynamic loads, and minor structural weight reductions.

1.6.3 Proton Third Stage

The third stage is powered by one RD-0210, developing 583 kN (1.31 x 105 Lbf) thrust, and a four-nozzle vernierengine that produces 31 kN (6.9 x 103 lbf) thrust. Total third-stage dry mass is approximately 4,185 kg (9,226 lbm);total third stage propellant load is approximately 46,562 kg (102,652 lbm).

Modifications to the Proton third stage, include structural reinforcement of the aft portion of the stage in order tocarry greater payload and aerodynamic loads, and minor structural weight reductions.


Page 1-11

1.6.4 Proton Flight Control System

Guidance, navigation, and control of the Proton K during operation of the first three stages is carried out by a tripleredundant closed-loop analog avionics system mounted principally in the Proton's third stage, using data collectedfrom a distributed set of inertial reference platforms. This system also provides for flight termination in the event of amajor malfunction during ascent.

For Proton M, the principal modification to the Proton K's first three stages is the incorporation of a digital flightcontrol system based on modern avionics technology. This digital system replaces the analog flight control hardware ofthe Proton K, which, although it has achieved a high demonstrated reliability, is based on obsolete 1960's eraelectronic designs. The new system eliminates the need to maintain a unique and limited production capability forProton K avionics, and allows for simplified control algorithm loading and test. It also enables greater ascent programdesign flexibility with respect to vehicle pitch profile and other parameters. The system is being developed by the Marsand Proton Control Systems Companies.

The self-contained inertial control system uses a precision three-axis gyro-stabilizer and a three-channel voting on-board digital computer. The digital control computer resides on the Proton M third stage and controls the flight of thefirst, second, and third stages.

1.7 BLOCK DM FOURTH STAGE

The Block DM fourth stage, normally uses liquid oxygen and kerosene, although the stage can be operated with severaldifferent kerosene-type fuels. The configuration of the Proton K/Block DM is depicted in Figure 1.7-1. The BlockDM is shown in cutaway, within its aerodynamic shrouds ("middle and lower adapters") in Figure 1.7-2, and as itappears in flight in Figure 1.7-3. The Block DM and its predecessor, the Block D, have together flown more than 200times.

The Block DM is optimized for multi-burn space transfer operations. Its main engine (model number 1lD58M)delivers a vacuum thrust of 83.5 kN (1.88 x 104 lbf), is gimbaled to provide three axis control during powered flightoperations, 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(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(33,180 lbm). It is three-axis stabilized in unpowered flight by a storable bipropellant (N2O4/UDMH) attitudecontrol system, comprised of two "SOZ" (or "micro") thruster units located at the base of the Block DM. The fourthstage can achieve a maximum of 1.5 rpm rotation at the moment of spacecraft separation. It should be noted, however,that the Block DM can only rotate approximately +/-180 degrees from a "reference" orientation, due to limitations ofthe stage's gyro platform; it cannot undergo continuous rotation. Guidance, navigation, and control of the fourth stageare provided by a triple redundant digital avionics package, which can be ground commanded in flight, if necessary.

The Block DM equipment bay provides the payload adapter and electrical interfaces to the customer’s spacecraft. Theinterface between the stage and its payload adapter is 2000 mm in diameter, allowing the Block DM to accommodatelarge diameter payload adapters and a static bending moment about this interface of 11,000 kg-m. The Block DMpayload fairing is a two piece composite structure developed specifically to meet the volume and environmentalrequirements of commercial spacecraft. Payload fairing usable volume geometry is provided in Appendix D of thisMission Planner’s Guide.


Page 1-12

Figure 1.7-1: Proton K/Block DM Major Hardware Elements

PayloadFairing

Payload

Adapter

Fourth StageBlock - DM

Forward fourthstage shroud

44

.3m

lon

g

Third stage4.1 m ∅

(1) RD - 0210

Aft fourthstage shroud

Second stage4.1 m ∅

(4) RD - 0210

Coreoxidizertank

Strap-onfuel tank1.7 m∅

(6) RD-253

First stage20.2m long7.4 dia m ∅

The third stage is equipped with onefixed single-chamber liquid propellantrocket engine developing 0.6 MNthrust and one liquid propellant controlrocket engine, with four gimbalednozzles, developing 30 kN thrust

The second stage is equipped with fourgimbaled single-chamber liquidpropellant rocket engines developing atotal thrust of 2.3 MN

The first stage is equipped with six gimbaledsingle-chamber liquid propellant rocketengines developing a total thrust of 9 MN

The fourth stage is equipped withone liquid propellant rocket enginedeveloping 86 kN thrust, and two"micro" engine clusters for attitudecontrol and ullage maneuvers

(1) 11D58M


Page 1-13

Figure 1.7-2: Proton DM Upper Stage Within Its Aerodynamic Shrouds

φ3700 m m

φ4100 m m

φ2000 m m

6280 mm

1

1

2

3

4

5

6

4900 mm

1

6

2

3

4

5

6

Plane of spacecraft adapter/

Block DM joint

1. Equipment bay

2. Oxidizer tank

3. Fuel tank

4. Auxilary propulsion for attitude control

("SOZ " units)5. Main engine

6. Middle and lower adapters (shrouds)


Page 1-14

Figure 1.7-3: Block DM As It Appears In Flight


Page 1-15

1.8 BREEZE M FOURTH STAGE

The Breeze M upper stage, which is derived from the Breeze K stage flown on the Rokot, offers substantially improvedpayload performance and operational capabilities over the Block DM flown on the current Proton K. The Breeze Mprogram was initiated in 1994 by the Khrunichev Space Center and has the full support of the Russian government. Anisometric view of the Proton M/Breeze M is shown in Figure 1.8-1.

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 atotal propellant mass of 19,800 kg. It consists of two main elements:

1) a core section (derived from the original Breeze K stage) that accommodates a set of propellant tanks, thepropulsion system, and the avionics equipment bay, and

2) a toroidal auxiliary propellant tank that surrounds the core section, and which is jettisoned in flight followingdepletion. Use of an external drop tank substantially improves the performance of the Breeze M stage.

Figures 1.8.-2 and 1.8-3 illustrates the layout and dimensions of the Breeze M.


Page 1-16

Figure 1.8-1: Proton M/Breeze M Major Hardware Elements

PayloadFairing

Payload

Adapter

Fourth StageBlock - DM

Forward fourthstage shroud

44

.3m

lon

g

Third stage4.1 m ∅

(1) RD - 0210

Aft fourthstage shroud

Second stage4.1 m ∅

(4) RD - 0210

Coreoxidizertank

Strap-onfuel tank1.7 m∅

(6) RD-253

First stage20.2m long7.4 dia m ∅

The third stage is equipped with onefixed single-chamber liquid propellantrocket engine developing 0.6 MNthrust and one liquid propellant controlrocket engine, with four gimbalednozzles, developing 30 kN thrust

The second stage is equipped with fourgimbaled single-chamber liquidpropellant rocket engines developing atotal thrust of 2.3 MN

The first stage is equipped with six gimbaledsingle-chamber liquid propellant rocketengines developing a total thrust of 9 MN

The fourth stage is equipped withone liquid propellant rocket enginedeveloping 86 kN thrust, and two"micro" engine clusters for attitudecontrol and ullage maneuvers

(1) 11D58M


Page 1-17

Figure 1.8-2: Breeze M (With Auxiliary Propellant Tank)

? 4 0 0 0 m m

? 24 9 0 m m

2 6 1 0 m m

1

2

3

4

5

6

7

1. E qu ip m en t B a y

2. C or e stage

3. A u x il ia ry pr opell an t tan k

4. O xidiz er ta nk

5. F u e l tan k

6. A u xil ia ry pr opul sio n fo r a tt it ude contr ol

7. M ai n e ng ine

Figure 1.8-3: The Breeze M And Breeze M Core As They Appear In Flight

The Breeze M uses nitrogen tetroxide (N2O4) and unsymmetrical-dimethylhydrazine (UDMH) as propellants.Propulsion for the Breeze M consists of one pump-fed, gimbaled main engine developing 19.62 kN of thrust, four"impulse adjustment thrusters" of 396 N thrust each for making fine corrections to the main engine impulse, and 12attitude control thrusters of 13.3 N thrust each. The main engine can relight up to eight times per mission, and isequipped with a backup restart system that can fire the engine in the event of a primary ignition sequence failure.


Page 1-18

During two flights of the Phobos space probes in 1988 and three flights of the Breeze K on the Rokot in 1990, 1991,and 1994 the main engine demonstrated up to five restarts in flight. Following minor modifications to adapt the enginefor the Breeze M, eleven main engines have been ground tested, some up to 6,000 seconds total burn duration. TheBreeze M attitude control thrusters were previously used on the Kvant, Kristall, Spektr, and Priroda modules of theMIR space station, and are used on the Russian FGB and Service Module components of the International SpaceStation.

The control system of the Breeze M includes an on-board digital computer and three axis inertial measurement unit.It also incorporates GLONASS and NAVSTAR GPS navigation systems. Breeze M can perform preprogrammedmaneuvers about all axes during parking orbit and transfer orbit coast. The stage is normally three axis stabilizedduring coast, but can be rotated at up to 2 degrees per second for thermal control. During powered flight the upperstage attitude is determined by mission specific pitch, yaw and roll programs. Breeze M can perform separation of acustomer's spacecraft in either one of two modes:

1) attitude hold mode, during which the angular rates in relation to any of the coordinate system axes will not exceed0.5 degrees per second, and the spatial attitude error in relation to the inertial coordinate system will not exceed 1degree, or

2) spin-up mode, during which the stage can achieve a maximum angular rate with respect to the upper stagelongitudinal axis of 30 degrees per second, and the spin axis deflection from the upper stage longitudinal axis willnot exceed 0.05 degrees.

The typical Proton M/Breeze M mission profile is discussed further in Section 3 of this Mission Planner’s Guide.

The Breeze core structure provides the payload adapter and electrical interfaces to the customer's spacecraft. Theinterface between the stage and its payload adapter is 2490-mm in diameter, allowing the Breeze M to accommodatelarge diameter payload adapters and a static bending moment about this interface of 18,000 kg-m. The Breeze M stageis encapsulated within the payload fairing, along with the customer's spacecraft, allowing loads from the payloadfairing to be borne by a short (600-mm) spacer ring attached to the Proton third stage equipment section, rather thanby the Breeze M. The Breeze M payload fairing is a derivative of the payload fairing currently in use with commercialspacecraft on the Proton K/Block DM. Modifications consist of a redefined attachment geometry at the aft end of thefairing; the attachment and separation hardware, however, is essentially unchanged from the current design. Payloadfairing usable volume geometry is provided in Appendix D of this Mission Planner’s Guide.

1.9 PAYLOAD FAIRINGS

Figure 1.9-1 illustrates the available payload fairings for the three stage Proton K and Proton M. These fairingstypically enclose both the payload of the Proton and any supplemental orbital propulsion units employed by thepayload. The payload fairing is jettisoned after second stage separation, with the exact time determined by spacecraftfree-molecular heating and impact zone constraints. Normally payload fairing separation occurs at approximately 344seconds into flight for commercial missions.

Figure 1.9-2 illustrates the standard commercial payload fairing for use with the Block DM upper stage, and either theProton K or the Proton M.


Page 1-19

Figure 1.9-3 shows the standard commercial payload fairing available for use with the Breeze M upper stage. Theusable volume of this fairing is comparable to that of the payload fairing for the Block DM. Figure 1.9-4 givesdimensions for the long version of the Breeze M payload fairing. Use of this fairing results in a decrease inperformance to GTO of approximately 100kg.

Figure 1.9-1: Proton K and M Payload Fairings

9047.5 mm

L = 2245 mm

250 mm

250 mm

30°

300 mm

150 mm

1000 mm

100 mm

φ4100 mm φ4100 mm

150 mm

1600 mm

2840 mm250 mm300 mm

100 mm

2150 mm

11970 mm

12650 mm

14562 mm15882 mm

3752 mm

3000 mm


Page 1-20

Figure 1.9-2: Block DM Payload Fairing for Single Spacecraft


Page 1-21

Figure 1.9-3: Breeze-M Payload Fairing (Standard) For Single Spacecraft


Page 1-22

Figure 1.9-4: Breeze M Payload Fairing (Long)


Page 1-23

1.10 BAIKONUR INFRASTRUCTURE AND SERVICES

All Proton launches occur from the Baikonur Cosmodrome, which is located in the Republic of Kazakhstan in centralAsia. Baikonur is approximately 2,000 km (1,300 miles) southeast of Moscow. Located in a semiarid zone, annualtemperature and climate vary considerably. The Proton launch system is designed to reliably operate from just such aclimatic range.

1.10.1 Baikonur Infrastructure

All primary launch systems infrastructure at Baikonur are being maintained in accordance with standard KhrunichevSpace Center (KhSC) and Russian Space Agency (formerly Russian Space Forces) policies concerning operationalefficiency and system safety. A number of facilities in use to support the Proton and satellite mission launch campaignprocess have been upgraded to support “Western-style” launch campaign processes. Of primary importance areupgrades to Launch Complex 24 at Baikonur and the addition of new, integrated payload processing facilities in closeproximity to the launch complex. Figure 1.10.1-1 shows the locations of the major facilities at the BaikonurCosmodrome.

Figure 1.10.1-1: Overall Layout Of The Baikonur Cosmodrome

6

Yubileini Airfield

SC Processing Area(Facility 92A-50) Proton LV

Launch Complex 81(Pads 23 and 24)

Proton LVProcessing Area

Area 95Living Quarters

DM Upper StageProcessing Area

DM Upper StageFueling Area

Kranyi Airfield

Baikonur Town(Leninsk)

Sir-Darya River

N

Area 31

Area 92

75

9

4

2

90 km

3

1-2 24 km2-3 18 km2-4 18 km4-5 16 km4-6 23.5 km6-7 1.25 km7-8 9 km8-9 7 km

1

75 km

Legend:RailroadRoad


Page 1-24

1.10.1.1 Yubileini Airport

All launch campaign personnel, ground support equipment, and the satellite arrive at Baikonur by air via the Yubileiniairport. Yubileini operates as a Category 3 airport facility capable of supporting all large transport aircraft. Airportinstrumentation and support services have been brought inline with western requirements with regard to aircraftoperations and aircraft onload/offload. The facility is capable of supporting operations in most weather conditions.

Upon offload of the spacecraft and support equipment, ground handling equipment is available to place palettes onto aproperly configured series of rail cars for transportation to the payload processing facility. The aircraft offload/rail caronload process can take up to 8 hours. ILS offers an environmentally controlled rail car for thermal conditioning of thesatellite transportation container.

1.10.1.2 Area 92 Launch Vehicle and Satellite Processing

Area 92 contains the major processing facilities for both the spacecraft and the Proton launch system. Building 92-1 isthe facility in which Proton vehicles are integrated and tested before mating with the fourth-stage/payload “ascentunit.” Building 92A-50 houses the new SVPF (Figure 1.10.1.2-1). The Space Vehicle Processing Facility is covered ingreater detail in Section 6 of this Mission Planner’s Guide. Area 92A-50 payload processing facilities areapproximately 20 km from airport facilities.

Figure 1.10.1.2-1: Building 92A-50 Space Vehicle Processing Facility


Page 1-25

1.10.1.3 Area 31 Payload Processing

Area 31 was configured to enable nonhazardous and hazardous processing of Western communications satellites tosupport the first commercial Proton launches. All ILS launches through the end of 1997 were processed in Area 31facilities, specifically Building 40D for nonhazardous processing and Building 44 for hazardous processing.

Following transition to the SVPF, Area 31 has been retained in a backup mode to support contingency and/orprocessing “surge” commercial requirements. ICD’s for commercial spacecraft should continue to be written tosupport this facility. Area 31 will continue to support Russian government payloads such as the SOYUZ mannedsystem.

1.10.1.4 Area 81 Launch Complex

Area 81 contains both active commercial Proton launch pads. Launch Pad 23 is presently launching commercialProton K/Block DM vehicles in support of ILS. Pad 24 has been refurbished to support launch operations for theProton K/Block DM system and the Proton M/Breeze M vehicle. Launch operations on Complex 24 will commencein mid-1999 after refurbishment and upgrades are complete. Complex 23 will support the ILS Proton business as abackup pad, as business requirements dictate.

1.10.1.5 Area 95 Hotel Facilities

ILS and our KhSC partners have spent significant effort and resources upgrading personnel and logistics-relatedfacilities. Area 95 houses the three primary hotel facilities used by ILS and western launch campaign teams. HotelsFili, Cometa and Polyot are the primary accommodations for satellite contractor personnel.

In addition to the physical accommodations, enhancements have been made to the cafeteria services and utilitiesinfrastructure to better accommodate western launch crews. A water treatment plant is in operations 24-hours per day,offering water purified to U.S. standards. Additional support services, such as recreation equipment, have also beenimplemented.

ILS typically arranges for the accommodation of up to 30 customer and satellite contractor personnel per spacecraftduring the Baikonur launch campaign process, with each spacecraft campaign expected to last approximately 30 days.

1.10.2 Proton Launch Campaign

For typical commercial Proton launches, ILS has developed a standard 30-day launch campaign schedule for purposesof initial discussions with potential Proton launch system users. As shown in Figure 1.10.2-1, a 30-day campaign isbased on the general philosophy that the satellite is shipped from the satellite manufacturer’s facility in a near “ship-and-shoot”condition. For most missions in which ILS is under contract, the satellite is shipped to Baikonur in aconfiguration that requires minimal nonhazardous checkout. ILS launch campaign process goals reflect the desire tominimize launch campaign duration, without jeopardizing mission success. The 30-day cycle is consistent with thisphilosophy.


Page 1-26

Figure 1.10.2-1: Proton Launch Campaign for 92A-50

L-30 days

Spacecraft L-30 days L-28 days L-27 days L-26 days

Ground SupportEquipment L-34 days L-30 » L-19 days

L-3 months L-2.5 months L-2 months L-18 days

L-11 » L-10 days L-16 » L-14 days

L-9 » L-7 days

L-13 » L-12 days

L-6 days L-5 » L-4 days L-3 days

L-17 days L-18 days

Fuel SC at Fueling Hall withSC Fuel Carts & Return toTechnical Assembly Hall

Move GSE to EmbarkationFacility for Shipment to Russia

SC Loadedinto SC Containerat Factory

Rail Shipment toBaikonur

Move Fuel Chartsto Fuel Hall

Temporary Storage

CheckoutSpacecraftPyrotechnics

GSE Loaded in CargoContainers & onPallets at Factory

Move SC Container to Airportvia 5-Wheel Tow; Load Container into Transport Aircraft

SC

Propellant Shipment toRussia Port-of-Entry

ILS & SC Technical SupportTeam Onsite

• Transfer Payload to Launch Vehicle 4th Stage Refueling Area• Mate Spacecraft Vertically To Launch Vehicle 4th Stage• Tilt to Horizontal & Mate Fairing• Test Transit Cable• Load Spacehead Unit on Transporter

• Prepare & Mate Payload Section with Launch Vehicle• Checkout Electrical Connections• Transfer Launch Vehicle with Payload Section onto Erector Transporter

• Proton Rollout from Horizontal Building• Install on Launch Pad• Do General SC Checkout on Launch Pad• General Testing & Preparation of SC & Launch Vehicle• Air Conditioning & Batteries Charge Hook-Up• Electrical Tests of SC

• Final Operations on Spacecraft• Fuel Launch Vehicle Fourth Stage

Connect ElectricalGSE to SC

• Test TC&C Systems• Test Solar Arrays• Pneumatic & Pressure Tests• Test Deployment Mechanisms

SC & GSEArrives at AirportNear BaikonurLaunch Site

Transport SC to TechnicalAssembly Hall

Install SCin Cleanroom

L-2 days

L-1 » L-0 days

Integrated SC& LaunchVehicle Test

FlightReadinessReview

Countdown& Launch

Prepare SCfor Fueling & Transfer toFueling Hall

Transport PayloadSpaceheadTo Launch VehicleAssembly Area


Page 1-27

1.11 PLANNED ENHANCEMENTS

Beyond the Proton M/Breeze M, our Khrunichev partners have planned a number of possible upgrades to the Protonlaunch system. Payload fairings that are longer in length and/or wider in diameter have been assessed. A tandemlaunch system concept has been studied.

Additionally, a new liquid oxygen/liquid hydrogen high energy upper stage, to be used in place of the currentLOX/hydrocarbon Block DM or storable propellant Breeze M for some missions, is also under development.Maximum performance capability of the improved Proton is expected to be as follows:

LEO (200 km circ.) 22.0 metric tons

GSO (LOX/LH2) 4.5 metric tons

The above enhancements are currently projected to begin appearing in flight vehicles starting in the year 2001 and on.Additional detail can be found in Section 8 of this Guide.

1.12 PROTON PRODUCTION AND OPERATIONS-SLIDES

Figures S-1 to S-10 show manufacturing operations for first three stages of Proton at the Khrunichev plant outside ofMoscow. Figures S-11 to S-13 show the Block DM assembly area at Energia. Figures S-14 and S-15 show externaland internal views of the Proton Standard Commercial Payload Fairing. Figures S-16 to S-19 show the Breeze Mbeing manufactured at Khrunichev. Figure S-20 shows the arrival of a commercial spacecraft, in its shippingcontainer, at the Yubileni airport at Baikonur. Figures S-21 and S-22 show spacecraft processing operations in Area 92and Area 31, respectively. Figure S-23 shows the complete Proton launch vehicle, after integration of the Ascent Unitonto the first three stages, being lifted onto its transporter/erector. Figure S-24 illustrates the launch vehicle interfacesat the base of a Proton launch pad. Figure S-25 shows the Proton being erected onto the pad. Figure S-26 shows theProton’s Mobile Service Tower. Figure S-27 shows the Proton at the moment of lift-off.


S-1

PROTON

PRODUCTION AND OPERATIONS

SLIDES


S-2

Figure S-1: Tank Component Fabrication At Khrunichev Uses Automated Machining Centers

Figure S-2: First Stage Propellant Tank Automated Weld-Up


S-3

Figure S-3: First, Second, And Third Stage Sub-Assemblies Awaiting Integration

Figure S-4: Proton Interstage Component Fabrication


S-4

Figure S-5: First Stage Build Up On Proton Fixture

Figure S-6: Interstage Joining Second and Third Stages


S-5

Figure S-7: RD-253 High-Pressure Engine On The First Stage External Fuel Tank

Figure S-8: Proton Final Assembly Hall At Khrunichev


S-6

Figure S-9: Assembled First Stage Showing Hold Down Points and Aft End Services Connectors

Figure S-10: End-To End Test Of Assembled Proton At Khrunichev Factory


S-7

Figure S-11: The Block DM Undergoing Final Assembly And Testing At The Energia Plant In Korolev, Near Moscow

Figure S-12: The Completed Block DM Stage, Before Attachment Of The External Shroud


S-8

Figure S-13: Finished Block DM Stages In Their Aerodynamic Shrouds, Awaiting Shipment To Baikonur

Figure S-14: Proton Standard Commercial Payload Fairing External View


S-9

Figure S-15: Proton Standard Commercial Payload Fairing Internal View

Figure S-16: Breeze M Stage Structural Components In Manufacture


S-10

Figure S-17: Breeze M Core Stage Manufacturing

Figure S-18: Breeze M Avionics Bay Assembly


S-11

Figure S-19: Breeze M Stage Final Assembly

Figure S-20: Spacecraft Arrival at Baikonur


S-12

Figure S-21: Spacecraft Processing In Area 92 (SVPF)


S-13

Figure S-22: Spacecraft Processing and Encapsulation In Area 31

Figure S-23: The Complete Proton Launch Vehicle Being Lifted Onto Its Transporter/Erector


S-14

Figure S-24: Launch Vehicle Interfaces at Proton Launch Pad, Baikonur

Figure S-25: Proton Erection At The Launch Pad


S-15

Figure S-26: Mobile Service Tower (MST)


S-16

Figure S-27: Lift-Off Of The Proton K/Block DM Carrying A Commercial Communications Satellite


Page 2-1

2. VEHICLE PERFORMANCE

2.1 OVERVIEW

This section provides the information needed to make preliminary performance estimates for several variants of theProton family, into a variety of mission orbits. It is organized so as to provide the user with essential backgroundmission planning information; detailed performance tables and charts follow the text material.

Trajectory profile and operational mission characteristics of the Proton K and Proton M launch systems using theBlock DM and Breeze M upper stages are provided in the first sections of the chapter. Mission performance data,guidance accuracy data, and upper stage attitude control capabilities are found in the last half of this section.

2.2 PROTON LAUNCH SYSTEM CAPABILITIES

The Proton has been operational since 1970 (Pre-1970 launches were considered developmental), and has carried outmore than 246 launches as of December 1998. Over the last 50 launches, the Proton has achieved a success rate inexcess of 92%. The Proton launch vehicle is presently available in three variants: the three-stage Proton booster the Kand M variants offer similar performance in this configuration and the four-stage Proton K/Block DM and ProtonM/Breeze M variants. Proton K/Breeze M and Proton M/Block DM variants may also be implemented, should thisbe desireable.

The production Proton K 4-stage vehicle is equipped with a large upper stage known as the Block DM. The Block DMcan place spacecraft into a variety of orbits including low, intermediate, and high Earth circular, geotransfer,geosynchronous, sun synchronous, and inter-planetary trajectories.

The Proton M is presently completing development. Although identical in outward appearance to the existing ProtonK, it incorporates improvements to the avionics and structures of the first three stages. It also incorporates improvedengines that have been flying since 1996. The Proton M is capable of delivering 21.0 metric tons into a 200 kmcircular, 51.6° inclination orbit. The Proton M/Breeze M will become available for commercial use in 2000. TheBreeze M storable propellant upper stage offers enhanced performance and operational capability and is described inSection 1. Table 2.2-1 summarizes performance for Proton K, Proton M, Proton K/Block DM, and ProtonM/Breeze M to a range of mission orbits.

Table 2.2-1: Summary Proton Performance (PSM) to Reference Orbits

Mission Proton K (Standard) Proton K (Enhanced) Proton M

LEO (51.6 degrees) 19,760 kg (43,560 lb) 19,760 kg (43,560 lb) 21,000 kg (46,300 lb)

GSO 1,880 kg (4,145 lb) 1,880 kg (4,145 lb) 2,920 kg (6,435 lb)

GTO (1800 m/s from GSO) 4,350 kg (9,590 lb) N/A 6,220 kg (13,710 lb)

GTO (1500 m/s from GSO) 4,350 kg (9,590 lb) 4,910 kg (10,825 lb) 5,500 kg (12,125 lb)

Lunar 4,530 kg (9,987 lb) 4,530 kg (9,987 lb) 5,600 kg (12,345 lb)

Mars Transfer 2,940 kg (6,482 lb) 2,940 kg (6,482 lb) 4,800 kg (10,580 lb)


Page 2-2

2.2.1 The Baikonur Launch Site

The Proton launch complex, consisting of spacecraft and launch vehicle processing and integration facilities and fourlaunch pads (two of which are available for commercial use), is located at the Baikonur Cosmodrome. Baikonur,shown in Figure 2.2.1-1, is located approximately 2,000 km (1,300 miles) southeast of Moscow in the Republic ofKazakhstan. The Baikonur Cosmodrome measures approximately 90 km east-to-west, and 75 km north-to-south, andsupports many other launch vehicles, including the Soyuz, Vostok, Molniya, Zenit and Energia. Temperatures rangefrom -40°C to 45°C during the year.

Figure 2.2.1-1: The Baikonur Launch Site, Showing Available Direct Injection Inclinations

80o

60o

40o

20o

0o180o160o140o120o100o80o60o40o20o0o20o

CA SP IAN SE A

B L A C K S E A

M E D I T E R R A N E A N SEA

RED

SEA

MOSCOW

FLIGHT AZIMUTH: 22.5o

INCLINATION: 72.7o


INCLINATION: 64.8o


INCLINATION: 51.6o

BAIKONURCOSMODROME

(45.6oN)(TYURATAM)

2.2.2 Launch Availability

The Proton launch system is designed to operate under the severe environmental conditions encountered at Baikonur(Table 2.2.2-1). The Proton can be launched year around, and the time between launches from an individual pad canbe as short as 25 days. Proton has demonstrated a launch rate of four per month from multiple launch pads, and a longterm average launch rate of approximately twelve per year. The capability of the Proton system to launch in severeenvironmental conditions decreases launch delays and ensures that payloads reach orbit as scheduled to begin revenuegenerating activities. The short turnaround time between launches can ensure that spacecraft constellations aredeployed quickly, minimizing the time required to enter service.


Page 2-3

Table 2.2.2-1: Allowable Launch Environment Constraints

Temperature -40OC to 45OC

Maximum Launch Ground Winds (Commercial PLF) 16.5 m/s

Times Launches available year round

Turn Around Times 25 days per pad

Number of Pads 4 (2 Commercial)

2.2.3 Payload Fairings and Adapters

2.2.3.1 Payload Fairings

Multiple payload fairings types have flown on the Proton; these fairings are discussed in detail in Sections 1 and 4 ofthis document. ILS currently offers one standard commercial payload fairing for the four stage Proton K/Block DM.Two additional fairings are available for the Proton M/Breeze M.

Payload fairing jettison times are constrained to occur so that fairing hardware will impact in designated areas. ForProton K/Block DM and initial Proton M/Breeze M flights, fairing jettison occurs at approximately 342 to 344seconds (121-125 kilometers altitude) into the flight. Free molecular heating is below 1,135 W/m2 for these cases.

2.2.3.2 Payload Adapters

Available Proton payload adapters are shown in Appendix D. These adapters provide mechanical and electricalinterfaces compatible with established commercial launch industry standards and can accommodate a wide range ofpayload requirements. ILS can assist in the development of specialized adapters as required.

2.2.4 Upper Stage Capabilities

Two upper stages are currently available for commercial Proton missions, the Block DM and the Breeze M.

The Block DM has many unique capabilities that can accommodate unique mission requirements. The heavy liftcapability of the Proton’s first three stages, coupled with the Block DM’s high performance, allows the placement ofspacecraft directly into their final orbits, reducing the size and complexity of spacecraft propulsion systems. The BlockDM’s multiple start capability, designed for up to 7 starts in flight (with 5 demonstrated to date), and a minimum 24-hr orbital lifetime, can increase mission utility through unique mission design. Block DM orbital lifetime can beextended through battery upgrades and other hardware modifications.

The Breeze M upper stage expands upon the capabilities enabled by the Block DM. The main engine of theBreeze M can be restarted up to 8 times in flight and allows the stage to offer high precision placement of spacecraftinto orbit. With storable propellant, Breeze M orbital lifetime is limited only by available on board battery power andis currently in excess of 24 hours. The jettisonable, toroidal propellant tanks offer significant mission design flexibilityand enable launch services to be offered for low and high energy delivery requirements.


Page 2-4

2.3 PROTON ASCENT PROFILE

2.3.1 Proton Booster Ascent

The first three stages of the Proton launch vehicle use a standard ascent trajectory to place the orbital unit (upper stageand/or payload) into a 200 km (108 nmi) circular parking orbit inclined at either 51.6°, 64.8°, or 72.7°. A standardascent trajectory is required to meet jettisoned stage and payload fairing impact point constraints. The use of a standardascent trajectory also simplifies lower ascent mission design and related analysis, thereby increasing system reliability.Once a payload is in the standard parking orbit at one of the three available inclinations, it can be transferred to itsfinal orbit by either the Block DM or Breeze M.

Figure 2.3.1-1 pictorially illustrates a typical Proton ascent into the standard parking orbit. Table 2.3.1-1 lists the timeof occurrence for major ascent events for a typical launch. The six Stage 1 RD-253 engines are ignited atapproximately T-1.6 sec. and are commanded to 40% of nominal thrust. Thrust is increased to 100% at T-0 sec.Liftoff confirmation is signaled at T+0.5 sec. The staged ignition sequence allows verification that all engines arefunctioning nominally before being committed to launch. The launch vehicle executes a roll maneuver beginning atT+10 sec. to align the flight azimuth to the desired direction. The vehicle incurs its maximum dynamic pressure of3,890 N/m2 (800 psf) approximately 65 sec. into flight. Stage 2’s four RD-0210 engines begin their ignition sequenceat 122 sec. and are commanded to full thrust when Stage 1 is jettisoned at 126 sec. Payload fairing jettison typicallyoccurs at 344.2 sec into flight, depending on spacecraft heating constraints. Stage 3’s vernier engines are ignited at 332sec. followed by Stage 2 shutdown at 334 sec. Stage 2 separation occurs after six small, solid retro-fire motors areignited at 335 sec. into flight. Stage 3’s single RD-0210 engine is ignited at 339 sec. and burns until shutdown at 567.1sec. The four vernier engines burn for an additional 10 sec. and are shutdown at 585 sec. After a 10 second coast, theStage 3 retro-fire motors are ignited and Stage 3 is separated from the upper stage or spacecraft. Figure 2.3.1-2 showsthe ascent vacuum instantaneous impact points and ground track. Figure 2.3.1-3 shows the ascent telemetry coverageprovided by the CIS ground stations. Figure 2.3.1-4 shows the times and values for the vehicle’s inertial velocity,altitude, and dynamic pressure.


Page 2-5

Table 2.3.1-1: Typical Booster Ascent Event Times

Event Description Event Time (sec)

Stage 1 ignition - 10% thrust -1.60

Begin stage 1 thrust to 100% -0.00

Lift-off 0.57

Stage 1 thrust to 100% 1.00

Stage 2 ignition 116.91

Stage 1 / 2 separation 121.11

Stage 3 vernier engine ignition 330.00

Stage 2 engine shutdown 332.70

Stage 2 / 3 separation 333.40

Stage 3 main engine ignition 335.80

PLF jettison 344.20

Stage 3 main engine shutdown 567.11

Stage 3 vernier engine shutdown 577.11

Stage 3 upper stage separation 582.01


Page 2-6

Figure 2.3.1-1: Typical Proton Booster Ascent

qmax

T=65V=125.1H=10.5nx=1.97q=35265

Separationof 1st stageT=126.73V=1649.4H=43.5nx=3.6/0.4q=3350

Separationof 2nd stageT=335.12V=4331.9H=148.11nx=2.8/0.04q=0

Separation of payload fairingT=344.2V=4340.5H=157.7nx=0.9q=0

Separationof 3rd stageT=589.0V=7477.4H=228.0nx=0.13/0q=0

Ignition No. 1of 4th stagenx=0.44

Burnoutnx=0.9

Ignition No. 2of 4th stagenx=0.9

Burnout andseparationnx=1.7/0

T = time (seconds)V = velocity (meters per second)H = altitude (kilometers)L = horizontal distance (kilometers)q = dynamic pressure (Pa)nx = axial acceleration at end of prior

stage/beginning of next stage

1750 x 6 = 10500 kN

583 x 4 = 2332 kN

583 + 31 = 614 kN

Retro

83.5 kN 83.5 kN

Retro

L = 1985L = 310

Figure 2.3.1-2: Typical Proton Booster Ascent Ground Track and Vacuum Impact Points

-180-165

-150-135

-120-105

-90 -75 -60 -45 -30 -15 0 15 30 45 60 75 90 105120

135150

165180

-90

-75

-60

-45

-30

-15

0

15

30

45

60

75

90

Moscow

Vacuum Impact Points

Longitude, degrees

Stage 1 Jettison PLF Jett Option 1

Stage 2 Jettison PLF Jett Option 2

Orbit Injection

La

titu

de

, d

eg

ree

s


Page 2-7

Figure 2.3.1-3: Typical Ground Tracker Acquisition Times for Proton Ascent to the Support Orbit

D s h u s a l i ( D Z h S )

K o l p a s he v o ( K L P )

U s s ur ick (U S K )

0 1 0 0 2 00 3 00 4 0 0 5 0 0 6 0 0

T i m e f r o m L i f t o f f ( s e c o n d s)

Figure 2.3.1-4: Typical Proton Lower Ascent Altitude, Inertial Velocity, Acceleration, and Dynamic Pressure

Alti

tud

e (k

m)

Dyn

ami

c pr

essu

re

(N/m

2 )

Acc

eler

atio

n (

g)V

elo

city

(m

/sec

)


Page 2-8

2.3.2 Block DM Trajectory and Sequence

The Block DM transfers payloads to a variety of mission orbits. This section describes the two options enabling theBlock DM to transfer a payload to a geosynchronous transfer orbit.

Figures 2.3.2-1 and 2.3.2-2 illustrate a typical Block DM ascent into a geosynchronous transfer orbit. The Block DMis delivered to the 200 km circular parking orbit with a 51.6-deg inclination. Once on orbit and separated from theProton third stage, the Block DM executes a 15-min duration maneuver to properly align its longitudinal axis for thefirst burn. After the alignment maneuver, the Block DM enters into a stabilized flight mode. Twenty-five min after thelongitudinal alignment maneuver, the Block DM executes a 180-deg turn about the roll axis to compensate forpossible gyroscopic drift. This maneuver can also help with spacecraft thermal management. Forty min after the rollmaneuver, the Block DM reaches the first ascending node, and the two SOZ unit’s four 2.5-kg axial loading enginesbegin a 300-sec burn to settle the stage’s propellants. After this settling burn, the main engine ignites, raising thetransfer orbit apogee to geosynchronous altitude. The first main engine burn lasts approximately 450 sec. After enginecutoff, the Block DM executes another alignment maneuver to place the longitudinal axis in the correct orientation forthe second burn. The Block DM then enters stabilized flight for the approximately 5 hr and 15 min transfer requiredto reach transfer orbit apogee. At approximately 2.5 hr into this transfer ellipse, another 180-deg rotation about theroll axis is executed. After reaching the transfer orbit apogee, the Block DM initiates another 300-sec propellantsettling burn followed by a main engine burn of approximately 230 sec. duration to circularize the orbit and reduce theinclination to 0 deg. The Block DM then commands spacecraft separation 14.8 ±0.05 sec after the end of the finalburn. Figure 2.3.2-3 depicts the upper ascent ground track and Figure 2.3.2-4 shows available tracker coverageprovided by the CIS ground stations. Telemetry (including state vector) data can also be collected from the Block DMduring transfer orbit flight by Russian geostationary spacecraft. Table 2.3.2-1 gives the times and values of the BlockDM attitude maneuvers. These maneuvers may be modified to assist in spacecraft thermal management; the BlockDM can perform a maximum of 11 such maneuvers between the 1st and 2nd burn during a normal mission. with amission maximum of 14. The standard mission scenario for commercial spacecraft involves retention of the SOZ unitsto provide attitude control following main engine shutdown. This allows spacecraft separation to be delayed until afterthe Block DM completes reorientation to a customer specified separation attitude.

To more optimally deliver western satellites to orbit, the Block DM is also capable of performing a suborbital burn inorder to enable larger masses to be delivered to low earth orbit. The larger mass is made up of the heavier payload andmore fuel on the Block DM stage for orbital maneuvers. The “enhanced” Block DM performance level is enabledwith the suborbital burn.

Normally, injection into final orbit occurs at 90O however, longitude placement for geosynchronous orbits can becontrolled with an increment of 12.25 deg by allowing the Block DM and payload to coast in the standard parkingorbit for the required number of revolutions, and initiating the first transfer orbit burn on either the ascending ordescending node. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parkingorbit revolutions are within the lifetime constraints of the Block DM, giving complete 360-deg coverage.


Page 2-9

Figure 2.3.2-1: Proton/Block DM Upper Ascent

Event description(block DM upper ascent)

Event time(sec)

Block 4 upper adapter JettisonSOZ unit first setting burn (SOZ1)Block 4 DM first burnSOZ unit engine shutdownSOZ unit second burn (SOZ2)Block 4 DM second burnSOZ unit engine shutdownBlock 4 DM separation

63754375732573724,80025,10525,10525, 345

Begin attitudemaneuver forblock DM 1st burn

180° rotationabout roll axis

Begin propellantsettling burn(MES1 - 300sec)

Block DM1st burn(MES1)

Begin altitudemaneuver forblock DM2nd burn

MECO2

Block DM/payloadseparationMECO2+14.8sec

180° rotationabout rollaxis

Begin propellantsettling burn(MES1 - 300sec)

Block DM2nd burn(MES 2)

MECO1


Page 2-10

Figure 2.3.2-2: Block DM Two-Impulse Transfer to Geosynchronous Transfer Orbit

it= variableParking Orbit

DV2

iref= 51.6o

4

1

Target OrbitHa= 36,000 km

Hp= variable km

SC Separation 6

DV3

7

GSO

IntermediateTransfer Orbit

1 - Launch from Baikonur Cosmodrome and booster ascent

2 - (DV1) For sub-orbital burn cases, f irst Block DM main engine fairing

3 - Coast to ascending node of parking orbit

4 - (DV2) Block DM main engine firing to raise orbit from LEO to GTO

5 - Coast to apogee of transfer orbit (5.25 hours)

6 - (DV3) Block DM main engine f iring to optimally raise perigee and reduce inclination

7 - Spacecraft separation from Block DM

3

2

DV1

5


Page 2-11

Figure 2.3.2-3: Proton K/Block DM Upper Ascent Ground Track to GSO

B o o s t e r

K O

B u r n 3

A Y 1

B u r n 2

Figure 2.3.2-4: Ground Tracker Acquisition Times for Proton Ascent to a GSO

Shelkovo (SHLK)

Petropavlovsk-Kamchatski(PPK)

Ussurick

Ulan-Uden (ULD)

Kolpaashevo (KLP)

Dshusali (DZhS)600 5,60

010,600 15,60

020,600 25,600

Time from liftoff(seconds)


Page 2-12

Table 2.3.2-1: Typical Block DM Attitude Maneuvers for Geosynchronous Mission (90O East Longitude)

Start Time of AttitudeManeuver (hr.min.sec)

Change inPitch Angle

(deg)

Change inYaw Angle

(deg)

Change inRoll Angle

(deg)

Duration ofManeuver

(sec)

Comments

00.10.44 43.00 - - 86

00.10.59.9 33.02 7.54 -13.25 269 Attitude alignment for FirstBlock DM burn

00.17.51 - - 180.00 900 IMU Compensation

00.35.45 - - -180.00 -900 IMU Compensation

01.01.19 12.07 - - 60 First Block DM burn

-01.13.28 -20.90 - - 420 First Block DM burn

01.21.36 9.00 - - 48 First Block DM burn

02.26.45 - 180.00 - 900 IMU Compensation

03.38.37 180.00 - - 900 IMU Compensation

04.50.10 - 180.00 - 900 IMU Compensation

06.18.33.5 7.78 68.37 2.6 273 Second Block DM burn

2.3.3 Breeze M Trajectory Sequence

The Breeze/payload unit is placed into a high-energy suborbital state by the third stage of Proton. After jettison of thethird stage, the Breeze upper stage performs a small propulsive maneuver to deliver itself and the attached satellite to astandard low earth parking orbit. After a coast of approximately 45 minutes, the Breeze stage performs the second offour propulsive maneuvers. This second main burn is used to begin the process of raising the apogee of the transfer orbitto geosynchronous altitude.

Due to the thrust of the Breeze M stage (19.6 kN), the optimum mission profile results by splitting the "apogee raising"propulsive maneuver into two smaller burns. The first burn raises apogee altitude to approximately 5,000 km - 7,000km; the actual value is mission and satellite mass dependent. After one full revolution in this intermediate transferorbit (approximately 2.5 hours), the third main burn of the Breeze M main engine occurs raising apogee altitude togeosynchronous. Perigee altitude and inclination are adjusted somewhat based on the trajectory optimization processthat occurs during the mission integration process.

During all nonthrusting periods, the fourth stage of Proton is able to align the satellite to an attitude that is compatiblewith the thermal and solar incidence requirements for the satellite. Roll maneuvers can be programmed to ensure evenheating/cooling of the satellite surfaces.

After a coast of approximately 5.2 hours, the Breeze M provides its fourth and typically final propulsive maneuver,raising perigee altitude and lowering orbit inclination to the optimum extent. After completion of the Breeze M fourthburn, the satellite is oriented and separated from the upper stage. Total elapsed time from launch for a typical BreezeM mission is approxiamately 10 hours.


Page 2-13

Normally, injection into final orbit occurs at 90O however, longitude placement for geosynchronous orbits can becontrolled with an increment of 12.25 deg by allowing the Breeze M and payload to coast in the standard parking orbitfor the required number of revolutions, and initiating the first transfer orbit burn on either the ascending or descendingnode. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parking orbitrevolutions are within the lifetime constraints of the Breeze M, giving complete 360-deg coverage.

Figure 2.3.3-1 illustrates the main characteristics of the trajectory for a Proton M/Breeze M launch to geosychronoustransfer orbit.

Figure 2.3.3-1: Typical Breeze M Flight Profile to Geosynchronous Transfer Orbit

2.3.4 Collision and Contamination Avoidance Maneuver

The Block DM and Breeze M can perform a variety of maneuvers to minimize the possibility of recontact with orcontamination of a customer’s spacecraft. The separation event provides a typical relative velocity between thespacecraft and the upper stage of 0.3 meter/sec.


Page 2-14

2.3.4.1 Block DM Upper Stage

Approximately 2 ½ hours after spacecraft separation, a 300-sec burn of the SOZ units is performed to increase therelative separation velocity between the upper stage and the spacecraft. This is followed by a burn of opposing pairs ofSOZ thrusters to SOZ propellant depletion (with zero net delta-velocity). Finally the LOZ tank residuals are ventedthrough a pair of forward facing nozzles located on either side of the stage near the bottom of the LOZ tank. Thesenozzles are canted at 15 degrees from the vertical in such a way as to spin up the stage whole simultaneously adding arelative velocity increment between spacecraft and upper stage of up to 5 m/sec.

2.3.4.2 Breeze M Upper Stage

Approximately 1 ½ hours after spacecraft separation, the Breeze M stage performs an attitude change maneuver to re-orient the stage. A small propulsive maneuver is made to increase relative velocity between the upper stage andspacecraft. After completion of the maneuver, the upper stage propellant tanks are depressurized and the stage is madeinert. Relative velocity increments between the spacecraft and upper stage are similar to those for Block DM areachieved.

2.4 PERFORMANCE GROUNDRULES

A number of standard mission groundrules have been used to develop the reference Proton K and Proton Mperformance capabilities identified in this document. They are identified in this section.

2.4.1 Payload Systems Mass Definition

Performance capabilities quoted throughout this document are presented in terms of payload systems mass. PayloadSystems Mass (PSM) is defined as the total mass delivered to the target orbit, including the separated spacecraft, thespacecraft-to-launch vehicle adapter, and all other hardware required on the launch vehicle to support the payload(e.g., harnessing). Table 2.4.1-1 provides masses for the standard Proton adapter systems. Data is also provided forestimating the performance effects of various mission-peculiar hardware requirements. As a note, the performanceeffects shown are approximate.

Table 2.4.1-1: Launch Vehicle Mission Peculiar Hardware

Block DM Breeze M

∅1666 mmtwo piece adapter

175 kg (386 lb) ∅1666 mm adapter 175 kg (386 lb)

∅1194 mm adapter500 mm height

120 kg (265 lb) ∅1194 mm adapter 150 kg (331 lb)

∅1194 mm adapter625 mm height

130 kg (287 lb) TBD adapter 200 kg (441 lb)


Page 2-15

2.4.2 Payload Fairings

All performance quotes in this document are based on use of standard payload fairing options. For Proton K/BlockDM the standard commercial payload fairing is used. For Proton M/Breeze M the baseline “short” payload fairingoption that enables a usable volume similar to Block DM is used. For Proton M/Breeze M, a performancedegradation of approximately 100 kg results for high-energy geosynchronous transfer missions when the “long”payload fairing option is used.

In all cases, the payload fairing is jettisoned at a location that ensures a free molecular heating rate after jettison of<1135 W/m2.

2.4.3 Mission Analysis Groundrules

All Proton mission estimates provided in this document assume launch from the Baikonur Cosmodrome. LaunchComplex 23/24 at Baiknour is located at the 46.1 degrees north geodetic latitude and 63.0 east longitude. All identifiedaltitudes are based on an Earth radius of 6378.16 km (3443.93 nmi).

Caution must be exercised in deriving performance estimates for missions whose inclinations differ from thosepresented. The first three stages of the Proton launch system can only deliver payloads directly into, or near, thestandard low earth parking orbit at an inclination of 51.6 degrees, 64.8 degrees, or 72.7 degrees. All other inclinationscan be reached only through an orbital plane change maneuver. Performance estimates should not be made based oninterpolation between performance values derived from different parking orbit inclinations.

2.4.4 Performance Confidence Levels

Proton missions are targeted to meet the requirements of each user. Historically, Proton missions have been targetedbased on a conservative 3-sigma confidence level (or greater) that the mission objectives would be achieved. All Protonperformance information contained in this document assumes a 3-sigma confidence level.

2.5 DIRECT INJECTION LEO MISSIONS

The three stages of the Proton vehicle typically inject their payload into the standard approximately 200 km (108 nmi)circular parking orbit at inclinations of 51.6°, 64.8°, or 72.7°. Table 2.5-1 shows the performance into these threeavailable orbits.

The Proton launch vehicle can only be launched along the flight azimuths that yield orbit inclinations of 51.6°, 64.8°,and 72.7°, as shown in Figure 2.2.1-1. The flight azimuths are constrained by available stage and fairing impact pointsalong the flight trajectory. If a mission requires an inclination different from 51.6°, 64.8°, and 72.7°, the inclinationchange is performed by the Block DM or Breeze M, or the spacecraft after achieving Earth orbit.

Table 2.5-1: Proton Booster Performance to Low Earth Orbits (Direct injection, no upper stage)

Orbit Inclination Proton K Proton M

186 x 222 km, 51.6 deg 19,760 kg 43,560 lb 21,000 kg 46,300 lb

175 km circular, 64.8 deg 19,300 kg 42,550 lb 20,610 kg 45,435 lb

170 km circular, 72.7 deg 18,900 kg 41,668 lb 19,975 kg 44,035 lb


Page 2-16

2.6 GEOSYNCHRONOUS TRANSFER MISSIONS

2.6.1 Launch to Geosynchronous Transfer Orbit

The high-energy geosynchronous transfer mission is the standard mission profile for most commercial Protonlaunches. Variable mass satellites are delivered to a geosynchronous transfer orbit with a 35,786 km apogee altitude, a 0degree argument of perigee, and a variable orbit perigee and inclination consistent with the liftoff mass of the satelliteand the delivered launch vehicle performance. From this point, the satellite will perform the remaining perigee raisingand inclination reduction to reach geosynchronous orbit.

For reference purposes, ILS has established a reference goesynchronous transfer mission performance quotation basedon an orbit that is 1,500 m/sec delta-velocity from geosynchronous orbit. This reference mission is indicative of thegeosynchronous transfer missions used by vehicles launched from low inclination launch sites. The reference orbitassumes a 5,500 km perigee altitude, a 35,786 km apogee altitude, a 25.0 degree orbit inclination, and a 0.0 degreeargument of perigee.

Proton K/Block DM Performance - Performance to elliptical transfer orbits with GSO apogees is shown in Figure2.6.1-1 for the Proton K/Block DM. Data is shown that represents launch vehicle performance as a function ofpayload systems and residual delta-velocity from targeted transfer orbit to geosynchronous orbit. Analyses have beenconducted to determine the near optimum orbit that can be achieved with Proton given a spacecraft mass and thevarious performance variants of the Proton launch system. Given a payload mass and launch from the BaikonurCosmodrome, the upper stage of Proton delivers the satellite to a high-energy GTO that results in the minimum delta-velocity remaining to reach GSO. The derived perigee altitudes and orbit inclinations are provided in Table 2.6.1-1 forthe standard Block DM 2-Burn mission profile. Table 2.6.1-3 provides parametric GTO performance data for a fullrange of perigee altitude and orbit inclination combinations.

For specific payload mass ranges, the Proton booster ascent profile can be tailored to enable a suborbital burn of theBlock DM upper stage enabling a higher upper stage/payload mass to be delivered to low earth orbit. The 3-Burnmission is executed by the Proton K booster delivering a 23,000 kg (50,700 lb) orbit unit “stack” to a high-energysuborbital state. The Block DM, after separation from the third stage of Proton, performs a main engine firing toachieve a low earth parking orbit. The Block DM mission then progresses similarly to the two burn profile. In thisinstance, 3 firings of the Block DM main engine result in a higher performance capability results as the Block DMreaches LEO with a higher propellant load than if it were offloaded and delivered directly to LEO by the boosterstatges. Table 2.6.1-2 identifies the optimum geosynchronous transfer orbit versus payload relationship for the“Enhanced” 3-Burn mission capability. The 3-burn profile is enabled with payloads that are greater than or equal to4,600 kg (10,141 lb). Additionally, the Block DM structural load limitations, as specified elsewhere in this document,must be adhered to in order for this capability to be enabled.

Proton M/Breeze M Performance – For Proton M/Breeze M, three performance configurations are identified. Figure2.6.1-2 plots optimum GTO payload versus Delta-Velocity to GEO for these three performance variants of ProtonM/Breeze M. The first configuration (Configuration 1) represents the performance capability for Proton K/Breeze Mfor initial flights which are to commence in early 1999. In this configuration, the Proton K booster delivers the BreezeM to orbit. A payload systems mass of 4800 kg can be delivered to a reference GTO in this case. With introduction ofthe Proton M booster, vehicle performance to GTO increases to 5200 kg PSM. Flights 5, 6 and 7 of the Breeze M, tobe launched during the second half of 2000, will deliver this performance capability to GTO and are represented by theConfiguration 2 data. Table 2.6.1-4 provides tabular data corresponding to the data identified in the performance plot.Additionally, Table 2.6.1-5 provides parametric performance data for a full range of geosynchronous transfer missionsfor the mature flight configuration, Configuration 3, of Proton M/Breeze M.


Page 2-17

Figure 2.6.1-1: Proton K/Block DM Performance To Representative Geosynchronous Transfer Orbits

2500

3000

3500

4000

4500

5000

5500

500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900

Delta-V to GSO (m/s)

Pay

load

Sys

tem

Mas

s (k

g)

Enhanced

Standard


Page 2-18

Table 2.6.1-1: Proton K/Block DM Performance To Representative Geosynchronous Transfer Orbits

Delta-V to GSO Orbit Inclination Perigee Altitude Payload Systems Mass (PSM)

(m/s) (deg) Hp (km) (2-Burn)

600 7.0 16600 2732

700 8.2 14400 2887

800 9.7 12600 3046

900 11.3 11000 3210

1000 13.1 9600 3382

1100 15.0 8300 3560

1200 17.0 7200 3740

1300 19.0 6100 3925

1400 21.1 5100 4143

1500 23.3 4200 4350

1600 25.7 3400 4350

1800 31.0 2100 4350

Transfer Orbit Parameters as Specified with Apogee Altitude of 35,786 km;

Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m2;Launch Mission Duration Approximately 6.5 hours.

Table 2.6.1-2: Proton K/Block DM Three-Burn Mission Performance to Representative Geosynchronous TransferOrbits

Delta-V to GSO Orbit Inclination Perigee Altitude Payload Systems Mass (PSM) (m/s) (deg) Hp (km)1343.6 19.83 5600 46001375.0 20.52 5310 46661400.0 21.08 5082 47191425.0 21.64 4859 47721450.0 22.21 4640 48251475.0 22.79 4425 48791500.0 23.37 4217 49321525.0 23.96 4016 4985

1500.0 25.00 5500 4910

Transfer Orbit Parameters with Apogee Altitude of 35,786 km;Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m2;Launch Mission Duration Approximately 6.5 hours.


Page 2-19

Figure 2.6.1-2: Proton M/Breeze M Performance To Representative Geosynchronous Transfer Orbits

3000

3500

4000

4500

5000

5500

6000

6500

500 600 700 800 900 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900

Delta-V to GSO (m/s)

Payl

oad

Syst

ems M

ass (

kg)

Config 1

Config 2

Config 3


Page 2-20

Table 2.6.1-3: Proton K/Block DM Parametric Geosynchronous Transfer Performance Date (2 Burn Mission)

Inclination, deg. 0.0 5.0 10.0 15.0 20.0 25.0

Perigee altitude, km

200 3450 3680 3920 4170 4350 4350

500 3430 3655 3895 4150 4350 4350

1000 3390 3615 3860 4115 4350 4350

2000 3310 3540 3780 4040 4310 4350

4000 3150 3380 3625 3880 4150 4350

6000 3010 3240 3475 3725 3985 4250

8000 2880 3100 3335 3580 3830 4085

10000 2760 2980 3210 3445 3690 3930

12000 2650 2870 3090 3320 3555 3790

15000 2505 2715 2935 3150 3380 3600

19000 2340 2545 2750 2965 3175 3385

23000 2200 2400 2600 2805 3000 3200

27000 2080 2275 2470 2670 2860 3050

35786 1880 2060 2245 2430 2610 2780

Transfer Orbit Parameters as specified with Apogee Altitude of 35,786 km;Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m2;Launch Mission Duration Approximately 6.5 hours.


Page 2-21

Table 2.6.1-4: Proton M/Breeze M Performance to Representative Geosynchronous Transfer Orbits

Delta-V to GSO (m/s) Inclination (deg) Perigee Alt (km) Config 1 Config 2 Config 3

600 7.0 16600 3181 3581 3825

700 8.2 14400 3339 3739 3990

800 9.7 12600 3504 3904 4160

900 11.3 11000 3673 4073 4335

1000 13.1 9600 3850 4250 4518

1100 15.0 8300 4036 4436 4710

1200 17.0 7200 4221 4621 4903

1300 19.0 6100 4416 4816 5104

1400 21.1 5100 4616 5016 5311

1500 23.3 4200 4821 5221 5523

1600 25.7 3400 5034 5434 5743

1700 28.3 2700 5276 5676 5993

1800 31.0 2100 5496 5896 6220

1500 25.0 5500 4800 5200 5500

Transfer Orbit Parameters as Specified with Apogee Altitude of 35,786 km;Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m2;Launch Mission Duration Approximately 10.0 hours.

Configuration 1: Performance for Proton K/Breeze M Flights (Flights 1, 2, and 3) and contrained performancefor Proton M/Breeze M (Proton/Block DM backup)

Configuration 2: Performance for Proton M/Breeze M Flights 5, 6, and 7Configuration 3: Performance for Proton M/Breeze M Flights 8 and on

Proton Vehicle Variant Performance (kg)


Page 2-22

Table 2.6.1-5: Proton M/Breeze M Parametric Geosynchronous Transfer Performance Data (Configuration 3)

Inclination, deg. 0.0 5.0 10.0 15.0 20.0 25.0 30.0

Perigee altitude, km

200

500

1000

2000

4000

6000

8000

10000

12000

15000

19000

23000

27000

35786

Transfer Orbit Parameters as specified with Apogee Altitude of 35,786 km;Argument of Perigee of 0.0 deg; Payload Fairing Jettison on 3-Sigma qV < 1135 W/m2;Launch Mission Duration Approximately 10 hours.

TO BE SUPPLIEDAT A LATER TIME


Page 2-23

2.7 ORBIT INJECTION ACCURACY

Table 2.7-1 shows Block DM 3-sigma accuracy predictions for various missions. The accuracy predictions areenveloping values and mission-unique analysis will be performed to verify that payload accuracy requirements aresatisfied. Table 2.7-2 provides similar orbit injection accuracy values for the Breeze M upper stage.

Table 2.7-1: Block DM Upper Stage Orbit Injection Accuracies

Perigee Apogee Inclination Arg of Perigee Longitude ofAscending Node

Period

200 km CircularSupport Orbit

± 6 km ± 15 km ± 0.5O ± 0.25O ± 0.025O ±8 sec

10000 km CircularOrbit

± 45 km ± 30 km ± 0.5O ± 0.25O ± 0.1O ± 550 sec

5500 km X 36000 km@ 25.0 deg GTO

± 400 km ± 150 km ± 0.5O ± 0.25O ± 0.5O ± 100 sec

Eccentricity Longitude Inclination Period

Geostationary 0.009 ± 1 O 0.75 O ± 20 min

Table 2.7-2: Breeze M Upper Stage Orbit Injection Accuracies

Perigee Apogee Inclination Arg of Perigee Longitude ofAscending Node

Period

200 km CircularSupport Orbit

± 6 km ± 15 km ± 0.025O ± 0.3O ± 0.15O ±8 sec

500 km Circular Orbit ± 5 km ± 5 km ± 0.05O ± 0.3O ± 0.15O ± 100 sec

1000 km CircularOrbit

± 10 km ± 10 km ± 0.05O ± 0.3O ± 0.15O ± 100 sec

5500 km x 36000 km@ 25.0 deg GTO

± 400 km ± 150 km ± 0.5O - - ± 550 sec

Eccentricity Longitude Inclination Period

Geostationary 0.009 ± 1 O 0.75 O ± 20 min


Page 2-24

2.8 SPACECRAFT ORIENTATION AND SEPARATION

The Proton is upper stages are capable of aligning the spacecraft and separating in a 3-axis stabilized mode; with alongitudinal spinup, or with a transverse spinup enabled with either asymmetric separation springs or via upper stagemaneuvering. Tables 2.8-1 through 2.8-3 show the separation and pointing accuracies for the Block DM and Breeze Mupper stages, for all three separation options. The selected payload separation mechanism will affect separation rates.The orientation and separation conditions are typical values, and mission unique analysis will be performed to verifythat payload requirements are satisfied. The Proton upper stagges can accommodate Customer requirements for spinrates at separation of up to 9.0 deg/sec. Figure 2.8-1 illustrates separation attitude capabilities.

Table 2.8-1: Upper Stage Orbit Injection Accuracies, Option I

Reference Value

SC spin rate about Zsc axis -6 deg/s ± 1%

SC tip-off rate about Ysc and Zsc axes ± 0.8 deg/ s

Relative separation velocity ≥ 0.35 m/s

SC separation pointing vector error ± 5 deg

Table 2.8-2: Upper Stage Orbit Injection Accuracies, Option II

Reference Value

SC spin rate about Xsc axis -2 deg/s ± 1%

SC tip-off rate about Ysc and Zsc axes ± 0.7 deg/ s



Table 2.8-3: Upper Stage Orbit Injection Accuracies, Option III

Reference Value


SC tip-off rate about any of SC axes ± 1.8 deg/ s



Page 2-25

Figure 2.8-1: SC Separation Attitude

Up p e r S t ag eL o n g i t u d i na l ax i s

Sp a ce c ra ft P o si t i o nV e ct o r T ra ns v e r s e

Sp i n

Su n L i ne

Y

X

L o ng i t u d i n a l S p i n

2.9 LAUNCH VEHICLE TELEMETRY DATA

ILS will provide (in Formats I-V , Tables 2.9-1 through 2.9-5 respectively) Launch Vehicle state vector data followingstage 4 insertion into transfer orbit and Spacecraft separation.

Such data is then submitted to the customer by ILS at Baikonur via fax or voice to the Spacecraft Mission ControlCenter.

ILS has adopted standard formats regarding orbital state vector data that are provided to the launch services customerduring and after the launch mission. These standard formats enable the satellite operator to properly determine orbitalconditions at various times during the mission. The standard data is transmitted to the spacecraft Mission ControlCenter at relevant times

The data formats are:

a) Format I - preliminary within 30 minutes after main engine first cut-off, final within 60 minutes. (Table 2.9-1)

b) Format II - within 120 minutes after main engine first cut-off. (Table 2.9-2)

c) Format III - within 30 minutes after Spacecraft separation. (Table 2.9-3)

d) Format IV - within 120 minutes after Spacecraft separation. (Table 2.9-4)

e) Format V - within 30 minutes after deorbit maneuver. (Table 2.9-5)


Page 2-26

The data in Formats I and III are preliminary and subject to clarification in Formats II and IV.

Table 2.9-1: Format I - Preliminary State Vector Data Provided Following Upper Stage 1st Burn

Item Units

Magnitude of DV for main engine 1st firing m/sec

Upper stage 1st burn cutoff (actual) Date and Time (hr, min, sec (GMT))

Roll angle deg

Yaw angle deg

Pitch angle deg

Roll angular velocity deg/sec

Yaw angular velocity deg/sec

Pitch angular velocity deg/sec

Table 2.9-2: Format II - Transfer Orbit Parameters Following Upper Stage 1st Burn

Item Units

Launch time Date and time(hr, min, sec (GMT))

Upper stage 1st burn cutoff (actual)

Semi-major axis km

Eccentricity ---

Inclination deg

Right ascension of the ascending node deg

Argument of perigee deg

Argument of latitude deg

Perigee altitude km

Apogee altitude km

Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation isthe estimated time of the Block DM 1st burn cutoff.


Page 2-27

Table 2.9-3: Format III - Preliminary Vector Data At Separation Epoch

Item Units

Magnitude of DV for main engine 2nd firing m/sec

Upper stage 2nd burn cutoff (actual) Date and Time (GMT)

Actual time of Spacecraft separation from 4th stage Date and Time (GMT)

Roll angle deg

Yaw angle deg

Pitch angle deg

Roll angular velocity deg/sec

Yaw angular velocity deg/sec

Pitch angular velocity deg/sec

Table 2.9-4: Format IV - Vector Data At Separation Epoch

Item Units

Launch time Date and time(hr, min, sec (GMT))

Separation time Date and time(hr, min, sec (GMT))

Estimated spacecraft separation time Date and time(hr, min, sec (GMT))

Semi-major axis km

Eccentricity ---

Inclination deg


Argument of perigee deg

Argument of latitude deg

Perigee altitude km

Apogee altitude km

Spacecraft +X axis right ascension deg

Spacecraft +X axis declination deg

Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation isthe estimated moment of spacecraft separation.


Page 2-28

Table 2.9-5: Format V - State Vector Data At Separation Epoch

Item Units

Fourth Stage third ignition time(UTC) for deorbitmaneuver

Date and time(hr, min, sec (GMT))

Fourth Stage third burn cutoff time (UTC) for deorbitmaneuver

Date and time(hr, min, sec (GMT))

De-orbit DV m/s

Note: Osculating elements of the orbit are referred to True Equator and Equinox of the liftoff epoch. The moment of osculation isthe estimated moment of spacecraft separation.

Within two days following separation, the SC contractor will provide spacecraft derived state vector data to ILS asshown in Table 2.9-6.

The SC contractor will provide SC rotation data about the SC X,Y and Z axes from the point of SC acquisition until15 minutes after acquisition. This data will be provided at [Launch + 15 Days].

Table 2.9-6: Spacecraft Supplied Separation Data

Spacecraft Supplied Separation Data

Parameter Units

Separation date (GMT) MDY

Separation time (GMT) HMS

Semi-major axis km

Eccentricity --

Inclination deg


True anomaly deg

Perigee radius km

Apogee radius km

Longitude of ascending node deg

Spacecraft separation spin axis relative right ascension deg

Spacecraft separation spin axis declination deg


Page 2-29

2.10 MISSION OPTIMIZATION/ PERFORMANCE ENHANCEMENTS

2.10.1 Nonstandard Mission Designs

The missions presented in the previous sections represent standard missions for the Proton K and M Boosters andBlock DM and Breeze M Upper Stages. ILS can take advantage of the unique capabilities of the Proton launch systemto design nonstandard mission profiles to meet unique mission and payload requirements.

2.10.2 Subsynchronous Transfer

Perigee velocity augmentation by the spacecraft can be used to increase payload weight to high Earth orbits. Thismaneuver consists of separating the spacecraft in a transfer orbit whose apogee is less than its final desired value. Thespacecraft then performs one or more burns to raise apogee, as well as a final burn to raise perigee. This maneuver isgenerally not used with the Proton launch system due to the Proton’s high throw weight capability, but it can beinvestigated if desired.

2.10.3 Super Synchronous Transfer

Use of a super synchronous transfer trajectory can increase performance into 24-hr orbits. The super synchronoustransfer trajectory takes advantage of the increased efficiency with which the inclination change can be performed at ahigher altitude. A number of vehicle hardware and satellite operational constraints interact with missions utilizingsupersynchronous transfer. ILS is able to asses supersynchronous missions by specific request and with detailed satelliteconfiguration data.

The long coast life and multiple restart capabilities of the Block DM and Breeze M can also assist in constellationphasing, thereby reducing spacecraft propellant usage.

ILS also has the resources available to develop special hardware items, such as dispensers for multiple spacecraft, tomeet unique mission requirements.


Page 3-1

3. SPACECRAFT ENVIRONMENTS

This section provides the ground and flight environments applicable for a Proton launch campaign and flight.

3.1 THERMAL/HUMIDITY

The thermal and humidity environment for the spacecraft is defined in this section, from transportation from theYubeleini Airport through launch base processing, launch and separation. SC component temperatures to be used forassessing thermal compatibility will be calculated by analysis using a SC thermal model provided by the Customer.

3.1.1 Ground Thermal Environment

Ambient temperatures at the Baikonur Cosmodrome are provided in Table 3.1.1-1. Facility and transportationtemperatures are provided in Table 3.1.1-2. During transport, the spacecraft is air-conditioned either by Customerprovided air-conditioning equipment or by a railcar mounted thermal control unit. While on the pad, thermal controlis provided by a pad air conditioning system and/or a liquid thermal control system (refer to Section 3.1.1.1 for adescription of the on-pad systems).

Table 3.1.1-1: 3σσ Ambient Temperatures at the Baikonur Cosmodrome

Month Max (OC) Min (OC)

January 8 -40

February 12 -38

March 24 -28

April 35 -12

May 42 0

June 46 8

July 46 9

August 42 8

September 38 0

October 30 -12

November 22 -30

December 13 -40


Page 3-2

Table 3.1.1-2: Spacecraft Thermal Environment

Location/event Item Temperature(deg. C)

Temperature Control

Min. Max.Spacecraft Container Inlet air to

container15 30 External air-conditioning unit provided by ILS/Khrunichev.

Flow rate of 2000 to 8000 cubic m/hr. Monitoring of thetemperature environment in S/C container is the SCcontractor’s responsibility. *

Fueling Hall92A-50Bldg 44

Spacecraftambient air

1517

2527

Building air-conditioning*

Payload Processing Area92A-50Area 31 rm 119


1517

2527


Fourth Stage Integration Area92A-50Area 31 rm 100A


1517

2527


Upper Stage fueling station/(Breeze M only)


15 25 External air-conditioning unit provided by ILS/Khrunichev.Maximum flow rate to 4000 m3/hr. *

Bldg 92-1 Spacecraftambient air

15 25 External air-conditioning unit provided by ILS/Khrunichev.Maximum flow rate to 4000 m3/hr. *

Encapsulated in fairing duringtransportation


15 25 External air-conditioning unit provided by ILS/Khrunichev.Preconditioning will be established by Khrunichev (takinginto account SC contractor recommendations) prior totransport to pad to guarantee temperature range duringerection. Maximum flow rate to 6000 m3/hr. *

Erection Spacecraftambient air

10 30 No active temperature control is provided.

On-pad with air-conditioningthrough umbilicalBlock DMBreeze M


1313

2723

Air-conditioning through umbilical, flow rates adjustable

from 5000 to 13000 M3/hr.*

On-pad following removal ofair-conditioning umbilical


10 30 Temperature control panels in fairing.Temperature could decrease to -5 deg. in Boattail area.

*Temperature is to be preset within the indicated range per SC contractor request and maintained accurate to + 2OC.


Page 3-3

3.1.1.1 On-Pad Thermal Control

An external thermal insulation shroud is placed around the fairing prior to pad rollout to provide additional insulationduring the erection of the Launch Vehicle on the pad when there is no active air conditioning. During transportation tothe pad, conditioned air is provided to the spacecraft from the Thermal Control Railcar. At the pad, the airconditioning is disconnected and the Launch Vehicle is erected. Following erection, the Mobile Service Tower isbrought up to the Launch Vehicle and the pad air conditioning is connected. This pad air conditioning is known as the“ASTR” for Air System, Thermal Regulation. Total time between disconnection of the Thermal Control Railcar airconditioning to the connection of on pad air conditioning is 4 hours maximum. A thermal analysis is performed toverify that under worst case ambient conditions, the spacecraft temperature will not exceed allowable temperaturelimits during the erection process.

The on-pad air conditioning system remains active 24 hours a day until approximately 1.5 hours prior to launch whenpreparations are begun for Mobile Service Tower rollback. To provide thermal conditioning of the fairing after MobileService Tower rollback, a liquid thermal control system is provided in the fairing. This system is known as the“LSTR” for Liquid System, Thermal Regulation. It consists of radiators on the fairing inside wall connected toethylene glycol filled pipes which run to a thermal control system in the launch pad complex. This system is activated 3hours prior to launch and purged with dry nitrogen 5 minutes prior to launch to insure that the lines are free of liquidprior to liftoff. Should the launch be aborted, the liquid system can be quickly reactivated and the Mobile ServiceTower will be brought up to renew air-conditioning within 2 hours. A schematic of both the liquid and air thermalcontrol systems is shown in Figures 3.1.1.1-1a and 3.1.1.1-1b along with an approximate operational timeline, forboth the Block DM and the Breeze M upper stages.


Page 3-4

Figure 3.1.1.1-1a: Fairing Air and Liquid Thermal Control System Schematic and Operations Timeline (Block DM)

Air system,thermal regulation

ASTR

Air inlet into ascent unit

T = 10 to +40oC

Launch cancellation

deg. C

27

22

17

12

Winter

Summer

Upper admissible temperature

Preparation tolaunchPreliminary

preparation

(~1.5 hr) (4-5 days) (2 hr) (1.5 hr) (2 hr)

Time

10' 5'

Launch

Thermalcover

mounting

Switchoff

thermo-stating

car

Switchon

ASTR

Switchon of

LSTR

Switchoff

ASTR

Switchoff

LSTR

Switchon

ASTR

Switchoff

LSTR

(15')(~4 hr)

Thermalcover

removal

Switchon

LSTR

Liftoff

Liquid system,thermal regulation

LSTR

Liquid flow rate = 0.250 - 0.9 m3/h

T = -10 to +80oC

Adapter

Temperaturesensors

emissivity = <0.1Air outlet

Heat exchanger panels

AscentUnit

~H/2

H

Air temperaturesensors

Air flow rate = 13,000 m3/h max

SC

Launch Vehicle installedon launch pad

Transportation 92-1 to

launch pad

(12-28 hr)


Page 3-5

Figure 3.1.1.1-1b: Fairing Air and Liquid Thermal Control System Schematic and Operations Timeline (Breeze M)

Air out let

Heat ex chang er panels

A dapt er

emissivity = <0.1

Launch cancellation

deg . C

27

22

17

12

Winter

Summer

Upper admissible temperature

Preparation tolaunchPreliminary

preparation

(12-28 hr) (~1.5 hr) (4-5 days) (2 hr) (1.5 hr) (2 hr)

Time

10' 5'

Launch

Thermalcover

mounting

Switchoff

thermo-stating

car

Switchon

ASTR

Switchon ofLSTR

Switchoff

ASTR

Switchoff

LSTR

Switchon

ASTR

Switchoff

LSTR

(15')

Transportation 92-1 to

launch pad

(~4 hr)

Thermalcover

removal

Switchon

LSTR

Liftoff

Air system,thermal regulation

ASTR

Air inlet into ascent unit

T = 10 to +40oC

Air flow rate = 13,000 m3/h max

LSTR

Liquid system,thermal regulation

Liquid flow rate = 0.250 - 0.9 m3/h

T = -10 to +80oC

Launch V ehicle installedon launch pad

Temperaturesensors

AscentUnit

Air temperaturesensors

SC


Page 3-6

3.1.1.2 Supplemental Air Conditioning for Spacecraft Battery Charging

Supplemental SC air-conditioning can be provided on the pad during SC battery charging operations if required. Upto 480 m3/hr can be provided through 2 access doors in the fairing at a temperature selectable between 10 and 16 deg.C. up until the time of Mobile Service tower rollback 1.5 hours prior to launch. The air inlets can be positioned inexisting fairing access doors in the aft section of the fairing. See Figure 3.1.1.2-1 below.

Figure 3.1.1.2-1: Supplemental Fairing Air Conditioning Schematic (Representative; detailed design conducted percustomer request)

35o

30

Fair ing/Block DM Interface Plane

1200

2100 7.5o

300

30o

160

260

840

Fai ring Access Door

Battery Radiator Cooling System Nozzle Location

3.1.2 Ascent

During Ascent, the launch vehicle will be exposed to aerodynamic heating. Following fairing jettison, the space craftwill be exposed to solar radiation and free molecular heat flux. A thermal analysis will be performed using theCustomer supplied SC thermal model to predict spacecraft temperatures during this phase of the mission. The heatflux density radiated upon the spacecraft by the internal surfaces of the NF should not exceed 500 W/M2 from the timeof launch until NF jettison. For commercial missions, the fairing is jettisoned at 342 - 344 seconds (121 – 125 kmalt.) into flight and the free molecular heat flux shall not exceed 1135 W/M2 at any time following fairing jettison.


Page 3-7

3.1.3 Orbit

Following injection into parking orbit, the spacecraft thermal environment is determined mainly by solar radiation,albedo and infrared earth radiation flux. Reorientation maneuvers of the Fourth Stage can be programmed to providedesired sun angles for maintaining thermal control. An integrated thermal analysis is performed to determinespacecraft temperatures as a function of time throughout the flight up to spacecraft separation.

3.1.4 Humidity

The ground humidity environment is shown in Table 3.1.4-1 below.

Table 3.1.4-1: Ground Humidity Environment

Location/event Item Relaative Humidity(RH) %

Humidity Control

Min. Max.

Inside SpacecraftContainer

Inlet air to container 35 60 ILS/Khrunichev provided air-conditioning

Fueling Hall Building air 35 60 Building air-conditioning

Payload Processing Area Building air 35 60 Building air-conditioning

Fourth Stage Integration Building air 35 60 Building air-conditioning

Encapsulated in fairingduring transportation

Air inside fairing 35 60 ILS/Khrunichev provided air conditioning

Fairing conditioning inBldg 92-1

Air inside fairing 35 60 ILS/Khrunichev provided air conditioning

Erection Air inside fairing 35 60 None

On-pad with air-conditioning throughumbilical

Air inside fairing 0.5 60 Air-conditioning through umbilical, flow rates

adjustable from 5000 to 13000 M3/hr.Max dewpoint approx 0 deg. C.

On-pad following removalof air-conditioningumbilical

Air inside fairing 0.5 60 None

Note: Should the relative humidity drop below 30%, SC contractor will be consulted prior to any work beginning in the vicinity ofthe spacecraft (inside the payload fairing after encapsulation).

3.1.5 Air Impingement Velocity

The air impingement velocity on the spacecraft surfaces shall not exceed 3m/sec during ground operations followingencapsulation through launch.


Page 3-8

3.2 CONTAMINATION ENVIRONMENT

3.2.1 Ground Contamination Control

The contamination environment around the spacecraft is controlled by use of Class 100 000 clean room facilities andstrict control of material cleanliness of flight hardware in proximity to the spacecraft. During transportation using ILSprovided air-conditioning systems and while on the pad, air is filtered to provide better than Class 100 000 particlecontent. Air cleanliness is monitored regularly in all areas where the SC is present to ensure particle count levels aremaintained within specification. In addition, witness plates can be mounted inside the fairing following encapsulationto monitor particle fallout inside the fairing up until the day prior to launch. The ground contamination environmentaround the spacecraft meets the cleanliness levels specified in Table 3.2.1-1.

Table 3.2.1-1: Ground Contamination Environment

Location/event Cleanliness LevelRequired*

Comments

Spacecraft container 100 000 ILS/Khrunichev supplied conditioned air.

SC Processing Facilities 100 000 Facility air conditioning

Bldg 92-1 100 000 Payload encapsulated, filtered air provided

Encapsulated in fairing duringtransportation and battery charging

100 000 Payload encapsulated, filtered air provided

Erection 100 000 Payload compartment sealed

On-pad with air-conditioning throughumbilical

100 000 Filtered air provided

On-pad following removal of air-conditioning umbilical

100 000 Payload compartment sealed

* Per FED Std 209E

3.2.2 In Flight Contamination Control

The Launch Vehicle systems are designed to preclude in-flight contamination of the SC. The Launch Vehiclepyrotechnic devices near the spacecraft used for fairing jettison and SC separation have sealed gas chambers and do notrelease significant contamination to the outside environment. The fairing liquid thermal control system pipes aresealed by automatic valves which close at fairing jettison. The third stage retro rocket plume does not result in anysignificant particle contact with the spacecraft due to their position on the aft end of the third stage and the orientationof the retro rocket plume 15 degrees away from the Launch Vehicle longitudinal axis. The thrusters are located on theaft section of the Fourth Stage and oriented perpendicular to the attitude control longitudinal axis (worst case) suchthat the plume does not contact the spacecraft while the spacecraft is attached to the Fourth Stage. Followingseparation, the Fourth Stage is attitude controlled to prevent reorientation relative to the spacecraft until the spacecraftis a safe distance away from the Fourth Stage. Finally, an evaporative cooling system in the Fourth Stage ejects a 20percent alcohol/water vapor periodically following third/fourth stage separation at the aft end of the Fourth Stage.This gas is diverted away from the spacecraft and therefor has no contaminating influence.


Page 3-9

3.3 PRESSURE

3.3.1 Payload Compartment Venting

During ascent, the payload compartment is vented through 4 venting orifices distributed equally about the cylindricalportion of the fairing. Maximum rate of pressure drop in the fairing will not exceed 3.5 kPa/s. A representativepressure drop profile inside the fairing during flight is given in Figure 3.3.1-1.

At the moment of fairing jettison, the pressure across the fairing halves shall not exceed 700 Pa.

The archimedes volume of the spacecraft to be taken into account for the venting analysis will be provided by theCustomer.

Figure 3.3.1-1: Typical Venting Profile During Ascent

Time, sec Pressure (Pascals)Atmosphere Inside Fairing

0 100000 10000010 98000 9900020 91000 9300030 80000 8300040 65000 7000050 49000 5500060 32500 3850070 19300 2650080 10000 1700090 5000 11000

100 2500 7500110 1000 5000120 0 3500130 0 2000

0.00E+00

2.00E+04

4.00E+04

6.00E+04

8.00E+04

1.00E+05

1.20E+05

0 10 20 30 40 50 60 70 80 90 100 110 120 130

Time, seconds

Pas

cals

Atmosphere

Inside Fairing


Page 3-10

3.4 MECHANICAL LOADS

The spacecraft is subject to various types of mechanical loads due to transportation and handling during the launchcampaign as well as due to the various flight events following liftoff. This section breaks down these events by type ofexcitation: quasi-static, sine/random, acoustic and shock.

3.4.1 Quasi-Static Loads

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 primarystructure and/or evaluation of compatibility for existing spacecraft with the Proton launch vehicle. The load factorswere derived for application at the center of gravity of a rigid spacecraft and generate a conservative estimate of flightmaximum interface axial, shear, and bending moments. The actual interface responses experienced during flight are afunction of the static and dynamic characteristics of the spacecraft, but these load factors have generally provenconservative for spacecraft in the weight range of 2000 - 4800 kg with c.g. height (above the interface) between 1.0 and

2.0 meters and where 85.0I

zm>

−; where: m=mass, z=c.o.g. location, I=maximum lateral moment of inertia. The

spacecraft cantilevered fundamental frequencies are assumed to be a minimum of 10 Hz lateral and 25 Hz axial toinsure applicability of both the ground Transport and the Flight load factors. Spacecraft that do not meet these criteriawill require preliminary analyses to generate loads environments for assessing compatibility with Proton. The Protonload factors are conservative, but do not include uncertainty factors. It is recommended that the spacecraft analyst useuncertainty factors in preliminary sizing of primary structure if the new design is not within the family of spacecraftintegrated on Proton.

3.4.1.1 Ground Loads

During ground transportation and handling operations, the spacecraft will be subjected to low frequency loads. Table3.4.1.1-1 provides the bounding quasistatic loads at the S/C center of mass in longitudinal and lateral axes.

Table 3.4.1.1-1: Ground Limit Quasistatic Load Environment-Transportation and Handling Operations

Operations Accelerations, (g) Safety FactorX Y Z

SC in container transported as a separate item + 1.0 1.0 + 1.0 + 0.4 1.5

SC transported as part of the Ascent Unit +0.5 1 + 0.5 + 0.4 1.5

SC transported as part of the Proton LV +0.4 1.0 + 0.3 + 0.15 1.5

Handling 0.15 1 + 0.5 1.5

Notes:In the transportation case, the axes are those of the transporter, namely:a) X axis runs along the direction of movement.b) Y axis is downward in direction of gravity.c) Z is lateral axis in right hand frame.

For handling:a) Y axis runs vertically along the lifting and lowering direction, respectfully.b) X axis runs along any lateral directions.c) Accelerations exist simultaneously along X,Y,Z.d) All accelerations are specified for a wind velocity W< 20 m/s.


Page 3-11

3.4.1.2 Flight Loads

In flight, the spacecraft will be subjected to low frequency input at the base of the spacecraft. The events which producethe worst case loading are liftoff and 1st/2nd stage separation. Figure 3.4.1.2-1 provides the bounding quasistatic loadsin longitudinal and lateral axes at the SC center of mass. Loads along perpendicular axes are independent.

Static and dynamic accelerations at the SC interface are measured by flight instrumentation and their maximum 3σvalues are as provided in Figure 3.4.1.2-2. Flight loads shall be evaluated by Coupled Loads Analysis using a dynamicmodel incorporating the spacecraft and launch vehicle.

Figure 3.4.1.2-1: Quasi-Static Design Load Factors

-4

-3

-2

-1

0

1

2

3

4

5

-3 -2 -1 0 1 2 3

Lateral, g

(Axi

al)

Lond

itudi

nal,

g

Block DM

Breeze M


Page 3-12

Figure 3.4.1.2-1: Quasi-Static Design Load Factors (Continued)

Block DM Quasi-Static Design Load Factors

Event Longitudinal, g Lateral, g

Lift-off 2.3 2.3 -2.30.3 2.3 -2.3

Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.21st/2nd stages before separation 4.3 1.4 -1.41st/2nd stages after separation (max compression) 3 1.4 -1.41st/2nd stages after separation (max tension) -2.8 1.4 -1.42nd/3rd stages separation 3 0.3 -0.33rd/4th stages separation 2.8 0.3 -0.3

Breeze M Quasi-Static Design Load Factors

Event Longitudinal, g Lateral, g

Lift-off 2.3 1.35 -1.350.3 1.35 -1.35

Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.21st/2nd stages before separation 4.3 0.9 -0.91st/2nd stages after separation (max 3 0.9 -0.91st/2nd stages after separation (max tension) -3 0.9 -0.92nd/3rd stages separation 3 0.3 -0.33rd/4th stages separation 2.8 0.3 -0.3


Page 3-13

Figure 3.4.1.2-2: Flight Limit Accelerations at the SC Interface

Event Longitudinal Longitudinal TransverseStatic g Dynamic g Dynamic g

Liftoff 1.5 1.5 -1.5 1.1 -1.1Wind and Blast (Qmax) 2.2 0.5 -0.5

Separation 1st/2ndBefore 1st stage booster separation 3.6 0.9 -0.9 0.9 -0.9

stagesAfter 1st stage booster separation 1 2 -2.8 0.9 -0.9

2nd stage engine cutoff 2.7 0.3 -0.3 0.3 -0.33rd stage engine cutoff 2.5 0.3 -0.3 0.5 -0.5

-3

-2

-1

0

1

2

3

4

5

-3 -2 -1 0 1 2 3

Transverse,g

(Axi

al) L

ongi

tudi

nal,g

Note: Used only to evaluate measured flight loads.


Page 3-14

3.4.2 Sine and Random Vibration Loads

At launch, when the propellant valves of the first stage engines are opened, reactive forces act on the liquid propellantin the tanks (for approximately 0.1 sec) causing launch vehicle lateral oscillations on the elastic pad supports.Prevailing oscillation frequencies are approximately 4 Hz with amplitudes of 0.3 g.

The engines operate at a preliminary thrust level that remains constant for approximately 1.6 sec. During this period,the Launch Vehicle experiences flexible bending oscillations brought about by uneven thrust among the six engines andunequal offloading of the pad supports. The prevailing frequencies are 5 to 7 Hz.

Longitudinal flexible body oscillations appear simultaneously with frequencies ranging from 5 to 15 Hz. They aremagnified as the engines are throttled up to full thrust within 0.5 sec as the Launch Vehicle leaves the pad.

During first stage flight, lateral dynamic loads are generated by wind gusts superimposed on steady state wind loadsgenerated by the jetstream. Launch Vehicle longitudinal flexible oscillations are produced at 10-12 Hz by the naturalrandom pulsation of the engine thrust. There is no pogo phenomenon. The maximum value of these oscillations basedon telemetry measurements is +/- 0.35 g’s.

From 0.5 to 0.6 sec before first stage cutoff, the four second stage engines start up and gain preliminary thrust. Becauseof the uneven thrust of the four engines, lateral reaction forces are generated, causing lateral flexible oscillations of theLaunch Vehicle body. These oscillations are influenced additionally by the first stage engines reacting to control systemcommands. The first stage cutoff is characterized by an abrupt decay from 90% to 20% within 0.03 sec which causessignificant flexible longitudinal oscillations of the Launch Vehicle second stage, driven by the preliminary thrust of itsown engines. The oscillations are additionally magnified due to the increase in thrust to 100%. These oscillations dampout within about 3 seconds.

Dynamic loads occuring during the propulsive events following first/second stage separation are enveloped by thepreceding events.

The above dynamic load environment can be represented by the quasi-sinusoidal vibration environment applied at theSC/LV interface plane shown in Figure 3.4.2-1. This can be considered a flight limit load environment.


Page 3-15

Figure 3.4.2-1: Equivalent Sine Levels at Spacecraft Interface - Flight Environment

Lateral Axes

Frequency (hz) Level (g's)5 10 0.3

10 20 0.420 40 0.640 100 0.7

Lateral Axes

0

0.2

0.4

0.6

0.8

1

1.2

0 20 40 60 80 100 120

Frequency (hz)

Leve

l (g'

s)

Longitudinal Axes

Frequency (hz) Level (g's)

5 8 0.38 20 0.8

20 40 0.640 100 0.9

Longitudinal Axes

0

0.2

0.4

0.6

0.8

1

1.2

0 20 40 60 80 100 120

Frequency (hz)

Leve

l (g’

s)


Page 3-16

During ground transportation, random excitation is produced by the rail system. The random vibration levels fordifferent 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.

Figure 3.4.2-2: Random Vibration Levels-Ground Transportation By Rail, SC In Container And SC Attached ToAscent Unit

Frequency (hz) Notes:

X-X Y-Y Z-Z X axis is in the direction of movement

2 0.000075 0.000150 0.000150 Y axis is in the vertical direction parallel to gravity field

4 0.000575 0.003300 0.000330 Z axis is in the direction that provide right a handed set

8 0.002000 0.003200 0.00066010 0.000600 0.003200 0.00080014 0.000280 0.000833 0.00033020 0.000275 0.000150 0.00032025 0.000275 0.000150 0.00031030 0.000275 0.000150 0.000300 *Durations are as follows:

35 0.000500 0.000150 0.000185 Transport SC Container Yubeleini to Area 31 120 minutes

40 0.000180 0.000150 0.000037 Transport SC Container Yubeleini to 92A-50 60 minutes

45 0.000125 0.000150 0.000037 Transport Ascent Unit Area 31 to 92-1 120 minutes

50 0.000125 0.000150 0.000037 Transport Ascent Unit 92A-50 to 92-1 5 minutesDuration (minutes) * * *

PSD (G2/hz)

- Transportation velocity less than or equal to 15 km/hr - Levels are at Container Base for container transportation and at SC/adapter interface for transportation as part of Ascent Unit

0.00E+00

5.00E-04

1.00E-03

1.50E-03

2.00E-03

2.50E-03

3.00E-03

3.50E-03

1 10 100Frequency (Hz)

Pow

er S

pect

ral D

ensi

ty (G

2/hz

)

X-X

Y-Y

Z-Z


Page 3-17

Figure 3.4.2-3: Transportation of SC in Contractor’s Container, Transportation of SC in KhSC Container,Transportation of Ascent Unit

y

x

y

x

y

x


Page 3-18

Figure 3.4.2-4: Random Vibration Levels-Ground Transportation by Rail, SC and Ascent Unit Attached to L/V

Frequency (hz)X-X Y-Y Z-Z

2 0.000002 0.000002 0.0000024 0.000002 0.000002 0.0000028 0.000002 0.000002 0.000002

10 0.000002 0.000002 0.00000214 0.000002 0.000020 0.00000220 0.000010 0.000001 0.00001025 0.000001 0.000001 0.000001 Notes:

30 0.000001 0.000001 0.000001 X axis is in direction of movement

35 0.000003 0.000001 0.000001 Y axis is in vertical direction parallel to gravity field

40 0.000001 0.000001 0.000001 Z axis is in direction to provide right handed set

45 0.000001 0.000001 0.000001 Transportation velocity less than or equal to 10 km/hr

50 0.000001 0.000001 0.000001 Levels are at SC to L/V interfaceDuration (minutes) 10 10 10

PSD (G2/hz)

0.00E+00

2.00E-06

4.00E-06

6.00E-06

8.00E-06

1.00E-05

1.20E-05

1.40E-05

1.60E-05

1.80E-05

2.00E-05

1 10 100

Frequency (Hz)

Pow

er S

pect

ral D

ensi

ty (G

2 /h

z)

X-X

Y-Y

Z-Z


Page 3-19

Figure 3.4.2-4: Transportation of Integrated Proton LV

y

x

3.4.3 Acoustic Loads

The launch acoustic loads arise from acoustic sound waves generated by the supersonic jets from the first-stage enginenozzles being diverted by the launch pad and flame deflectors. At transonic velocity and maximum aerodynamic drag,acoustic loads are caused by aerodynamic pressure pulsation effects on the payload fairing surface. The peak acousticloads do not act longer than 3 sec at liftoff and 50 sec while passing through the zone of maximum aerodynamic drag.Acoustic load characteristics normalized to the threshold pressure of 20 µPa are shown in Figure 3.4.3-1.


Page 3-20

Figure 3.4.3-1: Max Expected Acoustic Environment (Third Octave)

1/3 Octave Band Center Frequency (hz)

Acoustic Levels on Spacecraft (dB)

25 11731.5 12340 12750 126.563 127.880 131.6

100 132.4125 131.3160 132.1200 132.1250 132.1315 129.9400 129500 127630 124800 121

1000 1191250 1171600 114.52000 112.52500 1113150 1094000 1085000 1076300 105.58000 104

10000 103.5OASPL 141.4

100

110

120

130

140

10 100 1000 10000

Frequncy (Hz)

Leve

l (dB

)


Page 3-21

3.4.4 Shock Loads

Worst case shock levels are introduced into the spacecraft during the firing of the SC/adapter separation system. Thelevel is dependent on the type of clampband and the clampband tension. For the existing standard adapterconfigurations, three specific levels may be encountered as indicated in Figure 3.4.4-1. The 1194 separation systemsuse either a 26.6 kN or a 40 kN preload and the shock levels differ accordingly. The 1666 separation systems use a 30kN preload and the corresponding shock level is as indicated.

Shock loads during transportation in the SC container at the launch site shall not exceed the levels provided in Table3.4.4-1.

Table 3.4.4-1: Shock Loads During Transportation in SC Container

Event Acceleration amplitude, g

X axis runs along the direction of motion + 1.0

Y axis upward vertical axis 1.0 + 1.0

Z lateral axis + 0.4

One Semisinusoidal impulse duration, ms 30

Note: Assumes 5% damping (Q=10)


Page 3-22

Figure 3.4.4-1: Pyroshock Spectrum at Adapter/Payload Interface

100

1000

10000

100 1000 10000Frequency (Hz)

g's 40 kN 1194 Sep Syst.

26.6 kN 1194 Sep Syst.

30 kN 1666 Sep Syst.

Q= 10

Frequency (Hz) g's

40 kN 1194 26.6 kN 1194 30 kN 1666

Sep Syst. Sep Syst. Sep Syst.

100 150 100 150

600 - 1800 -

800 3000 - 3000

1200 5500 - -

1500 6500 - -

2000 5000 5000 -

3000 - - 3000

4000 5000 - -

6000 - - 3500

10000 8000 5000 4000


Page 3-23

3.4.5 Environmental Test Requirements

3.4.5.1 Acoustic Test Requirements

The spacecraft test requirements are as follow:

Table 3.4.5.1-1: Acoustic Test Requirements

Type of Test Levels Test Duration (seconds)

Acceptance Minimum of Figure 3.4.3-1 in each band 60

Qualification Acceptance levels + 3 db 120

Protoqualification Acceptance levels +3 db 60

3.4.5.2 Static and Sine Test Requirements

Static testing of primary structure is required as a qualification of the structure for flight. This static test mustdemonstrate the capability of the structure to withstand the worst case combination of quasi-static loads shown inTable 3.4.1.1-1 and Figure 3.4.1.2-1 or obtained from coupled loads analysis. Qualification margins for structurestatic testing of 1.1 to yield and 1.25 to ultimate must be assumed. For ground handling lift points, qualificationmargin to ultimate shall be 1.5 minimum.

Demonstration of the spacecraft secondary structures ability to withstand dynamic loads induced by flight events andground transportation is required for qualification and acceptance of the spacecraft design.

The following tests are required:

Table 3.4.5.2-1: Sine Test Requirements

Type of Test Levels Test Frequency Sweep Rate(octaves/min)

Sine sweep qualification Sine levels in Figure 3.4.2-1 x 1.25 2

Sine sweepprotoqualification

Levels in Figure 3.4.2-1 x 1.25 4

Sine sweep acceptance Levels in Figure 3.4.2-1 4


Page 3-24

Notching should be minimized, and as a test goal should not allow base inputs to decrease below the base equivalentlevel produced by the Final Coupled Loads Analysis multiplied by the appropriate Factor of Safety and should notallow spacecraft response to go below the worst case response predicted by the Final Coupled Loads Analysismultiplied by the appropriate Factor of Safety.

For Spacecraft secondary structure which does not attain the worst case level predicted from the Final Coupled LoadsAnalysis multiplied by the appropriate Factor of Safety, an analysis must show the capability to sustain the CLA resultmultiplied by the appropriate Factor of Safety.

3.4.5.3 Shock Test Requirements

A shock test using the adapter clamp system with the flight adapter and spacecraft is required at the SC manufacturer’sfacility in conjunction with a mechanical/electrical fitcheck for first of a kind spacecraft and the first follow-onspacecraft in a series. For this test the flight band will be tensioned to the same level as will be used for the flightinstallation per the manufacturers test procedure. Shock levels will be measured at a location on the adapter 120 mmabove the SC separation plane.

3.5 ELECTROMAGNETIC COMPATIBILITY

Electromagnetic Interference (EMI) emissions and susceptibility of the SC and the LV shall be individually controlledto the extent necessary to ensure EMC of the fully integrated system.

3.5.1 EMI Safety Margin (EMISM)

The integrated SC/LV system shall be designed to provide EMC with a minimum of 20 dB EMISM (vs. DC no-fireThreshold) for ordnance circuits and 6dB EMISM overall.

3.5.2 Radiated Emissions

The launch vehicle intentional emissons are described in Tables 3.5.2-1 and 3.5.2-2a and 3.5.2-2b. The SC needs to becompatible with these emissions.

The launch vehicle generated and launch base spurious EMI sources shall not exceed the levels of Figures 3.5.2-1a and3.5.2-1b in a plane 1 meter below and parallel to the SC/LV interface plane.

The SC generated and spurious EMI sources shall not exceed the levels of Figure 3.5.2-2 in a plane 1 meter below andparallel to the SC/LV interface plane.


Page 3-25

Table 3.5.2-1: Launch Vehicle RF Characteristics for Proton K/Block DM (TBC)

Parameters First StageTransmitter 1

First StageTransmitter 2

Second StageTransmitter

Third StageTransmitter 1


Block DMTransmitter

Block DMReceiver

Carrier Frequency (Mhz) 192 137 240 232 132 923 769

3db Bandwidth (Mhz) .256 3.0 .256 .256 3.0 0.5 0.5

Modulation type andcharacteristics PM APM/FM PM PM APM/FM PM PM

Transmitter output power(dbW) Max (at carrier freq.) 20.8 11.8 20.8 20.8 11.8 9.0

Receiver sensitivity at thecarrier freq (dbW), Nom. -137

Antenna gain, db -10 -10 -10 -10 -10 -10 9

Antenna description andpolarization

Circular

OMNI EP

Circular

OMNI EP

Circular

OMNI EP

Circular

OMNI EP

Circular

OMNI EP

Circular

OMNI EP

Circular

OMNI EP

Operating on pad? Yes Yes Yes Yes Yes Yes Yes

Operating in flight? Yes Yes Yes Yes Yes Yes Yes

Table 3.5.2-2: Launch Vehicle RF Characteristics for Proton M (TBC)

Parameters First StageTransmitter 1

First StageTransmitter 2

Second StageTransmitter



Tracking

Carrier Frequency (MHz) 192 137 240 232 132 3410

3db Bandwidth (MHz) 0.256 3.0 .256 .256 3.0 40

Modulation type andcharacteristics PM APM/FM PM PM APM/FM PM

Transmitter output power(dbW) Max (at carrier freq.) 20.8 11.8 20.8 20.8 11.8 9.0

Receiver sensitivity at thecarrier freq (dbW), Nom.

Antenna gain, dB -10 - 10 - 10 - 10 - 10 -9 within angles of+70 degrees

-10 within angles of+85 degrees


Omni-directional, circular limited view omniantenna

Operating on pad? Yes Yes Yes Yes Yes No

Operating in flight? Yes Yes Yes Yes Yes Yes


Page 3-26

Table 3.5.2-3: Launch Vehicle RF Characteristics for Breeze M (TBC)

Parameters Breeze MTelemetry

Breeze MTelemetry

Breeze MTelemetry

Breeze MReceiver

Glonass/ GPS

Carrier Frequency (MHz) 15150 1018.5 1020.5 5750-5760 1570-1640

3db Bandwidth (MHz) 0.5 0.128 0.128 50 55

Modulation type andcharacteristics PM FM FM PM PM

Transmitter output power(dbW) Max (at carrier freq.) 9.0 11.8 11.8

Receiver sensitivity at thecarrier freq (dbW), Nom. - 137 - 137

Antenna gain , dB 21 3 to 10(Parallel to

Xsc axis)

3 to 10(Parallel to

Xsc axis)

-7 3


right circularhorn

limited viewomni antenna




Operating on pad? No Yes Yes No Yes

Operating in flight? Yes Yes Yes Yes Yes


Page 3-27

Figure 3.5.2-1a: Launch Vehicle and Launch Base Pad Narrowband Radiated Emissions (Proton K/Block DM) (TBC)

Frequency (MHz) dBmV/m 3dB Band width (MHz)132 98 3137 74 3192 84 0.256232 100 0.256240 87 0.256769 25 0.5923 120 0.5

2000 70 N/A100000 15 N/A

0

20

40

60

80

100

120

140

100 1000 10000 100000

Frequency (MHz)

dBm

V/m


Page 3-28

Figure 3.5.2-1b: Launch Vehicle and Launch Base Pad Narrowband Radiated Emissions (Proton M/Breeze M) (TBC)

Frequency (MHz) dBmV/m 3dB Band width (MHz)132 98 3137 74 3192 84 0.256232 100 0.256240 87 0.256

1000 127 0.52000 70 N/A3400 123 0.5

15150 135 0.530000 15 N/A

0

20

40

60

80

100

120

140

100 1000 10000 100000Frequency (MHz)

dBm

V/m


Page 3-29

Figure 3.5.2-2: Launch Vehicle and Launch Pad Radiated Susceptibility Limits

Frequency (GHz) dBmV/m0.01 1200.1 120

0.75 601 129

20 80

10

30

50

70

90

110

130

1.00E-01 1.00E+00 1.00E+01 1.00E+02Frequency (GHz)

dBm

V/m

3.5.3 RF Transmitter/Receiver Systems EMC

Standard RF system compatibility analyses will be performed which shall insure integrated system EMC ofsimultaneously operated SC and LV transmitters and receivers during time frames where such operations are necessary.


Page 4-1

4. SPACECRAFT INTERFACES

4.1 MECHANICAL INTERFACES

4.1.1 Structural Interfaces

The SC to LV structural/mechanical interfaces include a payload adapter, a separation system, umbilical connectors,separation switches and bonding straps. The structural/mechanical interfaces are defined for each adapter system inAppendix D of this Mission Planner’s Guide.

The LV coordinate system is shown in Figure 4.1.1-1 with a representative SC.

Figure 4.1.1-1: LV Coordinate System

4.1.2 General SC Structural and Load Requirements

4.1.2.1 Design Criteria

The SC and LV interface structure shall support the SC during the maximum load condition without yielding. Theclearance between the flanges of the SC and the adapter prior to clampband tensioning shall not exceed 0.6 mm. Thegeometry of the spacecraft flange is provided in Appendix D for a temperature of 21° C. The surface flatness of the SCinterface ring shall be less than 0.3 mm.

4.1.2.2 SC Stiffness

The spacecraft primary structural stiffness shall be such that the minimum fundamental lateral and axial modefrequencies shall be greater than 10 Hz and 25 Hz respectively as cantilevered from a rigid interface. The SC/LVinterface is assumed to behave linearly under all loading conditions.


Page 4-2

4.1.2.3 SC Interface Loads

The SC lifting device and structure shall be capable of lifting the SC plus the payload adapter and the separationsystem. Maximum adapter, separation system and other mass to be lifted by SC = 220kg.

Loads affecting the SC at the SC/LV interface include the adapter springs and the SC/LV electrical umbilicalconnectors. The adapter spring forces and the SC/LV electrical umbilical connector forces are provided in the adapterAppendix D of this Mission Planner’s Guide.

4.1.2.4 SC Center of Gravity Offset Requirements

TBS

4.1.3 Fairing Interfaces

This section provides a description of fairing interfaces including generic fairing useable volume, allowable access doorlocations and RF window locations.

4.1.3.1 Fairing General Description

A general layout of the Proton M/Block DM Commercial Fairing is provided in Figure 4.1.3.1-1. A general layout ofthe Breeze M Standard and Long Commercial Fairings are provided in Figures 4.1.3.1-3 and 4.1.3.1-5.

Figure 4.1.3.1-2 provides the layout for the Proton/Block DM generic useable volume. Figures 4.1.3.1-4 and4.1.3.1-6 provide the layouts for the Proton/Breeze M Standard and Long Commercial fairing generic useablevolumes. These generic useable volumes do not take into account any specific adapter configuration. Specific adapterswill alter the bottom portion of the useable volume in order to take into account required adapter clearances forinstallation and required flight clearances with the adapter structure. Specific useable volumes tailored to individualadapter systems are provided in the adapter Appendix D of this Mission Planner’s Guide.

The definition of useable volume used throughout this Mission Planner’s Guide is as follows:

Useable Volume: The spacecraft static envelope (maximum dimensions of unloaded spacecraft, includingmanufacturing tolerances and expansion of thermal blankets) must not protrude beyond the useable volume, exceptwhere it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements due to ground or flightloads and deviations caused by an imperfect installation of the spacecraft on the Fourth Stage may protrude beyond theboundaries of this useable volume. An assumption made is that spacecraft dynamic displacements will not exceed 50mm. This must be verified by coupled loads analysis.


Page 4-3

4.1.3.2 Fairing Access Locations

Four fairing access doors are located on the lower boattail of the fairing structure of the Block DM vehicle and aredimensioned as shown in Figure 4.1.3.1-1. These doors are nominally used for access to the Block DM. The Customermay use these doors for access to spacecraft related interface equipment. These access requirements need to becoordinated and agreed upon with ILS in the mission specific ICD. Up to 2 access doors may be provided in thefairing 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 vehicleversions. These access locations may be accessed at times coordinated with ILS from the time of fairing encapsulationup to the beginning of Launch Vehicle fueling on the launch pad.


Page 4-4

Figure 4.1.3.1-1a: Proton/Block DM Commercial Fairing General Layout (Sheet 1 of 3)


Page 4-5

Figure 4.1.3.1-1b: Proton/Block DM Commercial Fairing General Layout (Sheet 2 of 3)


Page 4-6

Figure 4.1.3.1-1c: Proton/Block DM Commercial Fairing General Layout (Sheet 3 of 3)


Page 4-7

Figure 4.1.3.1-2a: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)


Page 4-8

Figure 4.1.3.1-2b: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)


Page 4-9

Figure 4.1.3.1-2c: Generic Proton/Block DM Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)


Page 4-10

Figure 4.1.3.1-3a: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 1 of 3)


Page 4-11

Figure 4.1.3.1-3b: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 2 of 3)


Page 4-12

Figure 4.1.3.1-3c: Proton/Breeze M Standard Commercial Fairing General Layout (Sheet 3 of 3)


Page 4-13

Figure 4.1.3.1-4a: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)


Page 4-14

Figure 4.1.3.1-4b: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)


Page 4-15

Figure 4.1.3.1-4c: Generic Proton/Breeze M Standard Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)


Page 4-16

Figure 4.1.3.1-5a: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 1 of 4)


Page 4-17

Figure 4.1.3.1-5b: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 2 of 4)


Page 4-18

Figure 4.1.3.1-5c: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 3 of 4)


Page 4-19

Figure 4.1.3.1-5d: Proton/Breeze M Long Commercial Fairing General Layout (Sheet 4 of 4)


Page 4-20

Figure 4.1.3.1-6a: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 1 of 3)


Page 4-21

Figure 4.1.3.1-6b: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 2 of 3)


Page 4-22

Figure 4.1.3.1-6c: Generic Proton/Breeze M Long Commercial Fairing - Useable Volume Dimensions (Sheet 3 of 3)


Page 4-23

4.1.4 GN2/Dry Air Purge Option

For an additional fee the Customer can obtain a GN2/dry air purge at the adapter interface via special pneumaticfittings. GN2 can be provided via Customer provided gas bottles during operations in the PPF and on the Launch Padup to Mobile Service Tower rollback. At this time, the line can be connected to a dry air source running through theLV to provide a dry air purge up to liftoff. Characteristics of this purge system are as follows:

Item Characteristic

Number of fittings 1

Type fitting pneumatic 0.172 inches internal diameter, 0.281 inches external diameter, 303 CRES material(provided by customer)

Period of operation a)accessible by Customer during payload operations in PPF up to on pad prior to MobileService Tower rollback (including during transportation operations as mutually agreed uponbetween ILS and Customer)

b)connected to ILS dry air source through LV from Mobile Service Tower rollback to launch

Operational gas Gaseous nitrogen or air

Gas Content

Particulate Size <50 microns

Hydrocarbon Content Maximum Condensable Hydrocarbons-

5.0X10-4% by mass

Helium Content At Standard Atmosphere Concentrations

5.0X10-4% maximum by volume

Filtration Preliminary purification and availability of filter at system outlet with mesh of 25-50 microns

Temperature -30 degrees C to +30 degrees C

Humidity Requirement Maximum Dew Point Temperature = -55 degrees C

Flow rate at SC/LV interface 450-650 cm3/min

Maximum pressure dropfrom SC/LV interfacethrough SC

0.048 Pa

For a typical mechanical interface layout, see Appendix D of this Mission Planner’s Guide.


Page 4-24

4.2 ELECTRICAL INTERFACES

Electrical interfaces include the SC/LV airborne interfaces, EGSE interfaces, and telemetry/command links.

4.2.1 Airborne Interfaces

Electrical umbilical interfaces are used primarily for providing power to the spacecraft from Customer ground powersupplies located in the Vault under the launch pad. They are also used for hardline telemetry and command linksbetween the spacecraft and the Customer ground support equipment located in the Bunker. Additionally, the Customerhas an option to provide spacecraft measurements to the Launch Vehicle telemetry system on the Fourth Stage viathese umbilicals.

4.2.1.1 Electrical Connectors

Two 37 or 61 umbilical connectors are provided at the Spacecraft interface with the Launch Vehicle adapter. Theconnectors are spring-loaded and at separation will disconnect from the adapter.

Appendix D of this Mission Planner’s Guide describes the standard adapters also provides the type, location andmechanical configuration for these connectors.

Umbilical connector brackets provide +/- 2 mm adjustment in longitudinal direction and +/- 4 mm adjustment inthe lateral axes.

4.2.1.2 Separation Verification

Two diametrically opposed separation microswitches are provided on the top adapter interface flange. Refer toAppendix D for specific locations and mounting configuration for each specific adapter. At separation, themicroswitches will open a circuit and the Launch Vehicle telemetry will detect this as the separation event.

In addition, continuity loops are provided in each umbilical connector on the spacecraft side. At separation, theumbilical connectors will disengage, thereby opening these circuits and providing a redundant indication of separationto the Launch Vehicle telemetry system.

4.2.1.3 Interface Electrical Constraints

All SC and LV electrical interface circuits shall be constrained at least 20 seconds prior to SC separation such that thereis no current flow greater than 100 mA per wire during the Separation event.

4.2.1.4 Accelerometer Measurements

Five accelerometers are mounted near the top of the adapter interface flange to record acceleration from liftoff untilstage three/four separation. Their positions are illustrated in Section 4.2.1.7. Three accelerometers measurelongitudinal loads and two measure lateral loads. Refer to Section 4.2.1.7 for characteristics of these telemetrychannels and Appendix D for the installation of the accelerometers on the adapter.

4.2.1.5 Separation Microswitches

Two diametrically opposed separation microswitches are provided, located on the top adapter interface flange. Refer toAppendix D for angular locations and for cross-section views.


Page 4-25

4.2.1.6 Pre-Separation “Dry Loop” Commands

The Customer may choose as an optional service up to 2 primary and 2 redundant in flight commands in the form ofrelay closures for initiating spacecraft commands during flight. The command for closure will be issued after launchand before SC/LV separation. Timing and signal characteristic requirements need to be provided by the Customer nolater than at L-12 months. Characteristics of this command are as follows:

Item Characteristic

Type of relay Single pull single throw

Actuation time Any time from liftoff up to separation

Pulse duration up to 210 ms

Timing accuracy +/- 0.03 ms

Allowable max voltage through relay contact at relay closure 36 Volts

Allowable max steady state current through SC/LV interfacecontact

1 Ampere

A schematic of this dry loop command is provided in Figure 4.2.1.6-1.

Figure 4.2.1.6-1: Dry Loop Functional Schematic

K7 K13 K8 K14

K11

K13

K7

2 6 1 5 11 15 12 16

K11

K14 K8

K9

K12

K15

K9

K15 K10

K12

K16

K16

K10

J1 J2

PO2A

SC/LV Interface


Page 4-26

4.2.1.7 Launch Vehicle Telemetry, Command and Power

The launch vehicle provides the spacecraft separation command and the power for initiating the separation system.There are no launch vehicle power or command lines which pass across the spacecraft separation plane.

Table 4.2.1.7-1 provides a description of the telemetry which is used during ground handling. Table 4.2.1.7-2 providesthe characteristics and location of each flight telemetry sensor. Finally, Figures 4.2.1.7-1 to 4.2.1.7-9 show thelocations of each sensor on the LV or ground transportation device.

4.2.1.7.1 Optional SC Telemetry through LV Telemetry System

There is a Customer option for monitoring up to 2 spacecraft parameters during flight by routing up to 2 spacecrafttelemetry points through the umbilicals to the Third and Fourth Stage telemetry system. Characteristics are describedbelow:

Item Characteristic

Data channel voltage range TBD

Data channel sample frequency 8000 Hz

Data channel measurement window From lift-off command to third stage separation

Allowable round trip impedance from umbilical connector tospacecraft TM point (ohms) to maintain measurement errorwithin +/- 7%

2 kohm (any combination of resistance, impedance andinductance)

Minimum isolation required between signal line and spacecraftstructural ground (Megohms)

TBD

Data recorded for these TM points will be provided to the Customer as part of the Post-Launch Report submissionand in electronic format.


Page 4-27

Table 4.2.1.7-1: Instrumentation Quantities and Locations for Ground Operations

AccelerationsAccelerometer Location andMeasurement Directions

AmplitudeMeasurementDynamicRange, G’s

FrequencyMeasurementRange, Hz

Transport by Motor Vehicle orRail, Spacecraft Mounted in

Location: In the area of the attachment of the containeron shock pallet to the Transport Vehicle

Shipping Container on ShockPallet

VerticalLateralLongitudinal

1.0 (TBX1)1.0 (TBY1)1.0 (TBZ1)

Up to 50 Hz

Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface RingMounted on LV Fourth Stage(Fourth Stage and FairingOnly)


1.0 (TBX3)1.0 (TBY3)1.0 (TBZ3)

Up to 50 Hz

Support Point of Fairing Assembly at Cylinder-NoseCone TransitionVerticalLateralLongitudinal

1.0 (TBX4)1.0 (TBY4)1.0 (TBZ4)

Up to 50 Hz

Spacecraft to SCA Separation Joint InterfaceVerticalLateralLongitudinal

1.0 (TBX)1.0 (TBY)1.0 (TBZ)

Up to 50 Hz

Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface RingMounted on Proton LaunchVehicle Assembly


1.0 (TBX6)1.0 (TBY6)1.0 (TBZ6)

Up to 50 Hz

Support Point of Proton First Stage at Aft RingVerticalLateralLongitudinal

1.0 (TBX5)1.0 (TBY5)1.0 (TBZ5)

Up to 50 Hz

Spacecraft to SCA Separation Joint InterfaceVerticalLateralLongitudinal

1.0 (TBX)1.0 (TBY)1.0 (TBZ)

Up to 50 Hz

Temperature Temperature Sensors Location Measurement Range, °C

All ground events 2 Thermocouples mounted on spacecraft support ring -10°C to +40°C (TA5, TA6)All ground events 2 Thermocouples measuring bulk air temperature inside

Khrunichev container-10°C to +40°C (TBK1, TBK2)

All ground events Temperature of inlet/exit air from thermal conditioningcar

0-50 °C

On Pad Temperature on AdapterTemperature of Air Supplied to Fairing

-50 to 80 °C (TA5, TA6)-50 to 80 °C (T20, T21)


Page 4-28

Table 4.2.1.7-1: Instrumentation Quantities and Locations for Ground Operations (Continued)

Humidity Humidity Sensors Location Measurement Range,

Transportation in Container Relative humidity inside shipping container 0-90%All transportation events Relative humidity of inlet, exit air from KhSC thermal

conditioning car0-80%

Contamination Contamination Sensors Location Measurement Range,

All ground events Particulate size at inlet/exit from air conditioning car (seefacility drwg)

0.5 microns/5 microns andhigher

All ground events Witness Plates (2) located inside fairing on boat-tail sectionOn Pad Access to fairing air supply for manual reading of

contamination levels0.5 microns/5 microns andhigher (SENSOR LABEL**)


Page 4-29

Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Typical)

Purpose Quantity, measurementrange

Sensor designation and location Time of operation Sensorsamplingfrequency

Temperature sensors

Environmentunder fairing

1 0...150C2 -40...100C2 -40...100C

T1 Nose internal surfaceT22 (T20) Cylinder, IV pl.T23 (T21) Cylinder, II pl.

From liftoff commandto nose fairing separation(Temperature sensors in

0.3 Hz

Ext. surfacethermalinsulation

1 -40...600C1 -40...500C1 -40...200C

T26 30 degree coneT25 20 degree coneT27 Cylinder

parenthesis are used onthe ground only).

Externalsurfaces of thecellularconstructionskin andthermalinsulation

1 -40...200C1 -40...200C2 -40...200C1 -40...200C1 -40...200C

1 -40...200C1 -40...200C2 -40...200C1 -40...200C1 -40...200C

1 -40...100C1 -40...100C2 -40...100C1 -40...100C1 -40...100C

1 -40...100C1 -40...100C

External skinT2 30 degree coneT3 20 degree coneT5,T4 Cylinder I pl.T6 Cylinder III pl.T7 Inverse cone I pl.Internal SkinT8 30 deg. cone I pl.T9 20 deg. cone I pl.T11,T10 Cylinder I pl.T12 20 deg. cone I pl.T13 Inverse cone I pl.Thermal insulationT16 30 deg. cone I pl.T28 20 deg. cone I pl.T17,T29 Cylinder I pl.T18 Cylinder III pl.T19 Inverse cone I pl.Panel liquid thermoregulating systemheater (LTS)T14 20 deg. cone I pl.T30,T15 Cylinder I pl.

Constructionof adapters

4 -10...77C TA1TA2 adapterTA3 upper frameTA4

From liftoff commandto spacecraft separation

1.6 Hz

2 -90...100C TA5TA6 adapter middle

To liftoff command(ground sensors)

continuous

4 -90...100C TA7TA8 adapter bottomTA9 frameTA10

From liftoff commandto spacecraft separation

1.6 Hz

Spacecraft separation sensors

Adapter

J1, J2

2

2

Separationplane

To spacecraftseparation

50-100 Hz


Page 4-30

Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Continued)



Nose fairing separation sensors

Nose fairing 24

4

1-25-67-81-23-4

To nose fairingseparation

50-100 Hz

Pressure sensors

Nose fairing 4 0...780mm Hg

1 20 degree cone,2 external surface34

From liftoff command tonose fairing separation

50 Hz

1: 0...400 mm Hg1: 0...780 mm Hg

5 Inverse cone,6 external surface

1: 0...780 mm Hg1: 0...250 mm Hg

1 20 degree cone,2 internal surface

1: 0...780 mm Hg1: 0...50 mm Hg

3 Inverse cone,4 internal surface

4: 0...780 mm Hg 1 Above vents6810

7: 0...780 mm Hg 2 Below vents3457911

Acoustic loads sensors

Nose fairinginternal andexternalacousticpressuremeasurement

2 places30-2000 Hz120-155 dB

2 places30-2000 Hz125-165 dB

AB-1 In the internalAB-2 volume between

SC and nosefairing. AB-1 on KhSC

adapter, AB-2 on fairingAH-1 On the noseAH-2 fairing external

From liftoff command tofirst stage separation

8000 Hz


Page 4-31

Table 4.2.1.7-2: Instrumentation Quantities and Locations for Flight Events (Continued)



Vibration loads sensors (low frequency)

Low frequencyvibrationmeasurementalong X, Y, andZ axes

3 places along X up to 32 Hzfrom -2...+4 g

2 places along Y & Z up to 32 Hz±0.6 g

KX-1 OnKX-2 adapter nearKX-3 SC interfaceKY-4KZ-5

KX-1KX-2KX-3KY-4KZ-5

From liftoff command tothird stage separation

At time of each stageseparation

200 Hz200 Hz200 Hz200 Hz200 Hz

400 Hz200 Hz200 Hz100 Hz100 Hz

Vibration loads sensors (high frequency)

Highfrequencyvibrationmeasurementalong X and Yaxes

2 places15...2000 Hzup to 8 gup to 8 g

On adapter

BX-CTBY-CT

From liftoff command tothird stage separation

8000 Hz


Page 4-32

Figure 4.2.1.7-1a: Locations of Measurement System Sensors on Block DM Fairing (Typical)


Page 4-33

Figure 4.2.1.7-1b: Locations of Measurement System Sensors on Breeze M Fairing (Typical)

TBS


Page 4-34

Figure 4.2.1.7-2: Locations of Measurement System Sensors on 1666 Adapter System (Typical)


Page 4-35

Figure 4.2.1.7-3: Locations of Measurement System Sensors on 1194 Adapter System (Typical)

TBS


Page 4-36

Figure 4.2.1.7-4: Instrumentation During Transportation of SC in Contractor’s Container

Rail Transportation of SC in SC contractor’s container from Yubeleini to SC Processing Facility (70 km at≤ 15 km/hr)

TBX1, TBY1, TBZ1

Conditioned Air InletConditioned Air ExhaustThermal Rail Car

Air Temperature Relative Humidity Particle Count

y

x

Figure 4.2.1.7-5: Instrumentation During Transportation of SC In KhSC Container

Rail Transportation of SC in KhSC container to/from Fueling Hall (0.5 km at ≤ 5 km/hr). (For fueling operations inArea 31 only)

TBX1, TBY1, TBZ1

Conditioned Air Inlet

Exhaust to Atmosphere (2)

Thermal Rail Car

Air TemperatureRelative HumidityParticle Count

T21

T20

(internal to containeropposite to T20)

(internal to containeropposite to T21)

y

x


Page 4-37

Figure 4.2.1.7-6: Instrumentation During Transportation of Ascent Unit

Rail Transportation of SC with upper stage from SC processing facility to area 95: 70 km at ≤ 15 km/hr.

TBX, TBY, TBZ

TBX1, TBY1, TBZ1

TBX3, TBY3, TBZ3

Conditioned Air InletConditioned Air ExhaustThermal Rail Car

Air Temperature (inlet, exit)Relative Humidity (inlet, exit)Particle Count (inlet, exit)

TA5, TA6

T20, T21

y

x

Figure 4.2.1.7-7: Instrumentation During Integration of Ascent Unit To LV

Temp., Humidity, Particle Count and Witness Plate measurements during the Ascent Unit mate to the LV.

TBX, TBY, TBZ Conditioned Air Inlet (2)

Air TemperatureRelative HumidityParticle Count

TA5, TA6

T20, T21

Witness Plates

Exhaust to Atmosphere

y

x


Page 4-38

Figure 4.2.1.7-8: Instrumentation During Transportation of Integrated Proton LV

Temp., Humidity, Particle Count and Witness Plate measurements during transport from Area 95 to the Launch Pad.

Con

ditio

ned

Air E

xhau

stTB

X, T

BY,

TBZ

Filte

r Blo

ck

TBX

1, T

BY1

, TB

Z1

TBX

3, T

BY3

, TB

Z3

Con

ditio

ned

Air

Inle

tW

itnes

s Pla

tes

TA5,

TA

6T20,

T21

Air

Tem

pera

ture

(inl

et, e

xit)

Rel

ativ

e Hum

idity

(inl

et, e

xit)

Part

icle

Cou

nt (i

nlet

,exi

t)


Page 4-39

Figure 4.2.1.7-9: Instrumentation During On-Pad Operations

Temperature, Humidity, Particle Count, and Witness Plate measurements while on the launch pad.

Fairing Air Exhaust to Atmosphere(two at 420 x 500)

Conditioned Air Inlet to Fairing(Air sample access: continuousmonitoring of air used to measure temp., humidity, and contamination levels)

Witness Plates(access door on Fairing boat tail)

AI R TH ERM AL C ONTR OL SYST EM (ATC S) TemperatureHumidityParticle Count

Temperature insideFairing, T20 and T21 (see table for description,see fairing schematicfor exact location).

Temperature on Adapter,TA5 and TA6 (see table for description,see fairing schematicfor exact location).


Page 4-40

4.2.2 Launch Pad EGSE Interfaces

The two interface connectors described in Section 4.2.1 are wired to a Mission Specific wiring harness on the adapterwhich is connected to the LV flight umbilical harness running the length of the vehicle to an interface connector Ø06 atthe bottom of the first stage. From here, ground cabling connects the umbilical to an interface panel in the Vault underthe launch pad where the Customer electrical interface equipment is located. As can be seen from Figure 4.2.2-1, thereare test access connectors on the Fourth Stage which permit access to the umbilical from the Mobile Service Tower upto 8 hours prior to launch. These can be used to interface Customer battery charging power supplies on the MobileService Tower with the SC. They can also be used to connect with wiring in the Mobile Service Tower to provide aparallel path with the flight LV umbilical to reduce overall resistance drop from the SC to Customer GSE for highcurrent power lines.

The Launch Pad interfaces include connections from the base of the Proton LV (and connections at station 43.85 onthe Mobile Service Tower if required) to ground wiring interfacing with SC EGSE. ILS provides all necessaryelectrical harnesses and cables between the SC/LV IFD’s and the SC EGSE interface enables in the Vault and on theMST. Figure 4.2.2-1 provides a block diagram of the electrical interfaces available between the payload, LaunchVehicle and the ground systems.

Figure 4.2.2-1: Spacecraft to Launch Vehicle and Ground Systems Electrical Interfaces

Mobile ServiceTower

Vault

CustomerMPS

Bunker Rm 250

CustomerSTE

Spacecraft

Level 6

O 06

O 1A

37 or 61 pin Deutsch Connectors

JunctionBox X1

Mission Specific Wiring

Mission SpecificWiring

Notes: 1) Cable length from level 6 of MST to Vault: 400 meters2) Cable length from SC connectors to O 06: 50 meters3) Cable length from O 06 to Vault: TBD meters


Page 4-41

4.2.2.1 EGSE Interface Electrical Constraints

The maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.

SC contractor equipment needs to provide protection against exceeding 100 V at spacecraft umbilical interface. It alsoneeds to provide continuous monitoring and recording of this bus voltage.

All EGSE interfaces to the SC shall be current limited by the Customer to preclude damage to the LV ground andairborne systems in the event of a short circuit . SC STE shall shut off power within 0.2 sec if Imax is exceeded by 50%.

SC and GSE will be de-energized prior to mating and demating of umbilical connectors for electrical checkouts andflight mating.

All SC EGSE electrical interface circuits shall be constrained at least 5 minutes prior to liftoff such that there will beno current flow greater than 100 mA per wire across the T-0 interface .

4.2.3 Telemetry/Command Links

An RF link for telemetry/command is provided between the SC on the pad and the Bunker. The Customer can choosebetween 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 tobunker RF/electrical interface is shown in Figure 4.2.3-1.

In order to confirm compatibility with the link, the following is required of the Customer:

a) The SC Checkout Station shall have 1 physical command interface and 1 physical telemetry interface.

b) SC GSE RF interface impedance shall be 50 OHM

c) Uninterrupted operation of RF devices shall not exceed 8 hours, with a 30 minute break before the next 8 hoursession.

d) The SC contractor shall provide to Khrunichev the measured coefficient values for TT &C signals via the RFwindow obtained during the RF channel checkup in the integration facility following the Ascent Unitencapsulation.

e) Prior to installation of the LV+Ascent Unit on the pad and following the delivery of the STE to the bunker, theSC manufacturer shall verify continuity between command RF link and STE and issue to Khrunichev theCertificate of Launch Pad Readiness to accommodate the LV and Ascent Unit.

f) After installation of LV+Ascent Unit and prior to the roll-up of service tower, the SC manufacturer, inconjunction with Khrunichev, shall check out the RF link between the Ascent Unit and STE. Such check out shallbe performed 20 minutes after the mating of the LV aft section. SC contractor to confirm functionality of the RFlink within 45 minutes.

g) At L-6 months, the SC manufacturer shall provide to Khrunichev two connectors for installation by Khrunichevon the existing RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructionson cable dressing and cable performances.


Page 4-42

Table 4.2.3-1a: C-Band RF Link Characteristics

Telemetry Link

Reference Value Note

Frequency range, GHz 3.95 ± 0.2Bandwidth, MHz > 250Signal polarization Left-hand circularRadio link output signal powermaximum. dBm minus 0.0 with Service Tower rolled back

minus 6.2 with Service Tower in placeminimum, dBm minus 30.0 with Service Tower rolled back

minus 36.2 with Service Tower in placeRadio link gain factor, dB minus 31.0 ±5

minus 42.2 ± 2with Service Tower rolled backwith Service Tower in place

Gain factor adjustment limit forradio link input, dB

30

Radio link output signal-to-noiseratio, dB*Hz

64

Command Link


Frequency range, GHz 6.42 ± 0.05Bandwidth, MHz > 200Signal polarization Right-hand circularRadio link output signal powermaximum, dBW/m2 minus 22.2 with Service Tower rolled back

minus 38.6 with Service Tower in placeminimum, dBW/m2 minus 72.2 with Service Tower rolled back

minus 88.6 with Service Tower in placeRadio link gain factor, dB minus 43.2

minus 38.6with Service Tower rolled backwith Service Tower in place

Gain factor adjustment limit for radiolink input, dB

30

Radio link output signal-to-noise ratio,dB*Hz

70


Page 4-43

Table 4.2.3-1b: Ku-Band RF Link 1 Characteristics

Telemetry Link


Frequency range, GHz 11.1 ± 0.2Bandwidth, MHz > 250Signal polarization Linear verticalRadio link output signal power with SC antenna input signal power of 0 dBWmaximum, dBm minus 31 with Service Tower rolled back

minus 37 with Service Tower in placeminimum, dBm minus 41 with Service Tower rolled back

minus 41 with Service Tower in placeRadio link gain factor, dB minus 78.3 ± 5

minus 80.0 ± 2with Service Tower rolled backwith Service Tower in place


30


118

Command Link


Frequency range, GHz 14.0 ± 0.05Bandwidth, MHz > 200Signal polarization Linear, horizontalRadio link output signal power with SCS antenna input signal power of

3 dBWmaximum, dBW/m2 minus 55.5 with Service Tower rolled back




Gain factor adjustment limits for radiolink input, dB

from -65 to -41from -67 to -43


123


Page 4-44

Table 4.2.3-1c: K-Band RF Link 2 Characteristics

Telemetry Link


Frequency range, GHz 12.2 ± 0.2Bandwidth, MHz > 250Signal polarization left-hand circularRadio link output signal power with SC antenna input signal power of 0 dBWmaximum, dBm 2 with Service Tower rolled back

minus 3 with Service Tower in placeminimum, dBm minus 8 with Service Tower rolled back

minus 7 with Service Tower in placeRadio link gain factor, dB minus 35 ± 5

minus 45 ± 2with Service Tower rolled backwith Service Tower in place


30


118

Command Link


Frequency range, GHz 14.0 ± 0.05Bandwidth, MHz > 200Signal polarization Right-hand circularRadio link output signal power with SCS antenna input signal power of 3 dBWmaximum, dBW/m2 minus 31 with Service Tower rolled back

minus 36 with Service Tower in placeminimum, dBW/m2 minus 41 with Service Tower rolled back

minus 40 with Service Tower in placeRadio link gain factor, dB minus 78.3


Gain factor adjustment limits for radiolink input, dB

from -50 to -80from -52 to -82


123


Page 4-45

Table 4.2.3-1d: Ku-Band RF Link 3 Characteristics

Telemetry Link


Frequency range, GHz 12.45 ± 0.25Bandwidth, MHz > 500Signal polarization Linear horizontalRadio link output signal powermaximum, dBm minus 8.0 with Service Tower rolled back

minus 14.4 with Service Tower in placeminimum, dBm minus 44.0 with Service Tower rolled back




30


118

Command Link


Frequency range, GHz 17.3 ± 0.05Bandwidth, MHz > 200Signal polarization Linear verticalRadio link output signal powermaximum, dBW/m2 minus 59.0 with Service Tower rolled back




Gain factor adjustment limit for radio linkinput, dB

30


155


Page 4-46

Figure 4.2.3-1: SC to Bunker RF/Electrical Interface Block Diagram

A048

RF link

BunkerUnderground Vault

InterconnectPanel

InterconnectPanel

InterconnectPanel

RF TMRF TC

KhrunichevRF

Rack

Launch Pad

SC

4.2.3.1 RF Window Requirements

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.

There are 2 RF window positions in the fairing to take into account the possible view angles required at each of the 2Proton launch pads. When the launch pad is designated, 1 out of the 2 windows will be covered with conductive enamelleaving 1 active window for transmission of the spacecraft T and C signal between the spacecraft and the Bunker. TheRF link between the spacecraft and the ground RF equipment is required before and after Mobile Service Towerrollback. Prior to rollback, the signals will be transmitted via a repeater on the Mobile Service Tower. Followingrollback, signals will be transmitted directly between the spacecraft and the Bunker antenna. During rollback, therecould be a maximum 20 minute outage of signal as the tower rolls through the line of site between the spacecraft andthe bunker antennas.


Page 4-47

4.2.4 Electrical Grounding

All payload preparation areas used by the SC as a launch base facilities used by SC and SC EGSE are equipped withearth referenced steel ground busses with equipment attach points (threaded studs). The resistance between any pointon these bars and the building earth ground is less than 4 ohms. The floor surfaces in the payload and hazardouspayload processing areas is anti-static and connected to the facility grounding system. The SC contractor shall provideall cables and attachment hardware required to interconnect the SC and support equipment with facility grounds.

4.2.5 Electrical Bonding

The resistance across the spacecraft/adapter separation plane shall be less than 10 milliohms at a current less than 10milliamps. This may be accomplished either by conductive surface contact between the spacecraft and adapterinterface ring (1666 adapters) or by the use of 2 bonding straps which incorporate a friction contact connector thatreleases upon spacecraft separation with a separation force of 40 ± 5 N. Signal and power grounds from the spacecraftare passed through the umbilical without connecting them to Launch Vehicle structure. Likewise umbilical shieldgrounds are isolated from the Launch Vehicle structure.

4.2.6 SC/LV Lightning Protection

All payload preparation areas used by the SC will be equipped with a lightning protection system for direct andindirect hits. Augmentation of the standard provisions for any necessary SC individual circuit protection shall beprovided by the SC contractor.

4.2.7 Static Discharge

During the entire flight through SC separation, no electrostatic discharge shall occur from either the LV or the SCsurface through the LV to SC interface plane.

4.3 FITCHECK OF MECHANICAL/ELECTRICAL INTERFACES

A fitcheck of electrical/mechanical interfaces with the flight adapter and spacecraft is required at the SCmanufacturer’s facility for first of a kind spacecraft and the first follow-on spacecraft in a series.


Page 5-1

5. MISSION INTEGRATION AND MANAGEMENT

ILS provides the Customer a SOW which define the management approach for a Customer Launch ServiceAgreement, the deliverables provided to the Customer during the course of the LSA and a schedule for all MissionIntegration activities. This section highlights these provisions.

5.1 MANAGEMENT PROVISIONS

5.1.1 Key Personnel

Immediately after execution of each LSA, ILS and the Customer will designate their respective Mission Managerswho will be responsible for performing all management functions related to the LSA.

ILS shall ensure that personnel necessary for the performance of this contract are made available to the program toperform the work in a timely fashion and to satisfy requirements of the contract and its exhibits.

5.1.2 Interface Control Document (ICD)

The ICD will be created by ILS based on a generic ICD template and the Customer provided IRD. It will provide theCustomer's technical requirements for the launch of their spacecraft and characteristics and constraints of the LaunchVehicle and Launch Site relating to the interface with the spacecraft.

5.1.3 Schedule Monitoring

ILS will create and maintain an interface activities milestone schedule that provides all key technical interfacemilestones necessary for successful completion of the contract. A typical mission integration schedule is as shown inFigure 5.1.3-1a and 5.3.1-1b for a non-recurring and recurring program respectively. It is based on the meetingschedule in Section 5.1.5.1 and the deliverable milestones provided in Sections 6 and 7. This integrated programschedule for a particular program will be presented and agreed upon between all Parties at the Kickoff Meeting andfurther changes shall be made as necessary and agreed upon at subsequent Technical Interface Meetings (TIM). In caseof changes to internal schedules, the other party will be promptly informed.


Page 5-2

Figure 5.1.3-1a: Baseline Integration Schedule (Non-Recurring Program)

MEETINGS Kickoff

ICD ReviewPDR

Site Visit No.1 Site Visit No.2Operations Review #1 Operations Review #2

ICD Sign-OffFitcheckShock TestCDR

LV PreshipSC PreshipSite Review

LVQR UpdateLV Roll-out AuthLV Loading Auth

ICD DEVELOPEMENTIRD (Amendment for specific SC)

Prelim ICDSigned ICD

PRELIMINARY DESIGNPreliminary SC Dynamic Model

CLA Report (if required)SC Thermal Model

PDR PackageCRITICAL DESIGN

Inputs to Final Analysis (except CLA)Final SC Dynamic Model

CLA ReportCDR Package

SPACECRAFT TESTINGSC Environmental Test Plan

Inputs to Fitcheck & Shock Test PlanFitcheck & Shock Test Plan

Fitcheck & Shock ReportSC Environmental Test Results

SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSINGSafety Submissions

All CertificatesLAUNCH CAMPAIGN

SC Launch Ops Plan (Prelim)SC Launch Ops Plan (final)

Participants List w/ Passport InfoSC Orbital Data

Launch Eval ReportMANAGEMENT

123456789

101112131415

161718

19202122

23242526

2728293031

3233

3435363738

39

KickoffPreliminary ICD ReviewPreliminary Design Review (PDR)Launch Site VisitsOperations ReviewICD Sign-OffFitcheckShock TestCritical Design Review (CDR)Launch Vehicle Preshipment RevSpacecraft Preshipment ReviewLaunch Site Acceptance ReviewLaunch Vehicle Quality ReviewUpdateLV Roll-out Authorization ReviewBoardLV Loading Authorization ReviewBoardIRD (Amendment for specific SC)Preliminary ICDICD prior to meeting to sign

Preliminary SC Dynamic ModelAnnex [x] of PDR package, CLAReportSC Thermal ModelPDR Package, all except Annex [x]

Inputs to Final Analysis (exceptCLA)Final SC Dynamic Model (testverified)Annex [x] of PDR package; CLAReportCDR Package; all except Annex [x]

SC Environmental Test PlanInputs to Fitcheck & Shock Test PlanFitcheck & Shock Test PlanFitcheck & Shock Test ReportSC Environmental Test Results

Safety SubmissionsAll Certificates

SC Launch Operations Plan(Preliminary)SC Launch Operations Plan (Final)Participants List w/ Passport InfoSC Orbital DataLaunch Evaluation Report

Management Report

AllAllAllAllAllAllAllAllAllAllAllAllAllAllAll

SCCILSILS

SCCILS

SCCILS

SCCSCCILSILS

SCCSCCILSILS

SCC

SCCAll

SCCSCCSCCILSILS

ILS

L-24 L-21 L-18 L-15 L-12 L-9 L-6 L-3 LAUNCH L+2

Revised:

Printed:

ID # Activity Resp

STANDARD INTEGRATION SCHEDULE-NON-RECURRING

31AUG98

08/21/98


Page 5-3

Figure 5.1.3-1b: Baseline Integration Schedule (Recurring Program)

MEETINGS

Kickoff

ICD Sign-Off

CDR

Fitcheck

Shock Test

LV Preship Rvw

SC Preship Readiness Rvw

Ops Review

Site Review

LVQR Update

LV Roll-out Auth

LV Loading Auth

ICD DEVELOPEMENT

IRD

ICD

CRITICAL DESIGN

Inputs to Final Analysis

Final SC Dynamic Model

CLA Report

CDR Package

SPACECRAFT TESTING

SC Environmental Test Plan

Inputs to Fitcheck & Shock Test Plan

Fitcheck & Shock Test Plan

Fitcheck & Shock Report

SC Environmental Test Results

SAFETY DATA AND CERTIFICATES REQUIRED FOR LICENSING

Safety Submissions

All Certificates

LAUNCH CAMPAIGN

SC Launch Ops Plan (final)

Participants & Info List

Orbital Data

Launch Eval Report

MANAGEMENT

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

30

Kickoff

ICD Sign-Off

Critical Design Review (CDR)

Fitcheck

Shock Test

LV Preshipment (Quality) Review

SC Preshipment Readiness Review

Operations Review

Launch Site Acceptance Review

Launch Vehicle Quality ReviewUpdateLV Roll-out Authorization ReviewBoardLV Loading Authorization ReviewBoard

IRD (Amendment for specific SC)

ICD prior to meeting to signrevision

Inputs to Final Analysis (exceptCLA)Final SC Dynamic Model (TestVerified)Annex [X] of CDR package; CLAReportCDR Package; all except Annex [x]

SC Environmental Test Plan

Inputs to Fitcheck & Shock Test Plan

Fitcheck & Shock Test Plan

Fitcheck & Shock Test Report

SC Environmental Test Results

Safety Submissions

All Certificates

SC Launch Operations Plan (final)

Participants List w/ Passport Info

Orbital Data

Launch Evaluation Report

Management Report

All

All

All

All

All

All

All

All

All

All

All

All

SCC

ILS

SCC

SCC

ILS

ILS

SCC

SCC

ILS

All

SCC

SCC

ILS

SCC

SCC

ILS

ILS

ILS

L-12 L-10 L-8 L-6 L-4 L-2 LAUNCH L+2

Revised:

Printed:

ID # Activity Resp

STANDARD INTEGRATION SCHEDULE - RECURRING

31AUG98

08/21/98


Page 5-4

5.1.4 Documentation Control and Delivery

ILS maintains an internal documentation and configuration control system for all LSAs. Deliverable documentationshall be maintained under this configuration control system.

All technical correspondence between ILS and the Customer relating to work on the LSA is strictly between theCustomer Mission Manager and the ILS Mission Manager.

5.1.5 Meetings and Reviews

ILS and the Customer will meet as often as necessary to allow good and timely execution of all activities related tolaunch preparation of each satellite. A preliminary meeting schedule is defined in Section 5.1.5.1 and meetingschedules will be updated through the course of the contract as part of the interface activities milestone schedulegenerated by ILS. Exact dates, locations, agendas, and participation are agreed upon in advance, on a case-by-casebasis, by the ILS Mission Manager and the Customer Mission Manager.

5.1.5.1 Interface Meetings and Reviews

The ILS Mission Manager chairs all meetings unless otherwise specified. ILS will provide meeting minutes at the endof each meeting signed by ILS and the Customer.

A baseline meeting schedule is provided in Tables 5.1.5.1-1a and 5.1.5.1-1b for a non-recurring and recurringprogram respectively. A non-recurring program is one with a first of a kind SC which requires 2 analysis cycles. Arecurring program is one with a similar SC which requires only one analysis cycle and no significant changes to thelaunch vehicle and the launch site.


Page 5-5

A description of each type of meeting is provided below:

Table 5.1.5.1-1a: Baseline Meeting Schedule for Non-Recurring Program

Meeting Date* Location

Kickoff L-24 TBD

Preliminary ICD Review L-20 TBD

Preliminary Design Review (PDR) L-14 Moscow

Launch Site Visit No. 1 L-12 Launch Site

Operations Review No. 1 L-12 TBD

ICD Signoff L-11 TBD

Fitcheck L-6 SC Manufacturer

Shock Test L-6 SC Manufacturer

Critical Design Review (CDR) L-6 Moscow

Launch Site Visit No. 2 L-5 Launch Site

Launch Vehicle Preshipment Review L-3 Moscow

Spacecraft Preshipment Review L-2 SC Manufacturer

Operations Review No. 2 L-2 TBD

Launch Site Acceptance Review L-2 Launch Site

Launch Vehicle Quality Review Update (TBC) L-7 days Launch Site

Launch Vehicle Rollout Authorization Review Board L-6 days Launch Site

Launch Vehicle Fueling Authorization Review Board L-1 day Launch Site

* Date: Launch Minus X Months


Page 5-6

Table 5.1.5.1-1b: Baseline Meeting Schedule for Recurring Program

Meeting Date* Location

Kickoff L-12 TBD

ICD Signoff L-10 TBD

Critical Design Review (CDR) L-6 Moscow

Fitcheck L-6 SC Manufacturer

Shock Test L-6 SC Manufacturer

Launch Vehicle Preshipment (Quality)Review L-3 Moscow

Spacecraft Preshipment Readiness Review L-2 SC Manufacturer

Operations Review L-2 TBD

Launch Site Acceptance Review L-2 Launch Site

Launch Vehicle Quality Review Update (TBC) L-7 days Launch Site

Launch Vehicle Rollout Authorization Review Board L-6 days Launch Site

Launch Vehicle Fueling Authorization Review Board L-1 day Launch Site

* Date: Launch Minus X Months

Kickoff

This meeting represents the formal start of the program. A description of overall LSA services will be presented as wellas management organization and preliminary program schedules. The Interface Requirements Document will bediscussed to kickoff generation of the ICD.

Preliminary ICD Review

The preliminary ICD will be reviewed and agreement reached on inputs to begin the Preliminary Analysis cycle.

Preliminary Design Review (PDR)

ILS will present all results of preliminary analyses and compare with ICD requirements.

Launch Site Visit No. 1

This first visit to the launch site will provide a first orientation to the Customer. A key goal is to verify compliance withICD requirements.

Operations Review

Review of requirements and corresponding implemention for launch base operations.


Page 5-7

Review of inputs to Final Analyses

Agreement will be reached at this meeting to all final analysis inputs prior to starting these analyses.

Fitcheck

This is a fitcheck of flight adapter and separation system hardware to the flight spacecraft at the SC manufacturer'sfacility.

Shock Test

This is an actuation of the flight type separation system with the flight spacecraft at the manufacturer's facility. It isdone in conjunction with the fitcheck.

Critical Design Review (CDR)

ILS presents all results from the final analysis cycle.

Launch Site Visit No. 2

This is the last visit prior to certification of the launch facility for this mission.

Launch Vehicle Preshipment Review

This meeting is held at Khrunichev as part of the quality control process. Khrunichev presents the quality status of allLaunch Vehicle hardware per design documentation.

Spacecraft Preshipment Review

This meeting is held at the SC manufacturer's facility and provides a status of the spacecraft readiness to ship to theLaunch Site.

Launch Site Acceptance Review

This meeting is held at the Launch Site prior to SC arrival to confirm the readiness of the Launch Site to begin thelaunch campaign. The compliance with requirements in the ICD and Operations Plan will be verified.

Launch Vehicle Quality Review Update

At this meeting, the data presented by Khrunichev at the Launch Vehicle Preshipment Review is updated to take intoaccount all subsequent activities up to L-7 days before launch (TBC).


Page 5-8

LV Rollout Authorization Review Board

A meeting is held at the Launch Site to confirm readiness to rollout the LV to the Launch Pad.

LV Fueling Authorization Review Board

A meeting is held at the Launch Site to confirm readiness to load the LV with propellants and confirm SC readiness tolaunch.

Action Item Control

ILS maintains a centralized action item control system for each LSA.

5.1.6 DTRA Oversight

ILS arranges for Defense Threat Reduction Agency (DTRA) oversight, as necessary for technical interchangeinvolving foreign nationals.

5.1.7 Quarterly Report

At the end of each contract quarter, ILS provides a report for each LSS covering management and technical progressincluding: subcontract status, technical status, mission integration schedule, production schedule and status for allmajor hardware, action item status, contract deliverable status. Major program issues are summarized and resolutionplans discussed.

5.1.8 Quality Provisions

Refer to Appendix B for a description of Quality Assurance provision in place for Proton launch services.

5.1.9 Launch License And Permits

ILS will obtain all necessary Russian Federation permits and approvals required for the processing and launch of theCustomer's spacecraft.

The Customer will obtain permits and approvals required to import and export the spacecraft and associatedequipment from its country of origin through the Port of Entry in Russia and Kazakhstan.

5.2 ILS DELIVERABLES

ILS provides the following deliverables during the course of each LSA. The baseline delivery schedule is provided inTable 5.2-1.


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Table 5.2-1: ILS Deliverable Schedule for a Recurring and a Non-Recurring Program

Recurring Non-Recurring

DOCUMENT Date* Date*

ICD DEVELOPMENT

Preliminary ICD L-21

ICD prior to sign off meeting L-11 L-12

PRELIMINARY DESIGN

PDR package PDR -10 days

CRITICAL DESIGN

CDR package CDR -10 days CDR -10 days

SPACECRAFT TESTING

Fitcheck and Shock Test Plan N/A L-9

Fitcheck and Shock Test Report N/A L-5


Safety Certificate L-3 L-3

LAUNCH CAMPAIGN AND LAUNCH

Orbital Data L+0 L+0

Launch Evaluation Report L+2 L+2

MANAGEMENT

Management Report Each quarter Each quarter

5.2.1 ICD Development

5.2.1.1 ICD

ILS shall provide the ICD and will maintain it up to date by issuing revisions as necessary. Preliminary ICD willcontain input data for the Preliminary Design. The signed ICD will contain input data for the Critical Design.

5.2.2 Preliminary and Critical Design

ILS will conduct all performance and mission analyses required for the proper implementation of the Customer’slaunch mission, including those analyses identified below.

The following analyses are conducted during the mission integration effort for each satellite launch mission. For firstof a kind spacecraft, one preliminary and one final analysis cycle will normally be conducted during each satelliteintegration effort. For follow-on spacecraft, one analysis cycle will be performed in most cases. Two analysis cycleswill be performed for following on SC where spacecraft and/or launch vehicle relevant parameters have changedsignificantly.


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Each cycle includes the analyses defined in Table 5.2.2-1:

Table 5.2.2-1: Design Review Analyses

Section Title Description

Design and Manufacturing A summary of the LV design concentrating on differences with previousvehicles. Emphasis is on specificities in adapter and payload compartmentdesign to meet specific spacecraft payload requirements.

Mission Design The flight design including maneuvers and maneuver sequence, orbitparameters and dispersions, collision avoidance

Thermal Analysis Integrated thermal analysis of combined operations (ground and flight) forspacecraft and launch vehicle hardware to ensure thermal compatibility.The spacecraft mathematical model is provided by the Customer per thethermal model specification provided by ILS

Separation Analysis Analysis of spacecraft separation including presentation of pertinentkinematic parameters and their dispersions during the separation event

CLA/Acoustic/ Shock LoadsEnvironment

1) Dynamic coupled-loads analysis. The spacecraft mathematical model isfurnished by the Customer according to the ILS provided dynamic modelspecification. The following events are analyzed:

a) Liftoff

b) Flight Winds and Gust

c) First/Second Stage Separation

For follow-on satellites of the same configuration, only one verificationcoupled loads analysis will be conducted unless significant launch vehicleconfiguration changes have occurred (e.g. if first launch occurs on BlockDM and next launch occurs with Breeze M).

2) Presentation of other load environments including acoustic, shock andground transportation loads

Contamination Analysis of ground and flight contamination sources and effect onspacecraft payload

RF Link and EMC Analysis of the RF link between the bunker and the pad and EMC analysisverifying compatibility between the spacecraft and LV systems

Clearance Analysis Clearance analysis between the spacecraft and the LV during flight to verifysufficient dynamic clearances

Venting Analysis Analysis of fairing depressurization during flight

Operations Detailed description of how Khrunichev will meet operationalrequirements specified by the Customer in the ICD


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One hardcopy (and a soft copy where possible) of all design documentation will be provided to Customer 10 days priorto review

Reports will be provided documenting the results of the above analyses. These reports will be provided for each analysiscycle and include the following topics: summary of results, detail of analyses performed, comparison of analysis resultswith ICD requirements. The analyses required may be reduced in scope if agreed between ILS and the Customer.

5.2.3 Spacecraft Testing - Fitcheck/ Shock Test Plan, and Report

ILS will provide an overall plan describing the Fitcheck/Shocktest and a description of the responsibilities and actionsfor each of the participants including Khrunichev, ILS, the SC manufacturer and SAAB (if applicable). ILS willprovide a summary report following the Fitcheck & Shocktest documenting the results.

5.2.4 Safety

ILS will prepare a Safety Data Package based on the safety data provided by the Customer Safety Submissions.

5.2.5 Launch Campaign and Launch

5.2.5.1 Orbital Data

ILS will provide the State Vector data as described in Chapter 2.

5.2.5.2 Launch Evaluation Report

ILS will provide a Launch Evaluation Report for each LSA documenting the results of ground processing of thespacecraft and the subsequent flight.

5.2.6 Management and Reports

For each LSA, ILS will provide a report each quarter documenting the status of management issues.

5.3 CUSTOMER DELIVERABLES

The Customer provides the following deliverables during the course of each LSA. The baseline delivery schedule isprovided in Tables 5.3-1.


Page 5-12

Table 5.3-1: Customer Deliverable Schedule for a Recurring and a Non-Recurring Program

Recurring Non-Recurring

DOCUMENT Date* Date*

ICD DEVELOPMENT

IRD L-12 L-24

PRELIMINARY DESIGN

Preliminary SC Dynamic Model L-24

SC Thermal Model L-19

CRITICAL DESIGN

Final SC Dynamic Model (test verified model) L-7 L-7

Inputs to Final Analyses (except SC Dynamic Model) L-12 L-11

SPACECRAFT TESTING

SC Environmental Test Plan L-11 L-11

Notching Profile for Finite Element Analysis Sine-test -2 months

SC Environmental Test Results L-4 L-4

Inputs to Fitcheck & Shock Test Plan L-10 L-10

Fitcheck & Shock Test Report L-6 L-5


Safety Submissions (Preliminary and Final SMPSP) L-12 to L-5 L-24 and L-5

Pre-Launch safety certificates L-3 L-3

LAUNCH CAMPAIGN AND LAUNCH

Spacecraft Launch Operations Plan (preliminary) L-16

Spacecraft Launch Operations Plan (final) L-8 L-8

Listing of Campaign Participants with Passport Information L-3 L-3

SC Orbital data L+2 days L+2 days


Page 5-13

5.3.1 ICD Development

The Customer will provide an IRD to ILS with interface requirements describing all pertinent design information onthe spacecraft characteristics, mechanical and electrical interfaces, and constraints necessary to define the integrationtasks and mission operation. This will be used to generate the preliminary ICD.

5.3.2 Preliminary and Critical Design

5.3.2.1 SC Dynamic Model

The Customer will provide a SC dynamic model conforming to the ILS dynamic model specification. For the secondanalysis cycle, this model will be a test verified version.

5.3.2.2 SC Thermal Model

The Customer will provide a SC thermal model conforming to the ILS thermal model specification.

5.3.2.3 SC Fluid Slosh Model

The Customer shall provide a SC fluid slosh model conforming to the requirements in Appendix C.

5.3.3 Spacecraft Testing

5.3.3.1 SC Environmental Test Plan and Results

The Customer will provide a test plan for ILS approval documenting the tests which will be performed by the SCmanufacturer to demonstrate compatibility with the Proton ground and flight environments. A summary of the resultsfrom these tests will be provided at test completion.

5.3.3.2 Fitcheck/ Shock Test Plan, Procedures and Report

The Customer will provide input to the ILS Fitcheck/Shock Test Plan. The Customer will provide a summary reportfollowing the Fitcheck & Shocktest documenting the results.

5.3.4 Required Safety Data and Certificates

5.3.4.1 Safety Submissions

The Customer will provide all necessary data required by the Safety Plan to demonstrate that the spacecraft systems,GSE and procedures are compatible with the Launch Site and flight safety requirements.

5.3.4.2 SC Reliability Certificate

The Customer will provide a certificate to confirm reliability and specifying the SC no-failure probability level forground operations and flight.


Page 5-14

5.3.4.3 SC Readiness Certificate

The Customer will provide a certificate to ILS following spacecraft environmental testing to certify that the spacecraftis capable of withstanding the loads environments imposed on it during ground processing at the Launch Site andduring flight as documented in the ICD and in the CLA report.

5.3.4.4 Certification of Orbital Position Approval

The Customer will provide to ILS a certificate demonstrating that approval has been given by an internationallyrecognized authority for the orbital position which the spacecraft to be launched under this LSA is to occupy.

5.3.4.5 Certificate Approving SC for Space Related Activities (TBC)

The customer will provide a certificate stating that the SC customer is authorized to conduct space related activities.

5.3.4.6 Certificate of Intent to File National Registry(TBC)

The Customer will provide a certificate stating intent to enter SC into the national registry of the owner's country.

5.3.5 Launch Campaign and Launch

5.3.5.1 Spacecraft Launch Operations Plan

The Customer will provide a plan which describes the spacecraft launch operations at the Launch Site.

5.3.5.2 Listing of Campaign Participants

The Customer will provide a list of all potential campaign participants 3 months prior to launch with all requiredpassport information. This list will designate primary and backup personnel.

5.3.5.3 Orbital Data

The Customer will provide SC state vector data complying with requirements in Chapter 2 of this Mission Planner’sGuide.

5.4 SPECIFIC CUSTOMER RESPONSIBILITIES

For each LSA, the Customer has the following responsibilities.

5.4.1 Campaign Duration

The campaign duration from spacecraft arrival to launch shall not exceed 42 days.

5.4.2 Spacecraft And Associated Ground Equipment

The Customer will provide at the Launch Site the spacecraft and associated ground equipment and personnel requiredto meet the contracted launch date.


Page 5-15

5.4.3 Final Spacecraft Data

For missions that employ the Block DM, the Customer shall also provide final spacecraft estimated dry and wet masstwo days prior to Block DM fueling. Prior to encapsulation, the Customer will supply ILS the actual satellite dry andwet masses.

5.4.4 Spacecraft Readiness

The Customer will provide a readiness to proceed with operations on launch minus 17 days (L-17 day), prior to thestart of Combined Operations; prior to rollout to the pad; and prior to fueling of the Launch Vehicle. These dates willbe coordinated with the Baikonur operations schedule.

5.4.5 Removal of SC Support Equipment

Unless prior arrangements have been made, the Customer shall remove from the Baikonur Cosmodrome all of itsground support equipment using Customer provided charter aircraft within 3 days of launch.

5.4.6 Evaluation Of Launch Vehicle And Associated Services

The Customer will provide to ILS, as soon as practical after launch, with all relevant available data from the launchnecessary to assist ILS in evaluating the performance of the launch vehicle and associated services provided under eachLSA.

5.4.7 Spacecraft Propellants

The Customer will procure spacecraft propellants to support the launch campaign and is responsible for shipment ofthese propellants to the Port of Entry into Russia (typically St. Petersberg) and through Customs. After the launchcampaign, the Customer will be responsible for removal of the propellants and associated equipment from the Port ofEntry. ILS will assist the Customer with Customs clearance procedures.

The Customer is responsible for propellant transportation costs.

5.4.8 Connectors

The Customer shall provide to ILS, flight and test connectors per mission specific requirements. These connectors willbe used for the assembly of flight and test harnesses.


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5.5 ILS SERVICES AND MATERIAL SPECIFICALLY EXCLUDED

ILS has no obligation to provide the following goods or services:

a) Receiving inspection of spacecraft elements and support equipment upon arrival at the launch site

b) Analysis of data generated by the spacecraft through its own telemetry system

c) Any spacecraft test equipment

d) Shipping cost associated with the spacecraft, its components, and support equipment (except while at the launchsite)

e) Replacement parts for the spacecraft or its support equipment

f) Installation, handling, or other responsibility related to spacecraft pyrotechnic systems or elements

g) Functional operation or installation of any spacecraft systems

h) Any tracking or commanding of the spacecraft after separation from the launch vehicle

i) Spacecraft propellants

j) Propellant sampling analysis. Facilities at or near the Launch Site are not equipped with equipment or technologynecessary for analysis of SC propellants. The Customer must plan for shipment and analysis of samples outside theRussian Federation if these analyses are required

k) Storage of spacecraft and all associated Customer GSE over 2 months

l) Storage of propellants over 4 months

m) Additional analyses over and above those specified in previous sections, caused by changes to the spacecraft designwhich are not in any way attributable to ILS and not required by the terms of this LSA and its' Exhibits

n) Additional analyses over and above those required in previous sections, caused by launch postponements requestedby the Customer (unless otherwise specified in the postponement provisions of the LSA)TBC

o) Changes to the LV and/or the launch site facilities as described in this Mission Planner’s Guide , caused bychanges to the spacecraft design which are not in any way attributable to ILS and not required by the terms of theLSA and its' Exhibits


Page 6-1

6. SPACECRAFT AND LAUNCH FACILITIES

6.1 FACILITIES OVERVIEW

The Baikonur Cosmodrome includes several facilities that are used for ILS launch campaigns (Figure 6.1-1). Thesefacilities include:

a) Yubeleini Airfield—Spacecraft (SC) and ground support equipment (GSE) arrival and departure

b) Area 31 and Area 92A-50—SC preparation and encapsulation

c) Building 92-1—Final integration of the Ascent Unit to Stages 1-3 of the Launch Vehicle Proton K

d) Building 92A-50—Final integration of the Ascent Unit to Stages 1-3 of the Launch Vehicle Proton M

e) Launch Complex Area 81, Pad 23 or 24—SC launch

The sub-sections that follow provide brief descriptions of these facilities. More in-depth descriptions of these samefacilities are provided later in this section.

6.1.1 Yubeleini Airport

Yubeleini Airport is located at the Baikonur Cosmodrome and is used for receiving the charter aircraft carrying the SCand GSE, as well as charter flights with campaign personnel. It is an internationally rated airport with a single 4.5 kmlong, 84 m wide landing strip oriented 60 degrees/240 degrees relative to North. The airport has an elevation ofapproximately 100 m above sea level.

A 140 m by 420 m pad is available next to a railhead for unloading aircraft and transferring equipment to rail convoys.Prior to aircraft arrival, this area is cleared and ground handling equipment is positioned. The pad also is equippedwith stationary spotlights for use in night operations.


Page 6-2

Figure 6.1-1: Baikonur Facilities Map

92A-50Facility

Bldg 92-1

Proton Launch Complex 81 (Pads 23 & 24)

Proton LaunchComplex 200(Pads 39 & 40)

HotelComplex

Yubileini Airfield

Area 254Area 31

Propellant Storage Facility Oxygen and Nitrogen Plant

Saturn Measurement Station

Hwy To Tashkent

Tyura-Tam Rail Station

Krainy Airfield

BAIKONUR TOWN(LENINSK)

Syr dar'ya River

Road:

Railline:

approximately 90 km- figure not to scale


Page 6-3

6.1.2 Building 92A-50

Building 92A-50 contains all facilities necessary for processing a SC from its arrival through its integration with theFourth Stage and encapsulation. The SC and GSE arrive at Hall 102, where the containers are cleaned. The SCcontainer is transported on the railcar or on its own wheels into Hall 101, where the SC is normally removed from thecontainer. The SC is then transported into Hall 103A on the SC transporter where it is installed onto its fueling/teststand. The SC remains in this hall for all subsequent testing and fueling operations. Following fueling, the SC istransported back to Hall 101 on a special transport dolly, where it is integrated with the Fourth Stage andencapsulated.

6.1.3 Area 31 Facilities

Area 31 includes three facilities that serve important functions in preparing spacecraft for launch.

6.1.3.1 Building 40, Hall 100; Area A, B and C

Building 40, Hall 100 is a large high-bay that is divided into three main zones: Area A, Area B and Area C:

a) Area A is a Class 100,000 clean area used for final preparation of the Ascent Unit and its components. It is alsoused for integrating the SC with the flight adapter and mating the SC to the Fourth Stage. Final encapsulation ofthe SC also takes place in this area.

b) Area B is a Class 100,000 airlock used to transition material from Area C into Area A and into Rooms 119, 120,and 121 of Building 40D.

c) Area C encompasses approximately half the building and is used for SC and GSE loading/unloading, as well asthe Fourth Stage and Ascent Unit components.

6.1.3.2 Building 40D, Rooms 119, 120, 121, and Offices

Building 40D is a three-story facility adjoining Building 40 that includes satellite preparation, ground equipment, andpersonnel/office areas:

a) Room 119 is a Class 100,000 cleanroom used for SC testing prior to loading. SC pressurization may occur in thisroom prior to transfer to the Filling Hall (Building 44) with a portable blast shield in place.

b) Room 120 is a Class 100,000 cleanroom used as a control room for test operations conducted in Room 119. Mostelectrical support equipment is located in this room.

c) Room 121 is a Class 100,000 cleanroom used for SC equipment storage, as required.

d) Offices are provided on the second and third floors for the SC Customer and ILS personnel.


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6.1.3.3 Building 44 (Fueling Hall)

Building 44 is a propellant loading facility that is used for loading the SC as well as the Fourth Stage. It is divided intothree main halls.

a) Room 1: Propellant Loading Area. A Class 100,000 clean tent is installed in this room to receive the SC for fueland oxidizer loading operations. Oxidizer propellant cylinders are thermally conditioned in this room in specialchambers prior to loading.

b) Room 2: SC Removal from Container and Container Storage Operations. This room is used for agitation ofpropellant cylinders, storage of the Thermal Transport Container, and storage of the GSE.

c) Room 3: Storage of Propellant Cylinders. Storage of fuel propellant cylinders occurs in this room in prior toloading.

In addition to the areas described above, a control room on the second floor (Room 58) houses SC Customer electricalground support equipment that is used to monitor spacecraft telemetry, charge SC batteries, and communicate withthe propellant-load team.

Fuel and oxidizer decontamination rooms are available for cleaning the GSE. These are located at either end of thebuilding and are supplied with water and nitrogen gas sources.

6.1.4 Building 92-1 and the Proton Launch Zone (Area 81)

Following encapsulation, the Ascent Unit is transported to Building 92-1 for integration with Stages 1-3 of theLaunch Vehicle. For the Breeze M configuration, the Launch Vehicle will return to Building 92A-50, Hall 111 forfinal electrical verification. Following integration, the Launch Vehicle is transported to the Proton Launch Complex(Area 81) for erection and launch. At the Launch Complex, two areas are used for Spacecraft GSE. Rooms 64 and 76(underground vaults) accommodate SC customer equipment providing power to the SC while on the pad. In theBunker, which is approximately 1 km from the launch pad, Room 250 is used for installation of SC customer electricalGSE required for launch.

6.1.5 Hotels

The Hotel Kometa, Hotel Fili, and Hotel Polyot, which are located in Area 95 near the Launch Complex, are used tohouse personnel during a launch campaign.

6.2 SPACECRAFT PROCESSING FACILITIES

This section describes the spacecraft (SC) processing facilities, which provide the capability to perform all requiredoperations from receipt of the SC through its encapsulation in preparation for launch on the Proton Launch Vehicle(LV) at the Baikonur Cosmodrome. These operations include off-loading in the SC technical zone, testing, fueling,mating to the Fourth Stage, payload encapsulation, and LV integration.

The SC processing facilities and Proton launch complex are located at the Baikonur Cosmodrome in the Republic ofKazakhstan in Central Asia, approximately 2,000 km southeast of Moscow. The annual temperature averages 13oC,ranging from -40oC in winter to 45oC in summer. Figure 1.3-1 depicts the overall layout of the Cosmodrome, showingthe facilities that are used for ILS launch campaigns. The specific SC processing areas at Cosmodrome describedinclude:


Page 6-5

a) Area 92, Building 92A-50 (Section 6.2.1)

b) Area 31, Buildings 40/40D (Section 6.2.2)

c) Area 31, Building 44 (Section 6.2.3)

d) Area 254, Building 254-1 (Section 6.2.4)

e) Area 92, Building 92-1 (Section 6.2.5)

The Baikonur Cosmodrome is equipped with spur-railroad service lines that are used for most transport. Specializedequipment is available for fueling, handling of compressed gases, and SC integration with the LV.

The main buildings within the technical complex are the integration and testing facilities. Assembly and integration ofthe various stages of the Proton LV are carried out in the Launch Vehicle Integration Building. Spacecraft preparation,testing, and integration with the launch vehicle fourth stage and the fairing are accomplished either in Area 92,(Building 92A-50) or in Area 31, (Buildings 40/40D), which are both Spacecraft Processing Areas. Spacecraft fuelingand pneumatics pressurization are accomplished in either Area 92, (Building 92A-50) or Area 31, (Building 44) theSpacecraft Fueling Hall). The LV fourth stage with the mated spacecraft and fairing are transferred to the LVTechnical Zone in Area 92, (Building 92-1), where they are horizontally mated to the assembled LV Stages 1-3. Theintegrated LV is then transported to the Proton Launch Zone (Area 81) for erection, checkout, and launch (see Section6.3: Launch Complex Facilities).

6.2.1 Facility 92A-50

Facility 92A-50 is located in Area 92 of the Baikonur Cosmodrome. The facility was modified and outfitted for thespecific needs of SC Customers, and provides the capability to perform required operations in one conveniently locatedarea. These operations include SC offloading, testing, fueling, mating to the Fourth Stage, and payload encapsulation.

This section describes the specific functional areas included within Facility 92A-50, as well as the equipment andservices available to SC Customers for pre-launch payload processing.


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6.2.1.1 Facility Layout and Area Designations

Building 92A-50 has been expressly modified and outfitted to efficiently complete all SC processing and encapsulationin a single building. The halls/rooms, facility systems, and equipment are sized to accommodate SCs of up toapproximately 4.5 m diameter, 10.0 m height, and loaded weights of up to 8,700 kilograms. Figure 6.2.1.1-1 depictsthe overall arrangement of all areas within the building used for commercial programs.

A SC, in the manufacturer’s shipping container, is delivered into Hall 101 on railcar, after having been cleaned in theReceiving Area. In Hall 101, the container is removed from the railcar and placed on the floor. After the railcar isremoved and the environment reestablished, the SC is removed from the shipping container and placed on atransporter to be moved into the Processing and Fueling Hall (Hall 103A). Once there the SC is placed on the fuelingisland and require no further movement in order to complete all necessary standalone assembly, checkout, propellantloading, and pneumatics servicing. When ready, the SC is moved by special transport dolly to the Integration Hall(Hall 101) for mating to the Proton upper stage and encapsulation inside the nose fairing.

Building 92A-50 is approximately 229 m long and 147 m wide. Only a portion of the building is used for commercialprograms.

The Receiving Area (Hall 102) is the primary entrance for the SC and associated equipment, and is located on the eastside of the building.

The main entry into 92A-50 for ILS and SC processing personnel is next to Hall 103A, on the west end of thebuilding, near the Change Room Area. An additional entrance on the north side of Hall 103A is used for ControlRoom equipment delivery.


Page 6-7

Figure 6.2.1.1-1: Building 92A-50 General Arrangement


Page 6-8

6.2.1.2 Receiving/Storage Area - Hall 102

The Receiving Area (Hall 102), or the Integration Area (Hall 101), may be used to offload the SC Container from itstransport railcar. External access for the SC and its GSE is provided by rail through two locally controlled, exteriorsliding doors located in the hall’s east wall. An area is provided outside of Hall 102 for initial wash-down of the railcarand SC container prior to its entry into the building.

An overhead crane is provided in Hall 102 to transfer SC GSE from their railcars to the floor for storage andoffloading.

Container cleaning is accomplished in Hall 102. An area of approximately 240 square meters is provided for non-Class100,000 storage of non-hazardous items.

The overall clear dimensions of Hall 102 are approximately 70.5 m by 36 m. The clear ceiling height is 25.85 m, andthe height of the overhead 50 T crane hook is 18.01 m. The SC unloading area is approximately 8.85 m wide and 34.1m long.

When ready, the SC Container on the railcar is moved, via railcar, from Hall 102 to the inside of Hall 101 where it istransferred to an air pallet for the move into the Processing and Fueling Area (Hall 103A). Alternatively, the SC maybe moved using its own wheeled container, a transport dolly.

6.2.1.3 Integration Area - Hall 101

Once the SC has been processed and fueled in Hall 103A, it is transported to the Integration Area (Hall 101), which isa Class 100,000 cleanroom. The Integration Hall is used to assemble the Ascent Unit, which involves the followingoperations:

a) mating the SC, Adapter, and Fourth Stage

b) checking system continuity

c) rollover of the fourth stage to horizontal

d) encapsulation within the Nose Fairing

Overhead bridge cranes are used to transfer the SC from the Transport Dolly to the rollover fixture, as well astransferring the integrated Ascent Unit from the rollover fixture to a railcar for delivery to the Integration Facility92-1.

This Hall is also used for electrical checkouts of the Breeze M stage prior to fueling.

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, thermalinsulation, and an anti-static floor coating.


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6.2.1.4 Spacecraft Processing and Fueling Hall - Hall 103A (Room 4101)

Hall 103A (also referred to as Room 4101), the Processing and Fueling Hall, is used for pre-encapsulation SCprocessing, including loading propellants and servicing pneumatics in the SC (Figure 6.2.1.4-1). Access to Hall 103Ais made available from Hall 103 through two sliding doors with a clear opening that is 9.5 m wide by 11.95 m high. A15 MT overhead bridge crane, is provided.

An 8 m by 8 m fueling island, located on the west side of Hall 103A, is used for oxidizer and fuel transfer operations. Itis surrounded by a grating-covered trench, which drains any fuel or oxidizer spills into separate waste tanks. Thegrating permits the passage of wheeled dollies.

The floor of Hall 103A has an anti-static coating and a load rating of 10 MT (3,000 kg/cm2) per truck axle. Allfinishes in Hall 103A use materials that do not react with propellants.

The wall between Hall 103A and Hall 103 includes a pair of large doors designed to withstand a 60 kilogram persquare meter overpressure load.

Rooms 4114 and 4116 provide rapid egress routes from Hall 103A, and the Pressurization Airlock (Room 4110)provides the standard egress route. Rooms 4114 and 4116 each have three emergency showers and eyewashes. TheSCAPE Shower Areas (Rooms 4121 and 4122) have showers for post-operation clean-up. A parking lot forambulances and fire trucks is located next to the rapid egress routes. The Pressurization Airlock (Room 4110) and thespace between the double doors between Hall 103 and 103A are pressurized with clean air in order to isolate Hall 103Aduring propellant loading operations.

The clear dimensions for the SC Processing and Fueling Hall 103A are 16.5 m wide by 22 m long.


Page 6-10

Figure 6.2.1.4-1: Building 92A-50 Spacecraft Processing and Fueling Area

BUILDING 92A-50SPACECRAFT PROCESSING &

FUELING AREA

.


Page 6-11

6.2.1.5 Fuel and Oxidizer Conditioning Rooms - Rooms 4112 & 4105

Room 4112, the Fuel Conditioning Room, is used for temporary storage of the campaign fuel (e.g., MMH, N2H4)and thermal conditioning of the fuel before loading. Room 4105, the Oxidizer Conditioning Room, is used fortemporary storage of the campaign oxidizer (e.g.N2O4) and thermal conditioning of the oxidizer before loading (Plate3.1-6). Both rooms have the capability to collect and dispose of propellant spills. Room 4112 contains no materialsthat react with fuel, and Room 4105 contains no materials that react with oxidizer.

The floors in both rooms have an anti-static coating and a load rating of 10 MT(3,000 kg/cm2 ) per truck axle. Thefloor elevations are the same as Room 4101.

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.

6.2.1.6 Fuel and Oxidizer Equipment Decontamination Rooms - Rooms 4111 & 4115

Room 4111, the Fuel Equipment Decontamination Room, is used to decontaminate fuel loading. Room 4115, theOxidizer Equipment Decontamination Room, is used to decontaminate oxidizer loading equipment. Both rooms havethe capability to collect and dispose of propellant spills. Room 4111 contains no materials that react with fuel, andRoom 4115 incorporates no materials that react with oxidizer.

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.

6.2.1.7 Control Room - Room 4102

Room 4102, the Control Room, is used for monitoring and controlling SC processing and fueling activities in Hall103A, as well as SC integration into the Ascent Unit in Hall 101.

A blast-resistant (bulletproof) viewing window is provided between the Control Room and Hall 103A for monitoringall processing and fueling operations. The wall between 103A and the Control Room is a welded, reinforced steelstructure that provides a hermetic seal.

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 entryvestibule, with inner and outer doors 2.9 m wide and 2.9 m high, is provided to facilitate equipment movement intothe Control Room.

The floors of the Control Room and all associated access corridors are designed for wheeled dollies. Forklifts may beused to bring equipment into the vestibule of the room, but they are not permitted to operate in the Control Roomitself. Temporary ramps are available to aide moving items from the entrance vestibule into the Control Room.


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6.2.1.8 Entrance/Lobby Area

The Entrance/Lobby Area includes the following rooms and features:

a) a street-level entrance

b) lobby area (Room 0301)

c) cloak room (Room 0302)

d) SC Customer security checkpoint with viewing windows and security sensor alarm panel (Room 0318)

e) tool storage room (Room 0319)

f) restroom (Room 0320)

Because these rooms serve general purpose administrative functions, they have an office-type environment.

6.2.1.9 Change Room Area

The Change Room Area consists of several rooms (0303 – 0317), including independent men’s and women’srestrooms, change and shower areas, a storage and issue room for cleanroom garments, a Personal ProtectiveEquipment storage and donning room, and a corridor with an air shower. Clean passage is available from the ChangeRooms to either Hall 103 or Hall 103A.

6.2.1.10 Pressurization Airlock - Room 4110

The Pressurization Airlock provides clean access between the Airshower (Room 0317) and Hall 103A. Duringpropellant loading operations in Hall 103A, the Airlock is pressurized slightly more that Hall 103A to prevent vapormigration from Hall 103A. SCAPE suited personnel use the airlock to access the SCAPE Showers and DoffingRooms, and the Pressurization Airlock and Corridor to the Airshower and Change Rooms can be used as anemergency egress route from Hall 103A, if necessary.

Room 4110 is 1.4 m by 3.5 m wall-to wall.

6.2.1.11 SCAPE Doffing Rooms and Showers - Rooms 4108, 4109, 4121 and 4122

The SCAPE Doffing Rooms; Rooms 4108 for Fuel and 4109 for Oxidizer, are available for donning and doffingPersonal Protective Equipment (PPE) for a propellant loading operation. As necessary, SCAPE Showers, Rooms 4121for Fuel and 4122 for Oxidizer, are available to decontaminate the PPE suits before doffing. The dedicated showersare plumbed to the respective liquid waste tanks.

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 mby 3.4 m and Room 4122 is 1.2 m by 3.4 m.


Page 6-13

6.2.1.12 Clean Storage Hall - Hall 103

Hall 103, the Clean Storage Hall, provides accessible storage for clean items supporting SC processing. It also providesa Class 100,000 corridor between Halls 101 and 103A.

A refrigerator/freezer is available for consumable storage. The unit has an internal volume of approximately 1 cubicmeter, and its operating temperature is adjustable from -18 to + 10°C.

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 ceilingheight is 15.0 m. At heights greater than 3 m above the floor, the width of Hall 103 restricted by HVAC ducting toabout 16 m.

6.2.1.13 Ordnance Storage

Khrunichev provides limited storage of ordnance required to support a launch campaign. The Ordnance Storage Roommay be accessed through a door located in the north wall of Hall 101.

Ordnance to be stored must meet the following criteria:

a) A maximum (TNT equivalent) quantity of 50 grams requiring a volume no more than 60 cm by 60 cm by 60 cmmay be stored in accordance with Russian Federation Standards

b) Only insensitive explosives are permitted, and each item must be individually packaged in U.S. Department ofTransportation-approved shipping and storage containers

c) The SC Customer must provide a certificate of conformance to the Hazard of Electromagnetic Radiation toOrdnance (HERO) Specification (MIL-I-23659)


Page 6-14

6.2.1.14 Offices and Conference Room Area - Rooms 1202 through 1209

An Office/Conference Room Area (Rooms 1202 - 1209) is located on the second floor of Building 92A-50 (Figure6.2.1.14). The functions of the seven constituent rooms are:

a) 1202, Facility Management/Security Control and Medical Office

b) 1203, SC Mission Management Office

c) 1204, Break Room

d) 1205, ILS Office

e) 1206, SC Contractors Office

f) 1207, SC Customer’s Office

g) 1208/1209, Conference Room

The clear dimensions of these rooms are as follows:

a) Offices - Rooms 1202 through 1207 are 8.9 m by 5.9 m

b) Conference Room 1208/1209 is 8.9 m by 11.9 m

c) The clear height of all rooms is 3.1 m

Restrooms are accessible from the corridor serving the Office/Conference Room Area; general access to the area is viastairs from the “street” entrance to the Change Room Area. Only essential personnel are permitted in theOffice/Conference Room Area during propellant transfer operations.

Two egress routes are available from the area: the normal route at the western end of the room block that exits to the“street” entrance to the Change Room Area; and an emergency evacuation route that exits east through theKhrunichev area of the building. A 2,000 kg capacity freight elevator, with a 1.8 m wide by 2.2 m high door openingand floor measuring 2.0 m by 3.0 m, is also available in the Khrunichev work area.

6.2.2 Area 31, Buildings 40/40D General Description

Buildings 40 and 40D are located in Area 31 of the Baikonur Cosmodrome. These facilities were modified andoutfitted with the specific needs of SC manufacturers and customers in mind and, along with Building 44, provide thecapability to perform all required SC operations, including off-loading, testing, fueling, mating to the Fourth Stage,and payload encapsulation. Figure 6.2.2-1 provides the general layout of Buildings 40/40D and their location relativeto the Propellant Filling Hall (Building 44).


Page 6-15

Figure 6.2.2-1: Spacecraft Processing Zone (Area 31)

Area C

300-500 m

SC fueling area

PropellantFilling Hall

H a l l 1 0 0

SCArea

Rm.119

Rm.120Rm.121

Area A

Area B

P A S -5 -0 9 1

Building 40Common Hal l

Building 40DPreparation Hall

Ascent Uni tInt egrati onArea

Rm.1

Building 44

This section describes the specific areas of Buildings 40/40D as well as the equipment and services available to SCCustomers for pre-launch payload processing. Payload processing refers to the final preparation of space payloads,upper stages, fairings, and related spaceflight support equipment. It is intended to provide sufficient information toenable customers, and potential customers, to make detailed plans for payload processing activities. It also serves as auseful reference for the facility areas, equipment, and services available during actual payload processing operations.


Building 40, consists primarily of Hall 100, which is divided into Areas A, B, and C. This facility is used to performthe following operations:

a) Off-loading/loading of the SC shipping container and support equipment

b) Processing of the Fourth Stage, the LV adapter system, and the Fairing

c) Ascent Unit integration (i.e., mating the Fourth Stage, the LV adapter system, the SC, and the Fairing, followedby electrical checkouts)

d) Verification of the SC (integrated in the Ascent Unit)

e) Transfer of the SC onto the rail transporter

Building 40D is a three-story facility that adjoins Building 40 and includes satellite preparation, ground equipment,and personnel/office areas.

Details of Buildings 40 and 40D are described and illustrated in the paragraphs that follow.


Page 6-16

6.2.2.2 Building 40, Hall 100 (Common Hall)

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 railin the manufacturer’s shipping container to Area C and transferred into environmentally controlled Class 100,000clean areas (Area B or Room 119) for SC processing and integration.

Hall 100 is approximately 120 m long and 30 m wide, with a height of over 18 m. The total usable area isapproximately 2,500 square meters. The room has 8.4 m wide, 10 m high doors to accommodate transport vehicles.Two overhead traveling bridge cranes are available for lifting and loading operations. This hall houses facilities forvertical SC integration and the necessary mechanical ground support equipment (GSE) for integration operations withthe Fourth Stage, the LV adapter system, the Fairing, and the assembled Ascent Unit. Figure 6.2.2.2-1 shows thehall’s general arrangement.

Figure 6.2.2.2-1: Building 40, Hall 100 (Common Hall) Layout

Two Traveling CranesHook Height = 14.5 mLoad Carrying Capacity 10/50 Tons

Room119

Room120

TravelingCrane Above

120 m

Area A

Area BArea C

Cleanroom forAscent UnitAssembly

Room 100

Locomotive

ThermalControl Car

RailwayTransporterwith SC

Room121

Rooms:100 = Common Hall119 = Payload Preparation Hall120 = STE Lowbay121 = Storage

Ascent UnitIntegrationStand

30 m

6.2.2.2.1 Area A

Area A, which measures approximately 11 m by 47 m, is a Class 100,000 clean area used for final preparation of theAscent Unit and its components. It also is used for integration of the SC with the SC flight adapter and then with theFourth Stage. Final encapsulation of the spacecraft takes place in this zone. It is separated from Area B by an 8 m highpartition that has an upper opening approximately 18 m long and 4 m high used for transferring cargo by cranebetween Areas A and B.

6.2.2.2.2 Area B

Area B measures approximately 19 m by 47 m. It is, essentially, a Class 100,000 airlock used to transition materialfrom Area C into the Class 100,000 Area A cleanroom, as well as through a door into Building 40D, Room 119.

6.2.2.2.3 Area C

Area C measures approximately 30 m by 73 m and constitutes approximately half of Building 40. It is not strictlyenvironmentally controlled and is used for loading/unloading SC and support equipment, as well as the Fourth Stageand Ascent Unit components.


Page 6-17

6.2.2.3 Building 40D, First Floor

The first floor of Building 40D contains the SC preparation hall consisting of Rooms 119, 120, and 121. These roomsare environmentally controlled, Class 100,000 clean areas. Details of this layout are provided in Figure 6.2.2.3-1. Thetotal usable area of Rooms 119, 120, and 121 is approximately 400 square meters. Change rooms and restrooms alsoare located on the first floor.

Figure 6.2.2.3-1: Building 40D, First Floor Layout

Room 119

Room 101

Room 110

Room 120

Room108

Room122

114

Room112

H = 4.5 mRoom 121

Room104

2 m

16 m

9 m

4 m(H = 5 m)

13 m

5.5 m(H = 7 m)

12 m

12 m

3 m

Rooms:101 - Personnel changing

room104 - Air shower108 - Security Office109 - Medical Office112 - Security Checkpoint116 - Restroom120 - Control Room121 - Storage122 - ILS Storage

5. 5 m

PA S- 5- 022

Room111

Main Entrance

Room109

The normal personnel entrance into Building 40D is through an outside door past a security checkpoint. This buildingis the only one completely controlled by U.S. security for the duration of a launch campaign. Security video monitorsare located in the first floor Security Office (Room 108). Access to cleanroom areas (Room 119, 120, 121) is throughan air shower (Room 104) into Room 121.


Page 6-18

6.2.2.3.1 Spacecraft Processing Room 119

Room 119 is used to place support equipment and to prepare the SC prior to propellant loading. The room is anenvironmentally controlled Class 100,000 cleanroom; details of its layout are provided in Figure 6.2.2.3-1. The roomis 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 toHall 100 for the SC. An overhead traveling bridge crane also is available in Room 119.

A header rack to supply compressed air, nitrogen, and helium (at 240 and 350 bars), a blast shield, and service units(to provide access to the SC) are provided to support stand-alone SC processing.

6.2.2.3.2 Control Room 120

Room 120 is used to place electrical support equipment. The room is an environmentally controlled Class 100,000cleanroom 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,5 m high door that opens on to Room 119.

6.2.2.3.3 Storage Room 121

Room 121 is used to place GSE or other spacecraft-related equipment as required. Like the other rooms in this block,Room 121 is an environmentally controlled Class 100,000 cleanroom. It is approximately 12 m long, 7 m wide, and10 m high.

6.2.2.3.4 Support Rooms/Areas

Support rooms on the first floor of Building 40D are shown in Figure 6.2.2.3-1 and include:

a) Personnel changing room (Room 101)

b) Air shower (Room 104)

c) Clean/used cleanroom garment storage

d) Toilet, washroom, and showers

The walls and ceilings of all support rooms are plastered and coated with water-based paint. The walls of toilet andshower rooms are faced with ceramic tiles throughout their entire height. The floors of the support rooms are cast-in-place concrete; the floors of the toilets and shower room are covered with ceramic tiles.

SC loading equipment, packing materials, and containers are stored in Area C or in Building 14, which is located nearBuilding 40. Sealand containers and pallets are stored outside Building 40.

6.2.2.4 Building 40D, Second/Third Floors

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 thephysical arrangements of these areas.


Page 6-19

Figure 6.2.2.4-1: Building 40D, Second Floor Layout

PA S- 5-023

Room 202

Room 201

Rm207

Rm206

Room205

Room 204

Room 203

Room206

Rooms:201 - ILS Office203 - 50 Hz Power

Conditioning Equipment

204 - Security Office206/7 Restroom208 - Khrunichev

Storage

Room207

Room 203

Figure 6.2.2.4-2: Building 40D, Third Floor Layout

P AS- 5- 024

Rooms: 301/307 = SC Customer Offices304 = Automatic Telephone Exchange305 = SC Customer Breakroom306 = Conference Hall311/312 = Restroom

Room303

Room 301

Room31 3

Room305

Room 307 Room 306

Room 305

Room 304

Room308

Rm311

Room312

Room311


Page 6-20

The second floor of Building 40D is used by the ILS team. Room 201 (48 m2) is used as the ILS Office Area, andRoom 204 (32 m2) is used for ILS Security. Room 203 houses 50 Hz power racks and is reserved for Khrunichevengineers. Rooms 206 and 207 are restrooms/washrooms, and Room 208 is used for Khrunichev storage.

The third floor of Building 40D is used by the SC Customer‘s team. It has two available offices: Rooms 301 (48 m2)and 307 (15 m2). Room 305 (42 m2) is normally used as a breakroom. Room 306 (42 m2) is used as a conferenceroom. Rooms 311 and 312 are restrooms, and Room 304 contains Khrunichev communications switch gear.

6.2.3 Area 31, Bldg 44 - General Description

Building 44 is a SC propellant loading facility located in Area 31 approximately 300 meters from Building 40. Thisfacility provides the capability to safely perform SC pre-filling operations, propellant thermal conditioning, and SCfueling in a clean environment. The fueling station has been upgraded to support the loading of the SC withappropriate propellants. (Propellants are supplied by the SC Customer and are loaded using the SC Customer’sequipment.)

This section describes the specific areas of Building 44, as well as the equipment and services, available to SCCustomers for payload propellant filling operations.


Building 44, otherwise known as the Propellant Filling Hall, is an 18 m wide by 72 m long highbay divided into threeequal size bays (Figure 6.2.3.1-1).

Commercial SC enter Building 44 through Room 3 and use Filling Room 1 for SC propellant loading. The SC isplaced on the Customer's fueling stand in a clean tent and secured to a tray platform. Room 1 includes a 5.5 m wide by7 m high sliding door and two emergency exits. Two emergency exits are equipped with panic hardware (each locknormally keeps the door closed but is immediately released when pressed from the inside).

In addition to the main high-bays, a Control Room on the second floor (Room 58) is used for electrical groundsupport equipment to monitor the SC. Fuel and oxidizer decontamination rooms are available for decontamination ofsupport equipment following loading. Areas for propellant thermal conditioning also are available.


Page 6-21

Figure 6.2.3.1-1: Building 44 (Filling Hall) Floor Plan


Page 6-22

6.2.3.2 Building 44, Filling Hall Room 1 (Clean Tent)

Room 1 in Building 44 is used for servicing of propellants and is approximately 24 m long, 18 m wide, and 15 m high.The SC entrance doors are 7 m high and 5.5 m wide.

A clean tent is installed in Room 1 to provide a clean zone for the SC and propellant support equipment. The cleantent consists of two parts: the SC Zone, which measures approximately 9.2 m by 7.6 m; and the Ground SupportEquipment (GSE) Zone, which measures approximately 4.6 m by 6.6 m. Two configuration options are available tothe SC Customer for the clean tent: a pass-through corridor may be installed between the SC and GSE zones tofacilitate personnel passage between the two areas; or a blast shield may be installed between the SC and GSE zones toprevent personnel passage between the two areas. The clean tent has three exits with ramps.

The walls of the clean tent are made of a transparent, non-particulate-forming, anti-static polymer that is compatiblewith propellant vapors. The walls are mounted on a frame made of stainless steel vertical beams and cross-members.The floor of the tent is made of aluminum, with a stainless steel center section, designed to interface with an SC or SCadapter. The aluminum floor has channels that are approximately 2 m apart and 3 to 8 cm deep that provide a pathwayfor liquid spills to drain to one end of the tent and to the spill tanks located underneath the floor of Building 44. Theoxidizer sump drain is located in Room 2, and the fuel sump drain is located in Room 1.

Conditioned air is supplied to the SC and GSE area through the side walls of the tent at two locations to achieve aClass 100,000 cleanroom. The air flows into the tent and leaves through outlets at the bottom of the walls. The tent ispressurized to between 9.8 and 28.6 Pa relative to the outside of the tent. HEPA filters are installed at the inlets of thestainless steel air ducts. Lighting in the clean tent is provided at an illumination of 100 dekalux. The top and front topwall of the tent may be opened to facilitate entry of the SC and GSE after tent installation.

6.2.3.3 Building 44, Room 58 (Control Room)

Room 58 on the second floor of Building 44 is the SC Control Room. Room 58 is approximately 6 m wide and 18 mlong and has a window that is covered with explosion-proof glass to facilitate observation of fueling operations. Anumbilical cable may be laid via a blast-proof penetration located between Room 58 and Fueling Hall Room 1.Telephone, CCTV television monitors, and vapor detection indicators are also provided in Room 58.

6.2.3.4 Building 44, Room 26 and 27 (Fuel/Oxidizer Decontamination Rooms)

Rooms 26 and 27 are available as propellant de-servicing areas to allow the SC Customer to decontaminate SCpropulsion GSE after propellant loading operations. Room 26, the fuel decontamination room, is approximately 5 mwide and 8 m long and is located adjacent to Filling Hall Room 1. Room 27, the oxidizer decontamination room, isapproximately 9 m wide and 9 m long and is located adjacent to Filling Hall Room 3. Access to these rooms isprovided through exterior entrance doors only.

Each decontamination room is supplied with gaseous nitrogen, which may be used to purge SC ground supportequipment. Floor drains in each room are connected to the facility’s appropriate contaminated water sump tanks.


Page 6-23

6.2.3.5 Building 44 Support Rooms/Areas

There are three support rooms on the first floor of Building 44 that are used by SC personnel. These facilities include:

a) Room 11, which provides the entrance to Room 1 and the locker room

b) Room 12, which is used as the staging area for SCAPE personnel

c) Room 14, which is used as the SCAPE suit donning area

6.2.4 Building 254 (TBD)

6.2.5 Area 92, Building 92-1 - General Description

The Proton LV Processing Room, in Building 92-1 at Area 92, is used to horizontally mate the assembled launchvehicle stages and strap-on elements.

6.2.5.1 Facility Layout & Area Designations

Building 92-1 is an industrial-type, unprotected building constructed from concrete, brick, and welded/riveted steelthat measures approximately 50 m wide and 120 m long (Figure 6.2.5.1-1). The building includes theintegration/assembly room and two laboratory annexes adjoining the assembly room. The building is provided withheating, ventilation, and fire-fighting systems, fire and security alarms, and special lighting. An overhead crane isavailable for handling the integrated LV.

The Proton LV Processing Room in Building 92-1 is used to:

a) Transfer the Ascent Unit from its railway transportation unit to a mating stand.

b) Mate the Ascent Unit to the Proton LV

c) Conduct electrical checks of the LV transit circuits and verify the hardware links between the Ascent Unit and theLV

d) Charge the SC’s onboard batteries

e) Install the Thermal Cover on the Ascent Unit

f) Transfer the integrated LV to the erector

g) Prepare the integrated LV for transport to the launch complex

The Proton LV Processing Room is approximately 30 m wide, 119 m long, and 22.9 m high. The room has three railtracks with a gauge of 1,524 mm. The central track and one of the lateral tracks are throughways, while the third trackterminates in the Processing Room. Along the building’s central axis, there are three double-leaf gates measuring 4.7by 5.6 m, and one electric-operated rolling (central) gate measuring 8.4 by 10 m.

Temperature and humidity in the room are maintained between 13 and 27oC and between 30 and 60 percent,respectively. While it is in the room, the temperature and humidity of the SC are maintained within required levels byusing the Thermal Conditioning Railcar, if necessary.


Page 6-24

Figure 6.5.2.1-1: Building 92-1

6.3 LAUNCH COMPLEX FACILITIES

Following integration of the Proton Launch Vehicle (LV), in Building 92-1, the launch vehicle is transported to theProton Launch Zone, Area 81, for erection, checkout, and launch. This section describes the facilities that are used forerection, checkout, and launch of the Proton LV at the Baikonur Cosmodrome. In particular, areas available at thelaunch complex for Spacecraft (SC) equipment and personnel are described. These areas may be used by the SCCustomer to test and monitor the SC and to charge SC batteries during final LV checkout and launch.

6.3.1 Area 81, Launch Pad 23 - General Description

Following integration, the Proton LV is transported to the Launch Complex, Pad 23, located in Area 81 for erectionand launch. Figure 6.3.1-1 shows a plan view of the launch complex. The launch area includes the following physicalfacilities, units, and systems that support processing and launching of the Proton LV:

a) Launch structure with launch pad (including underground vault)

b) LV Mobile Service Tower

c) Bunker (Room 250)

d) Facilities for support systems


Page 6-25

Figure 6.3.1-1: Proton Launch Zone, Area 81

BunkerRooms 250/251

PUG968

600 m

340 m

LaunchPad 23

LaunchPad 24


Page 6-26

6.3.2 Facility Layout & Area Designations

6.3.2.1 Launch Structure with Launch Pad (Including Underground Vault)

The launch structure and vault house equipment that supports the pre-launch processing of the LV. They provideelectric, pneumatic, and hydraulic links between the ground system testing equipment and on-board LV hardware viathe LV transit cables and pipes. The launch structure is designed to withstand LV first-stage engine plumeimpingement. The launch pad is intended for LV installation, erection, securing the LV in a vertical position.

6.3.2.1.1 Room 64 - Vault

Room 64 (Vault), along with Room 250 (Bunker), can be used to house the SC Customer’s ground support equipment(GSE). Room 64 measures approximately 5.1 by 5.6 m and is equipped with 60 Hz electrical power, grounding, andcommunications services. All launch campaign operations requiring the presence of personnel in the vault must becompleted prior to the start of LV fueling, which occurs approximately 7 hours prior to launch. All personnel arerequired to leave the vault by this time, and from then on, all vault equipment must be controlled remotely from thebunker (Room 250). Additional details about Room 64 are provided in Drawing 4.1-1. Plate 4.1-1 provides aphotographic overview of the vault’s interior.

6.3.2.1.2 Room 76 - Vault

Room 76 can be used to house the SC Customer’s GSE. Room 76 measures approximately 5.4 m by 10.8 m and isequipped with 60 Hz electrical power, grounding, and communications services. All launch campaign operationsrequiring the presence of personnel in the vault must be completed prior to the start of LV fueling, which occursapproximately 7 hours prior to launch. All personnel are required to leave the vault by this time, and from then on, allvault equipment must be controlled remotely from the bunker (Room 250).

6.3.2.2 Mobile Service Tower

The Mobile Service Tower provides access to the SC and LV and houses equipment to support SC and LV pre-launchprocessing and launch. The Mobile Service Tower includes service platforms, a gallery, service fixtures, two cargo-passenger elevators (500 kg rated load capacity each), and two cranes (rated load capacity 500 and 5,000 kg,respectively). An overall view of the Mobile Service Tower is provided in Figure 6.3.2.2-1.


Page 6-27

Figure 6.3.2.2-1: Proton Mobile Service Tower

30.400 m30.515 m


Page 6-28

6.3.2.3 Rooms 250/251 - Bunker

The Bunker (Rooms 250/251) is used to support a Proton launch. It is located 1.3 km from the launch pad andprovides protection for personnel and equipment during the launch. The bunker houses the LV and SC system testequipment and GSE required for pre-launch operations and monitoring of SC readiness. Air temperature andhumidity inside the bunker are controlled by an air-conditioning unit.

Room 250 is allocated for installation of SC equipment, and Room 251 is the ILS Security Office.


Page 7-1

7. LAUNCH CAMPAIGN

7.1 ORGANIZATIONAL RESPONSIBILITIES

In any given launch campaign several organizations are involved. What follows is a brief overview of the responsibilitiesof these organizations.

7.1.1 Khrunichev

a) Overall responsibility for coordinating work performed at the launch complex by the Strategic Rocket Forces (andNPO Energia if Block DM used)

b) Engineering support and quality inspection for all testing performed on Stages 1-3 of the Launch Vehicle, as wellas the adapters and fairing. For launches with the Breeze M Khrunichev is also responsible for Breeze Mengineering, inspection and test

c) Maintenance of Buildings 92A-50 (Halls 103A, 103, and 101) and 40/40D; and the hotel complex

d) Transportation and food services

e) Coordinating Baikonur and Leninsk medical services with Strategic Rocket Forces

f) Launch complex security

7.1.2 Strategic Rocket Forces

a) Maintenance of Building 44, Building 92-1, Building 92A-50 (Hall 102), and the Launch Complex

b) Provision of technicians for performing launch vehicle testing

c) Provision of quality inspectors

d) Launch Vehicle operations from integration in Building 92-1 through erection on pad and launch

e) Launch complex security

7.1.3 Energia (Block DM launches only)

a) Checkout and processing of the Fourth Stage at the Baikonur Cosmodrome from its arrival through the launch.


Page 7-2

7.1.4 ILS

a) Prime interface between the Customer and Khrunichev

b) Coordinating campaign schedules and operations with the SC Customer and Khrunichev

c) Logistics

d) Safety overview as an advisory function to SC and Customer management, as well as physical security of SC assetswhile at Baikonour processing facilities

e) Translation and interpretation services

f) Medical doctor and SOS evacuation

7.1.5 Spacecraft Customer

a) Spacecraft checkout and processing at the Baikonur Cosmodrome

7.2 CAMPAIGN ORGANIZATION

7.2.1 Contractual and Planning Organization

The fundamental contractual relationships among the principal parties in any given launch campaign are as follows:

a) ILS is a contractor to the SC Customer

b) Khrunichev is a subcontractor to ILS

All matters that could potentially affect the terms of the Launch Service Agreement (the contract) between a SCCustomer and ILS must be dealt with by the SC Customer and ILS. Matters affecting the terms of the subcontractbetween ILS and Khrunichev must be dealt with by ILS and Khrunichev. In particular, any issues involving possibleadditional costs must be mutually agreed upon through these contractual relationships.

ILS will coordinate all logistics support and operations planning with both the SC Customer and Khrunichev.

ILS may assign a safety engineer to monitor any given operation to ensure that all activities are carried out inconformance with the mutually agreed upon safety plan. This safety engineer is present for all hazardous operations.

Figure 7.2.1-1 provides a diagrammatic representation of a typical operational organization used during a launchcampaign.


Page 7-3

Figure 7.2.1-1: Organization During Launch Campaign

Customer MM

SC Subcontractor

SC Ops Manager ILS Ops Manager KhrunichevOps Manager

ILS MM KhrunichevProgram Director

Other Subcontractors

Russian StrategicRocket Forces

7.2.2 Organization During Combined Operations

Combined operations are those operations involving some combination of SC Customer organization, ILS, andKhrunichev personnel (e.g., Khrunichev adapter mating, Fourth Stage to payload mating, encapsulation, Ascent Unitcheckout, Ascent Unit integration to Stages 1-3, any operations on the pad which require payload faring access). Foreach such combined operation, one operation leader is assigned, either from Khrunichev or the SC Customer. Thisindividual directs the operation and ensures that it is carried out in conformance with the mutually agreed uponprocedures.

For each operation, one person from the SC Customer organization, ILS and Khrunichev is designated as team leaderfor their respective organizations. Agreements among organizations can only be reached among these three teamleaders.

Security personnel from either or both ILS and the Strategic Rocket Forces may be present during any operation ifrequired by the Security Plan.

ILS provides at least one interpreter for each combined operation. Special training is conducted with the SC Customerand Russian personnel for joint crane operations to ensure reliable communications between English and Russian-speaking personnel.

Either Khrunichev or the SC Customer provides a Quality Assurance Representative for each operation whodocuments any test discrepancies on a Quality Assurance Report.

Strategic Rocket Forces personnel implement many of the operations at the Baikonur Cosmodrome. The StrategicRocket Forces is a Khrunichev subcontractor and, as such, coordinates directly with Khrunichev and NOT with the SCCustomer or ILS personnel.


Page 7-4

7.2.3 Planning Meetings

ILS maintains the master schedule for combined operations planning and reviews it with the SC Customer andKhrunichev at a daily scheduling meeting. At this meeting, the current operations for the day are reviewed and the next3 day’s operations are agreed upon. Following each meeting, ILS revises the master schedule for the following day. Ata minimum, the following personnel must attend the daily scheduling meetings:

a) ILS Mission Manager or Launch Operations Manager

b) SC Customer Launch Campaign Manager

c) Khrunichev Mission Manager

d) Strategic Rocket Forces representative in charge of the facility

At certain stages in the campaign, ILS holds meetings to give the go-ahead for critical phases of a campaign. Thesecritical phases include:

a) SC offload and move to Integration Hall

b) SC Processing and Propellant Loading

c) SC Encapsulation

d) Launch

Two critical meetings require high-level concurrence prior to proceeding to the next phase of the campaign. These are:

a) Vehicle Readiness for Transport to Pad—5 days prior to launch

b) Vehicle Readiness for Propellant Load—10 hours prior to launch


Page 7-5

7.3 COUNTDOWN ORGANIZATION

The countdown organization is illustrated in Figure 7.3-1.

Figure 7.3-1: Countdown Organization

SC CustomerMissionDirector

ILSMission

Manager

EnergiaMissionDirector

(if Block DMused)

KhrunichevMissionDirector

KBOM *

ReadinessReview BoardAuthorization Strategic R ocket

Fo rces LaunchComm ander

AutomaticComputer-Control led

Sequencer

LIFT-OFFCOMM AND

Spacecraf tDirector

(SC Customer)

Fou rth StageDirector

(Energia) i fBlock DM

used)

Stages 1-3Director

(Khrunichev)

SC R eadyL - 10 min

Fourth StageReadyL - 2 min

LV R eadyL - 5 min

*Design Bureau for General Machine En gineering (Launch Site) Note: Launch Co mma nd er has so le autho rity to abor t launch

LaunchAuthorization


Page 7-6

The Strategic Rocket Forces directs the countdown, which follows a pre-approved script. The Launch Commanderreceives authorization to launch from the Readiness Review Board, which consists primarily of the five entities shownin Figure 7.3-1.

Certain organizations have pre-assigned abort capability. Each of these organizations is asked to acknowledge thereadiness of their subsystems on Launch Day according to the Launch Day Script. These subsystem readiness checksare as follows:

a) Stages 1-3 Readiness: Khrunichev

b) Fourth Stage Readiness: Khrunichev for Breeze M, Energia for Block DM

c) SC Readiness: SC Customer

Each organization designates a single individual to provide the above authorities on Launch Day, and eachrepresentative is vested with abort authority over the launch sequencer for their respective area of responsibility. Forexample, the SC Customer may abort the start sequence as late as 2.5 seconds prior to lift-off contact.

7.4 ABORT CAPABILITY

There is a ground circuit that, when closed, sends a command to an onboard switch that, in turn, activates on-boardpower to the first three stages and begins the sequence for first stage engine ignition. Normally, readiness signals aregiven in the following sequence:

a) Launch - 10 minutes: SC Ready

b) Launch - 5 minutes: Stages 1-3 Ready

c) Launch - 2 minutes: Fourth Stage Ready

Once all readiness signals are received, the command circuit is ready and is waiting only for the Launch Vehicle StartCommand to be issued automatically by the Start Timer Mechanism. This command is issued at 2.5 seconds prior toliftoff, at which time the final relay closes, sending a command transferring to on-board power and initiating the FirstStage engine ignition sequence.

The SC Customer is capable of aborting the Launch Vehicle Start Command issued by the Start Timer Mechanism.The launch aborts if this circuit is open when the Launch Vehicle Start Command is issued. If the circuit is openedfollowing issuance of the Launch Vehicle Start Command, the launch can no longer be aborted. The SC Customer canabort the launch countdown up to 2.5 seconds prior to lift-off contact.

The Fourth Stage Manager can abort the launch by removing power from the command circuit. He has this capabilityup to 2.5 seconds prior to lift-off contact.


Page 7-7

7.5 LAUNCH OPERATIONS OVERVIEW

This section provides an overview of Launch Vehicle (LV) and Spacecraft (SC) processing.

Figure 7.5-1 provides a similar schedule assuming the use of Building 92A-50 for SC processing. Figure 7.5-2 providesdetails of the SC processing timeline using Area 31 facilities.

The typical duration of a launch campaign from SC arrival to launch is 30 days, depending on SC manufacturer andCustomer requirements.


Page 7-8

Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50


Page 7-9

Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50 (Continued)


Page 7-10

Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31


Page 7-11

Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31 (Continued)


Page 7-12

7.5.1 Launch Vehicle Processing

The Proton LV stages and fairings are built in Moscow by Khrunichev and transported by rail to the BaikonurCosmodrome typically well in advance of the SC delivery date. After transportation of the Proton’s stages and fairingby rail, LV build-up takes place in an integration and test facility (Building 92-1) capable of supporting foursimultaneous Proton assembly and checkout operations. The Fairing is moved either to Building 40 or to Building92A-50, depending upon which facility is to be used for SC integration, prior to SC arrival, there it is stored andcleaned in preparation for encapsulation. The Fourth Stage is delivered to Area 254 for pre-launch checkout andtesting. The Fourth Stage is then delivered to the Building 44 in Area 31, the propellant fueling hall, where eitherkerosene is loaded (Block DM) or MMH and N2O4 (Breeze M). It is then moved to either Building 40 or 92A-50 forintegration with the SC. Payload adapters typically are delivered shortly before the processing cycle and cleaned andprepared in the Integration Hall (100A or 101) of whatever processing area is being used for that program.

In advance of SC arrival, Payload Processing Facilities undergo facility activation and certification. Buildings 40,40D, 44, or 92A-50 are verified to meet environmental control and cleanliness requirements, in addition tocommodities and power support requirements usually a week prior to the SC arrival date.

7.5.2 Spacecraft Preparations Through Arrival

Prior to campaign start, SC propellants are shipped by rail from St. Petersburg to the Baikonur Cosmodrome. Theyare stored in the same temperature-controlled railcars used for transport from St. Petersburg until required for fueling.A pre-determined number of hours prior to fueling, as defined by the spacecraft manufacturer, propellant containersare transferred to a storage/conditioning room for temperature stabilization.

The SC and its ground support equipment (GSE) arrive at Yubeleini Airport via a SC Customer-chartered aircraft,where they are offloaded and loaded onto railcars. These operations are supported by Khrunichev-supplied mobilecranes, K-loader, and 5 and 15-ton forklifts as required. After the SC container is placed on a railcar, it may beconnected to a thermal control railcar via two air duct flanges (inlet and outlet air flow) to provide thermalconditioning during transport. Some spacecraft containers are completely self-contained thermally andenvironmentally and do not require this support option. The thermal car also provides a dynamic load monitoringsystem, so use of the car may be limited to use of this system. SC Customer personnel effects may be transporteddirectly to the hotel by truck.

7.5.3 Area 31 (Buildings 40, 40D, 44) - Spacecraft Testing, Fueling, and Ascent Unit integration

The SC and its GSE are transported by rail convoy approximately 70 km to Area 31; transport duration isapproximately 7-10 hours, depending on time of day of transport.

The SC and its GSE are offloaded outside of Building 40, Room 100, Zone C using 5 and 15-ton forklifts. TheBuilding 40 bridge crane may be used for offloading large GSE such as the SC container. Spacecraft GSE andcontainer handling proceed as required for each unique spacecraft operations flow.


Page 7-13

A typical flow might progress as follows:

a) SC container offloaded from railcar in Rm 100C, using overhead bridge crane, and container lid removed

b) SC rotated to the vertical position and moved into Rm 119, placed on transporter

c) SC electrical checkout equipment moved into Rm 119

d) SC standalone testing, electrical and mechanical systems

Note: A portable blast shield is available for use if required for high-pressure leak checks

Following Fourth Stage filling in Building 44 (Filling Hall), Room 1 of the Filling Hall is readied to receive the SC.This includes setting up the clean tent and all fluid/gas systems. The SC may be bagged to protect it during any briefexposures to less than Class 100 000 environments. It is then rolled out of Building 40D Room 119 to Building 40,Room 100, Zone B, where it is transferred to the support on the base of the Thermal Transport Container. The SCmaintains its vertical orientation, as the Thermal Transport Container is designed to interface to the SC adaptermating surface. This container is used to transfer the SC to the Filling Hall for propellant fueling. Once on thecontainer base, the SC container cover is installed, and the container is lifted onto the railcar in Zone C for transport.The thermal control railcar may be used to provide conditioned air and dynamic load data during transportation. Useof this temperature control system may be waived by the SC Customer since the nominal time for transportation to theFilling Hall is of short duration (less than 1 hour). The Thermal Transport Container has been specifically designed toprovide thermal stability for an extended period of time.

The SC is moved to Building 44 (Filling Hall) 250 meters away, and the Thermal Transport Container cover isremoved in Room 2 and placed on support stands. The railcar is then rolled into Room 1 and the SC is hoisted fromthe Thermal Transport Container base onto the SC loading stand already installed in the clean tent. The clean tentceiling is closed and a Class 100 000 environment is established.

Propellants are loaded in accordance with the SC Manufacturer’s procedure. The facility provides passive vapor ventand active vapor extraction, as well as liquid waste disposal and commodities such as water, breathing air, and GN2.Propellant GSE is decontaminated using rooms in the Filling Hall specifically dedicated to this purpose. Hot and coldGN2 purge is available as required.

Following fueling, the process is reversed and the SC is returned to the Thermal Transport Container and then movedback to Building 40, Room 100 for further testing and processing. Electrical checkouts are made on the adapterelectrical harnesses. The SC is moved from Zone B to Zone A and placed onto the flight adapter (s).

The SC and its adapter is then mated to the Fourth Stage. For certain SC, this may require the use of an integrationstand commonly referred to as the “Gallows”. The Gallows allows taller payloads to be installed onto the FourthStage without structurally modifying Building 40D. The flight clamp band is installed and tightened to the correctflight-specific tension.

Following integration, an end-to-end electrical checkout is performed. The composite Ascent Unit is rotated to thehorizontal for encapsulation with the two fairing halves. Following encapsulation, an RF GO/NO GO test isperformed through the RF window. The Ascent Unit is then transferred to its railcar for transport to Area 92. Duringtransport, the Ascent Unit is thermally conditioned using the thermal conditioning railcar; input air to the Ascent Unitis provided at the base of the Fourth Stage and output occurs at the Fairing nose.


Page 7-14

7.5.4 Area 92 (Building 92A-50) - Spacecraft Testing, Fueling, and Ascent Unit Integration

In this processing scenario, the SC and its GSE are transported by rail convoy approximately 30 km to Building 92A-50, Room 102 , where external cleaning of the SC container is performed in Hall 102. Final cleaning is completed inthis area prior to moving the SC to the Integration Hall (Hall 101). Hall 101 is a Class 100,000 cleanroom, and theSC container cover may be removed here or in Hall 103A as required by the unique mission-specific SC processingflow.

A typical flow is as follows:

a) Move the SC container into Hall 101 on railcar

b) Remove container from the railcar and place on the floor in Hall 101

c) Remove the container lid

d) Roll container into 103A if on wheels or

e) Lift SC Container onto KhSC-supplied air pallet or Customer-supplied transport dolly to roll into 103A

f) Move SC on transporter into Hall 103A for processing

Note: Container lid removal may also be accomplished in Hall 103A as required by the Customer.

The SC container and lid may be stored in Hall 101. Electrical test equipment is brought into the control room bymeans of an external door which opens directly into the control room loading area. This is a small buffer zone betweentwo sets of double doors with a concrete floor.

After container removal, SC electrical testing, pneumatic testing, and propellant fueling occurs in Hall 103A. Pass-throughs from the control room are available for cabling. These cable feeds are verified to be leak-tight prior topropellant operations. The 60 Hz, 120 Vac power source is provided by an UPS. Typically the UPS is activated theweek prior to SC arrival and not deactivated until the SC leaves the facility and all parties agree that no furtherrequirement for it exists. A portable blast shield is available for high-pressure tests.

For propellant operations, the facility is configured with liquid waste aspirators, passive vent scrubbers, and a vapordetection system which alarms locally, in the control room, and at the Security Command Post. Breathing air issupplied by a single source which is sampled prior to operations. GN2, water, and shop air are provided on demand. Afire suppression system which will arm but not release on alarm is also active in Hall 103A. The command to activatethe suppression system deluge is made in the control room, and is not an automatic function of the alarm system. LN2is available with 24 hr call-up.


Page 7-15

The loaded SC is transported back to Hall 101 using the Transport Dolly. The SC is lifted from the transporter andmated to the SC adapter using the 50T bridge crane. The SC/adapter unit is then lifted and mated to the Fourth Stagewhich has been previously installed in the Integration Stand. The adapter clampband is installed and tensioned.Incremental electrical continuity tests are performed at each phase, with the final check being an end-to-end test withthe SC mated to the Fourth Stage.

After SC integration, final closeout operations and photographs are performed. The combined Fourth Stage/SC isrotated to the horizontal position on the integration stand and is encapsulated by the bi-conic payload Fairing. Afterthe upper fairing half is emplaced, an RF Go/NoGo test is performed to ensure that the SC link has not been disturbedand that the RF window is transparent to RF. This is performed as soon as the fairing half is mechanically emplacedand before continuation of encapsulation sealing operations. If any anomaly is found, the fairing may be removedrelatively easily at this point. After determining a good RF signal, encapsulation is completed.

After encapsulation and required RF testing, the integrated Ascent Unit is placed on a railcar for transport to Building92-1 (Integration Hall) for integration with the Proton Launch Vehicle. At this point, the processing flow for all SCbecomes a common flow, independent of SC Processing Facility.

7.6 LAUNCH VEHICLE INTEGRATION (BUILDING 92-1) THROUGH LAUNCH PADOPERATIONS

The Ascent Unit is transported to Building 92-1, where it is disconnected from the thermal conditioning car and liftedonto the integration dolly. Here, it is brought together horizontally with the mating surface of Stage Three of theintegrated Proton launch vehicle (LV). An end-to-end electrical check is performed on the SC/LV cables. Theintegrated LV is then transferred to the transportation/erection fixture and a thermal blanket is installed over theFairing to protect the payload from temperature extremes during periods when there is no active thermal control. Atypical launch flow requires 4 days in the Integration Hall. Activities in the Integration Hall are controlled by theStrategic Rocket Forces and are based on the LV Launch schedule.

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 theBreeze M, the LV will be transported first to 92A-50 for final integrated Breeze M tests. This will be followed bytransportation to the Breeze M Filling Station for top off of the low pressure MMH and N2O4 reservoirs on its way toLaunch Pad 24.

The first of two Inter-governmental Commission Meetings is held on Day L-6, prior to vehicle roll-out to the pad, toensure all agencies are go for pad roll-out. All LV agencies, including the SC Manufacturer will be called upon toprovide a launch readiness statement.

The LV along with the thermal control railcar is moved to the launch pad. The vehicle is erected, and the liquidthermal conditioning system is activated. An RF check of the SC telemetry and command links is performed from theBunker (Room 250) prior to Mobile Service Tower rollup. Once the Mobile Service Tower is in place, the Ascent Unitair-conditioning system is activated and the liquid system is turned off.


Page 7-16

7.7 LAUNCH PAD OPERATIONS

The Integrated LV is transported to the Launch Pad and erected in one piece at L-5 using the Launch VehicleTransporter. From L-5 on, the launch schedule is driven by the Strategic Rocket Forces overall countdown schedule.Coordination of SC-related pad activities is performed through the 7/701 Script. The 7/701 Script is generated byKhrunichev with SC/Customer input and should include all pad access requirements and requirements for RFradiation and commanding the SC. Operator “7” is the Khrunichev Program Director while Operator “701” is the SCPoint of Contact. Information required from the SC/Customer: RF radiation, battery charging, SC commanding, padaccess. Note that ILS functions such as pad access will also be coordinated through this script. Figure 7.7.1 provides atypical detailed on-pad operations flow.

The SC Customer is expected to participate in a Launch Countdown Rehearsal on Day 3 on pad. This countdownrehearsal is supported by a full booster vehicle launch crew countdown and requires the SC Customer to indicate SCreadiness to go at the required time. The rehearsal also includes a planned abort on a SC No-Go condition. SC fullfidelity countdown rehearsal is not required for this exercise, simply the operation of the readiness switch at theplanned time in accordance with the 7/701 script.

The second Inter-Governmental Commission Meeting is held at T-8 hours, to ensure all agencies are go for launchprior to propellant load of the booster vehicle. At T-8 hours, the launch pad is cleared of all non-essential personnel,and at T-6 hours, propellant load commences. At T- 2.5 hours, the pad is open for final closeouts and service towerremoval. At T-2 hours, all personnel are cleared from the hotel areas and should be in their final positions for launchi.e. Bunker, Viewing Area, and Comm Center. Note that personnel in the Bunker and the Comm center should belimited to essential personnel only.

The SC Customer participates in the final countdown by sending a SC readiness to launch signal at T-10 minutes, asnoted in the Countdown Organization discussion.

Active commanding of the SC is prohibited during critical booster vehicle functions. Propellant fueling of the boostervehicle, which starts at T- 8 hours, is one such time-frame. Other typical RF silence and no-command times areshown in Figure 7.7-2.


Page 7-17

Figure 7.7-1: Typical Launch Pad Operations Timeline


Page 8-1

8. PROTON LAUNCH SYSTEM ENHANCEMENTS

The Proton launch system is supported by the Russian space industry leaders Khrunichev State Research andProduction Space Center and Russian Space Complex Energia. Based on the long, successful history of the Protonlaunch system, ILS, Khrunichev, and Energia have continued to assess and conceptually study hardware and launchservices enhancements that would improve the competitiveness and capability of the Proton launch system.


Page 8-2

Figure 8-1: Proton Evolution Options

P r o t o n - K M

2 0 0 0 -2 2 . 0 M T 4 . 5 M T

Y e a r s o f o p e r a t i o nL E O c a p a b i l i t yG S O c a p a b i l i t y


Page 8-3

A number of enhancement options are being studied for the Proton system including enhanced payload fairing options,avionics enhancements, and upper stage modifications to further improve capabilities of Proton heavy lift launchsystem.

8.1 HIGH VOLUME PAYLOAD FAIRINGS

With the quantity of LEO, GTO, and planetary spacecraft on the Proton launch record, a number of payload fairingconfigurations have been developed in the past to uniquely meet the needs of the diverse Russian governmentprograms. Most notable in recent developments has been the launch of large low earth orbit payloads such as the MIRspace-station using the three stage Proton vehicle.

With adoption of the Breeze M upper stage, larger payload fairings can be easily mounted to the Proton boostervehicle. Unlike the Block DM configuration vehicle, the design of the Breeze M enables the stage to be completelyencapsulated inside the fairing. For the Breeze M configuration, the payload fairing is attached directly to the thirdstage of Proton enabling large volume/higher mass fairings to be successfully carried with the limits of the Protondesign. A number of payload fairing options have been assessed for development with the commercial ProtonM/Breeze M configuration. Figure 8.1-1 illustrates some of those options.

With these conceptual systems, usable volume diameters as great as 4.2-meters have been envisioned. Usablecylindrical volume lengths of up to 15,000-mm have been assessed and determined feasible with the existing controlauthority capability of the Proton launch system.

ILS and KhSC are willing to develop these concepts with the award of firm launch services contracts. These fairingoptions can be fielded in 30-36 months of authority to proceed and will support Proton M and Proton M/Breeze Mvehicle launches as early as mid-2001.


Page 8-4

Figure 8.1-1: Proton Larger Payload Fairing Concepts

TBS


Page 8-5

8.2 TANDEM LAUNCH SYSTEM

In addition to large volume payload fairing options, ILS and KhSC have studied and conceptually designed a tandemlaunch system payload carrier concept. The tandem launch system concept may offer attractive opportunities forlaunch of multiple payload constellations to low and intermediate altitude orbits. Figure 8.2-1 illustrates possibleinterfaces available with the tandem launch system concept.


Page 8-6

Figure 8.2-1: Breeze-M Launch Configuration With Tandem Launch Systems (TLS)

4395

2 14 5

1 0 3 0 0

1 9420

6 1 7 0

1 0 0

63 0

600

1 0 0

Φ 4350

2650

Φ 4100

Φ 2490

30

20


Page 8-7

8.3 AVIONICS SYSTEM MASS UPGRADES

With the introduction of the Breeze M stage, it becomes possible to remove the booster avionics system from the thirdstage of Proton and use a modified upper stage avionics system for control of the entire launch system from liftoff toseparation of the payload in the targeted orbit. Removal of the avionics system increases typical performance togeosynchronous transfer orbits by approximately 100 kg.

8.4 NEW CYROGENIC UPPER STAGE

In a the course of a vehicle development contract with another entity, Khrunichev has developed a liquid oxygen,liquid hydrogen upper stage. This upper stage can be adapted to fly on the Proton booster and could significantlyenhance launch system performance to high energy transfer orbits. Payload performance gains to GSO and GTO ofalmost 50% are possible over that available with the Breeze M upper stage.

8.5 SUMMARY

The mature partnership between Lockheed Martin and Russian partners Khrunichev Space Center and RSC Energia isnow enabling ILS to explore the next evolutionary steps in effectively supporting Proton commercial launch services.ILS is ready to discuss, with potential customers, straight forward enhancements in Proton launch services capabilityto meet near term commercial launch services needs.


Page A-1

A. PROTON LAUNCH SYSTEM HISTORY

A.1 BACKGROUND AND HISTORY

Development of the Proton launch vehicle was undertaken in the early 1960s, under the direction of the Sovietacademician, V. N. Chelomey. The first launch took place in July 1965. The two-stage 8K82 (D) version, last flownin 1966, was used to launch four flights of the Proton satellite series, from which the launch vehicle takes its name.The two-stage version has been superseded by the three stage Proton K also known as (D-1, SL-13) model and thefour-stage Proton K/Block DM (D-l-e, SL-12) model, both of which are currently in use. An improved version ofthe Proton (Proton M) is now in development, incorporating changes to the first three stages, as well as thedevelopment of a new storable propellantBreeze M upper stage. Figure A.1-1 shows the major variants of the Protonlaunch vehicle family.

Figure A.1-1: Proton Launch Vehicle Family

1965-196712.2 MT

--

8K82(Proton D)

1967- Present-

4.9 MT1.9 MT

Proton K/Block DM(Proton D-1-e)

196719.76 MT

--

Proton K(Proton D-1)

1999- Present21.0 MT5.5 MT

2.92 MT

Proton-MBreeze-M

Years of operationLEO capabilityGTO capabilityGSO capability


Page A-2

The Block DM fourth stage of the Proton was developed independently during the 1960s as the fifth stage of theRussian manned lunar launch vehicle, the N2-L3. It was originally known in Russia as the Block D ("block" is thecommon translation of the Russian word denoting a rocket "stage", while "D" is the fifth letter in the Russianalphabet). The vehicle was upgraded during the 1970s to the current Block DM (modernized) version.

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,with the D numbers having been applied by the Library of Congress and the SL numbers originating with theDepartment of Defense.

Proton has flown more than 235 missions, and has carried the Salyut series space stations and the Mir space stationmodules. It has launched the Ekran, Raduga, and Gorizont series of geostationary communications satellites (whichprovided telephone, telegraph, and television service within Russia and between member states of the IntersputnikOrganization), as well as the Zond, Luna, Venera, Mars, Vega, and Phobos inter-planetary exploration spacecraft.The Proton has also launched the entire constellation of Glonass position location satellites. All Russian geostationaryand interplanetary missions are launched on Proton. Approximately 90% of all Proton launches have been the four-stage version.

The Proton launch vehicle has a long history of outstanding reliability. From its first operational launch in 1970 to thepresent day, Proton has averaged a 92.5% success rate. Today the Proton launch vehicle has a 92% (moving average)success rate over its last 50 launches. The recent history of Proton's launch reliability is shown in TableA.1-2.

Table A.1-2: Proton 50-Launch Moving Average

Year Proton Block DM Proton (3 stage) Failures/Cause

1992 8

1993 8 1 booster second and third stage propulsion failure

1994 13

1995 6 1

1996 7 1 1 Block DM engine failure, 1 Mars 96 SC control failure

1997 8 1 Block DM propulsion failure

1998 5 1

Last 50 Proton/Block DM: 46/50 92% (Since 10 Sep 1992)

A.2 PROTON FLIGHT HISTORY

The Proton launch vehicle is one of the most reliable commercial launch vehicles available today, with a 92 percentreliability record over the last 50 launches.

Summary launch data by year are shown in Table A.2-1.


Page A-3

Table A.2-1: Proton Launch Record Summary (1970-1998)

Year Number ofLaunches

Number of Launches by Version Total Launches onAccrual Basis

Failures

4-Stage Version 3-Stage Version

1970 7 4 3 7 1 Proton K/Block DM& 1 Proton K

1971 7 5 2 14

1972 2 1 1 16 1 Proton K

1973 7 5 2 23

1974 6 4 2 29

1975 5 5 - 34 1 Proton K/Block DM

1976 5 3 2 39

1977 6 2 2 45 1 Proton K

1978 8 7 1 53 3 Proton K/BlockDM’s

1979 6 5 1 59

1980 5 5 - 64

1981 7 6 1 71


1983 12 11 1 93

1984 13 13 _ 106

1985 10 9 1 116

1986 8 7 1 124 1 Proton K


1988 12 13 _ 150 2 Proton K/BlockDM’s

1989 11 10 1 161

1990 11 10 1 172 1 Proton K/Block DM

1991 9 8 1 181

1992 8 8 _ 189

1993 7 7 _ 196 1 Proton K/Block DM

1994 13 13 _ 209

1995 7 6 1 216


1997 8 8 - 232 1 Proton K/Block DM

1998 6 5 1 238


Page A-4

A.3 DETAILED FLIGHT HISTORY

The Proton launch history since 1970 is shown in Table A.3-1. This historical data was verified by Khrunichev inDecember 1993, and subsequent launches confirmed on a case by case basis.

Table A.3-1: Proton Launch History

Date Proton Variant Payload Orbit Type Comments

4-stage 3-stage

6 Feb 1970 ww Cosmos Failed to orbit Command abort

8 Aug 1970 ww Unknown Failed to orbit No data

12 Sep 1970 ww Luna-16 Escape

20 Oct 1970 ww Zond-8 Escape

10 Nov 1970 ww Luna-17 Escape

24 Nov 1970 ww Cosmos-379 192 km x 1210 km at 51.9 deg

2 Dec 1970 ww Cosmos-382 2464 km x 5189 km at 51.9 deg

26 Feb 1971 ww Cosmos-398 201 km x 7250 km at 51.6 deg

19 Apr 1971 ww Salyut-1 200 km x 210 km at 51.6 deg

10 May 1971 ww Cosmos-419 145 km x 159 km at 51.5 deg

19 May 1971 ww Mars-2 Escape

28 May 1971 ww Mars-3 Escape



14 Feb 1972 ww Luna-20 Escape

29 Jul 1972 ww Salyut Failed to orbit

8 Jan 1973 ww Luna-21 Escape

3 Apr 1973 ww Salyut-2 207 km x 248 km at 51.6 deg No data


21 Jul 1973 ww Mars-4 Escape

25 Jul 1973 ww Mars-5 Escape

5 Aug 1973 ww Mars-6 Escape

9 Aug 1973 ww Mars-7 Escape

26 Mar 1974 ww Cosmos-637 LEO

29 May 1974 ww Luna-22 Escape

24 Jun 1974 ww Salyut-3 LEO

29 Jul 1974 ww Molniya-1S Elliptical orbit

28 Oct 1974 ww Luna-23 Escape

26 Dec 1974 ww Salyut-4 LEO

6 Jun 1975 ww Venera-9 Earth escape

14 Jun 1975 ww Venera-10 Earth escape

8 Oct 1975 ww Cosmos-775 LEO

16 Oct 1975 ww Luna Escape

22 Dec 1975 ww Raduga-1 GSO

22 Jun 1976 ww Salyut-5 LEO

9 Aug 1976 ww Luna-24 Escape

11 Sep 1976 ww Raduga-2 GSO

26 Oct 1976 ww Ekran-1 GSO

15 Dec 1976 ww Cosmos-881 and 882 LEO


Page A-5

Table A.3-1a: Proton Launch History (Continued)


4-stage 3-stage

17 Jul 1977 ww Cosmos-929 301 km x 308 km at 51.5 deg

23 Jul 1977 ww Raduga-3 GSO

4 Aug 1977 F Cosmos Failed to orbit

20 Sep 1977 ww Ekran-2 GSO

29 Sep 1977 ww Salyut-6 380 km x 391 km at 51.6 deg

14 Oct 1977 F Unknown Failed to orbit

30 Mar 1978 Cosmos-997 and 998 230 km x 200 km at 51.6 deg

27 May 1978 F Ekran Failed to orbit First stage failure


17 Aug 1978 F Ekran Failed to orbit Second stage failure

9 Sep 1978 ww Venera-l l Escape

14 Sep 1978 ww Venera-12 Escape

17 Oct 1978 F Ekran Failed to orbit Second stage failure

19 Dec 1978 ww Gorizont-1 20,600 km x 50,960 km at 14.3 deg Failure?

21 Feb 1979 ww Ekran-3 GSO

25 Apr 1979 ww Raduga-5 GSO

22 May 1979 ww Cosmos- 1100 and 1101 193 km x 223 km at 51.6 deg

5 Jul 1979 ww Gorizont-2 GSO


28 Dec 1979 ww Gorizont-3 GSO

2 Feb 1980 ww Raduga-6 GSO

14 Jun 1980 ww Gorizont-4 GSO

15 Jul 1980 ww Ekran-5 GSO

5 Oct 1980 ww Raduga-7 GSO

26 Dec 1980 ww Ekran-6 GSO

18 Mar 1981 ww Raduga-8 GSO

25 Apr 1981 ww Cosmos-1267 240 km x 278 km at 51.5 deg

26 Jun 1981 ww Ekran-7 GSO



30 Oct 1981 ww Venera-13 Escape

4 Nov 1981 ww Venera-14 Escape

5 Feb 1982 ww Ekran-8 GSO

15 Mar 1982 ww Gorizont-5 GSO

19 Apr 1982 ww Salyut-7 473 km x 474 km at 51.6 deg

17 May 1982 ww Cosmos-1366 GSO

23 Jul 1982 F Ekran Failed to orbit First stage failure


12 Oct 1982 ww Cosmos-1413 and 1415 19,000 km x 19,000 km at 64.7 deg

20 Oct 1982 ww Gorizont-6 GSO

26 Nov 1982 ww Raduga-11 GSO24 Dec 1982 F Raduga Failed to orbit Second stage failure


Page A-6

Table A.3-1b: Proton Launch History (Continued)


4-stage 3-stage

2 Mar 1983 ww Cosmos-1443 324 km x 327 km at 51.6 deg

12 Mar 1983 ww Ekran-10 GSO

23 Mar 1983 ww Astron-1 1,950 km x 201,100 km at 51.09 deg


2 Jun 1983 ww Venera-15 Escape

6 Jun 1983 ww Venera-16 Escape


10 Aug 1983 ww Cosmos-1490 and 1492 19,000 km x 19,000 km at 64.8 deg

25 Aug 1983 ww Raduga-13 GSO

29 Sep 1983 ww Ekran-II GSO

30 Nov 1983 ww Gorizont-8 GSO

29 Dec 1983 ww Cosmos-1519 and 1521 19,000 km x 19,000 km at 64.8 deg


2 Mar 1984 ww Cosmos-1540 GSO


29 Mar 1984 ww Cosmos-1546 GSO

22 Apr 1984 ww Gorizont-9 GSO

19 May 1984 ww Cosmos-1554 and 1556 19,000 km x 19,000 km at 64.8 deg

22 Jun 1984 ww Raduga-15 GSO

1 Aug 1984 Gorizont-10 GSO

24 Aug 1984 ww Ekran-13 GSO

4 Sep 1984 ww Cosmos-1593 and 1595 19,000 km x 19,000 km at 64.8 deg

28 Sep 1984 ww Cosmos-1603 836 km x 864 km at 71 deg

15 Dec 1984 ww Vega-1 Escape

21 Dec 1984 ww Vega-2 Escape

18 Jan 1985 ww Gorizont-ll GSO

21 Feb 1985 ww Cosmos-1629 GSO


17 May 1985 ww Cosmos-1650 and 1652 19,000 km x 19,000 km at 64.8 deg


8 Aug 1985 ww Raduga-16 GSO

27 Sep 1985 ww Cosmos-1686 291 km x 312 km at 51.6 deg

25 Oct 1985 ww Cosmos-1700 GSO

15 Nov 1985 ww Raduga-17 GSO

24 Dec 1985 ww Cosmos-1710 and 1712 19,000 km x 19,000 km at 64.8 deg

17 Jan 1986 ww Raduga- 18 GSO

19 Feb 1986 ww Mir 335 km x 358 km at 51.6 deg

4 Apr 1986 ww Cosmos-1738 GSO

24 May 1986 ww Ekran-15 GSO

10 Jun 1986 Gorizont-12 GSO


Page A-7

Table A.3-1c: Proton Launch History (Continued)


4-stage 3-stage


25 Oct 1986 ww Raduga-I9 GSO


30 Jan 1987 F Cosmos-1817 188 km x 210 km at 51.6 deg Fourth stage failed toignite


31 Mar 1987 ww Kvant-1 298 km x 344 km at 51.6 deg

24 Apr 1987 F Cosmos- 1838 to 1840 200 km x 17,000 km at 64.9 deg Fourth stage earlyshutdown

11 May 1987 ww Gorizont-14 GSO

25 Jul 1987 Cosmos-1870 237 km x 249 km at 71.9 deg





26 Nov 1987 ww Cosmos-1897 GSO


27 Dec 1987 ww Ekran-17 GSO

17 Feb 1988 F Cosmos-1917P1919 162 km x 170 km at 64.8 deg Fourth stage did not ignite

31 Mar 1988 ww Gorizont-15 GSO


6 May 1988 ww Ekran-18 GSO

21 May 1988 ww Cosmos 1946-1948 19,000 km x19,000 km at 64.9 deg

7 Jul 1988 ww Phobos-1 Escape

12 Jul 1988 ww Phobos-2 Escape

1 Aug 1988 ww Cosmos-1961 GSO

18 Aug 1988 ww Gorizont-16 GSO

16 Sep 1988 ww Cosmos-1970P1972 19,000 km x 19,000 km at 64.8 deg


10 Dec 1988 ww Ekran-I9 GSO

10 Jan 1989 ww Cosmos-1987P1989 19,000 km x 19,000 km at 64.9 deg

26 Jan 1989 ww Gorizont-17 GSO


31 May 1989 ww Cosmos-2022P2024 19,000 km x 19,000 km at 64.8 deg

21 Jun 1989 ww Raduga-l-1 GSO


28 Sep 1989 ww Gorizont- 19 GSO

26 Nov 1989 ww Kvant-2 215 km x 321 km at 51.6 deg

1 Dec 1989 ww Granat 1957 km x 201,700 km at 52.1 deg

15 Dec 1989 Raduga-24 GSO

27 Dec 1989 ww Cosmos-2054 Unknown


Page A-8

Table A.3-1d: Proton Launch History (Continued)


4-stage 3-stage


19 May 1990 ww Cosmos-2079P81 19,000 km x19,000 km at 65 deg

31 May 1990 ww Kristall 383 km x 481 km at 51.6 deg

20Jun 1990 ww Gorizont-20 GSO

18 Jul 1990 ww Cosmos-2085 GSO

9 Aug 1990 F Unknown Did not achieve orbit



8 Dec 1990 ww Cosmos-2109P11 19,000 km x 19,000 km at 64.8 deg



14 Feb 1991 ww Cosmos-2133 GSO


31 Mar 1991 ww Almaz-1 268 km x 281 km at 72.7 deg

4 Apr 1991 ww Cosmos-2139P41 19,000 km x 19,000 km at 64.9 deg


13 Sep 1991 ww Cosmos-2155 GSO

23 Oct 1991 ww Gorizont-24 GSO

22 Nov 1991 ww Cosmos-2172 GSO


29 Jan 1992 ww Cosmos-2177P79 19,000 km x 19,000 km at 64.8 deg

2 Apr 1992 ww Gorizon-25 GSO

14Ju11992 ww Gorizont-26 GSO

30 Jul 1992 ww Cosmos-2204-06 19,000 km x 19,000 km at 64.8 deg

10 Sep 1992 ww Cosmos-2209 GSO


27Nov 1992 ww Gorizont-27 GSO

17 Dec 1992 ww Cosmos-2224 GSO

17 Feb 1993 ww Cosmos-223?P3? 19,000 km x 19,000 km at 64.8 deg


27 May 1993 F Gorizont Did not achieve orbit 2nd and 3rd stagepropulsion failure

30 Sep 1993 ww Gorizont GSO

28 Oct 1993 ww Gorizont GSO

18 Nov 1993 ww Gorizont GSO

23 Dec 1993 ww Gorizont GSO

20 Jan 1994 ww GALS GSO

5 Feb 1994 ww Raduga - 30 GSO

18 Feb 1994 ww Raduga- 31 GSO

11 Apr 1994 ww Glonass l9,000 km x l9,000 km at64.8°

20 May 1994 ww Gorizant GSO


Page A-9

Table A.3-1e: Proton Launch History (Continued)


4-stage 3-stage

7 Jul 1994 ww Cosmos GSO

11 Aug 1994 ww Glonass 19,000 km x 19,000 km at 64.8°

21 Sep 1994 ww Cosmos - 2291 GSO

13 Oct 1994 ww Express GSO

31 Oct 1994 ww Electro GSO

20 Nov 1994 ww Glonass 19,000 km x 19,999 km at 64.8°

16 Dec 1994 ww Luch GSO

28 Dec 1994 ww F Raduga - 32 GSO

7 Mar 1995 ww Glonass 19,000 km x 19,000 km at 64.8°

20 May 1995 ww Spektr 335 km x 358 km at 51.6°

24 Jul 1995 ww Glonass 19,000 km x 19,000 km at 64.8°

31 Aug 1995 ww Gazer GSO

11 Oct 1995 ww Looch - 1 GSO

17 Nov 1995 ww GALS GSO

14 Dec l995 ww F Glonass 19,140 km x l9,l00 km at 64.8°

25 Jan 1996 ww Gorizant GSO

19 Feb 1996 F Raduga GSO Block DM propulsionfailure

9 Apr 1996 ww Astra 1F Hi-GTO Commercial

23 Apr 1996 ww Priroda 214 km x 328 km at 51.6 deg

25 May 1996 ww Gorizant GSO

6 Sep 1996 ww Inmarsat 3 F2 GSO Commercial

26 Sep 1996 ww Express GSO

16 Nov 1996 F Mars 96 Did not achieve escape trajectory Failure of Mars 96 controlsystem to initiate BlockD2 engine ignition

24 May 1997 ww Telstar-5 Hi-GTO Commercial

6 June 1997 ww Arak GSO

18 June 1997 ww Iridium LEO Commercial

14 Aug 1997 ww Cosmos-2345 GSO

28 Aug 1997 ww PanAmSat-5 Hi-GTO Commercial

15 Sep 1997 ww Iridium LEO Commercial

3 Dec 1997 ww Astra-1G Hi-GTO Commercial

25 Dec 1997 F AsiaSat-3 GTO Block DM Engine Failure

7 Apr 1998 ww Iridium LEO Commercial


8 May 1998 ww Echostar-IV Hi-GTO Commercial

30 Aug 1998 ww Astra 2A Hi-GTO Commercial

04 Nov 1998 ww PanAmSat-8 Hi-GTO Commercial

20 Nov 1998 ww Zarya (FGB) LEO RSA/NASA


Page A-10

Notes:

a) The SL-12 launch vehicle is also designated D-l-e and is the four-stage version of the Proton.

b) The SL-13 launch vehicle is also designated D-1 and is the three-stage version of the Proton.

c) The stated orbital parameters are approximate and included for reference only.

A.4 FAILURES CAUSES AND CORRECTIVE ACTION

Data provided below was provided by Khrunichev Space Center.

a) 1970: After 128.3 seconds of flight, 1st stage engine cutoff due to false alarm from the launch vehicle safety systemactivated by the engine pressure gage. Manufacturing defect. Additional check of gages introduced at point ofinstallation.

b) 1972: After 181.9 seconds of flight, 2nd stage automated stabilization system failure due to a relay short circuit inthe "pitch" and "yawing" channels caused by elastic deformation of the device housing (which operates in vacuum).Design defect. Design of instruments upgraded and additional testing undertaken.

c) 1975: Failure of 4th stage oxidizer booster pump. Manufacturing/design defect. Cryogen-helium condensatefreezing. Booster pump blowing introduced.

d) 1977: After 40.13 seconds of flight, spontaneous cutoff of 1st stage steering motor, loss of stability and enginecutoff at 53.68 seconds into the flight safety system command. Steering most failure due to spool-an-sleeve pairmanufacturing defect (faulty liner) which caused penetration of hard particles under liner rim and resulted inspool-and-sleeve seizure.

e) 1978: After 87 seconds of flight, loss of stability commenced due to error of 1st stage second combustion chambersteering gear. High temperature impact on cables due to heptyl leak into second block engine compartment. Leaklikely developed at heptyl feed coupling to gas generator. Coupling upgraded.

f) 1978: Flight terminated after 259.1 seconds due to loss of launch vehicle stability. Automatic stabilization systemelectric circuit failure in rear compartment of 2nd stage caused by hot gases leaking from second engine gas inletdue to faulty sealing of pressure gage. Gage attaching point upgraded.

g) 1978: After 235.62 seconds of flight, 2nd stage engine shutoff and loss of stability caused by a turbine part ignitingin turbo pump gas tract followed by gas inlet destruction and hot air ejection into 2nd rear section. Engine designupgraded.

h) 1982: At 45.15 seconds into the flight, major malfunctioning of 1st stage engine fifth chamber. Flight terminatedby launch vehicle safety system command. Failure caused by steering motor malfunctioning: first stage ofhydraulic booster got out of balance coupled with booster dynamic excitation at resonance frequencies. Hydraulicbooster design redefined.

i) 1982: 2nd stage engine failure caused by high-frequency vibrations. Engine design upgraded.

j) 1986: Control system failure due to brief relay contact separation caused by engine vibration. Upgrading includedintroduction of self-latching action capability for program power distributor shaft.

k) 1987: 4th stage control system failure due to component (relay) defect. Manufacturing defect. Remedial programintroduced at supplier's factory. Inspection made more stringent.

l) 1987: 4th stage control system failure due to control system instrument defect. Manufacturing defect. Devicemanufactured at the time of transfer from developer's pilot production to a factory for full scale production.Remedial program introduced at relevant factory. No recurring failures recorded.


Page A-11

m) 1988: 3rd stage engine failure caused by destruction of fuel line leading to mixer. Unique manufacturing defect.Inventory rechecked.

n) 1988: 4th stage engine failure due to temperature rise in combustion chamber caused by penetration of foreignparticles from the fuel tank. Manufacturing defect. Remedial program introduced at point of manufacture toprevent penetration of foreign particles into tanks. No recurring failures recorded.

o) 1990: 2nd stage engine shutoff due to termination of oxidizer supply. Fuel line clogged by a piece of textile(wiping rag). Remedial program introduced to prevent wiping rags from being left inside engine and launchvehicle.

p) 1993: 2nd and 3rd stage engine failures. Multiple engine combustion chamber burn-through caused by propellantcontaminants. Remedial program introduced to modify propellant specifications and testing procedures. Alllaunch site propellant storage, transfer, and handling equipment purged and cleaned.

q) 1996: Block DM 4th stage second burn ignition failure. Remedial program involved corrective actions to preventtwo possible causes. The first involved introduction of redundant lockers, revised installation procedures, andincreased factory inspections to prevent a loosening of a tube joint causing a leak that would prevent engineignition. The second involved additional contamination control procedures to further precule particulatecontamination of the hypergolic start system.

r) 1996: Block D2 4th stage engine failure during second burn due to malfunction of Mars 96 spacecraft controlsystem, and associated improper engine command sequences. Unique configuration of spacecraft and 4th stage.Remedial program includes stringent adherence to established integration and test procedures.

s) 1997: Block DM 4th stage engine failure resulting from improperly coated turbopump seal. Remedial programincludes removal of unnecessary (for < 4 burn missions) coating.


Page B1-1

B1. QUALITY SYSTEM

TBS


Page C1-1

Introduction

This section describes the data required from the spacecraft customer to determine the compatibility of the spacecraftwith the Proton launch vehicle. Providing this data in full constitutes providing an Interface Requirements Documentwhich is a contractual document provided by the Customer at the beginning of a mission integration cycle. Forpreliminary feasibility studies, a smaller subset of this information can be provided as indicated in the table.

The requested information is provided in the sequence the data appears in the Interface Control Document in order tosimplify the creation of this document once a contract is signed.

C1. GENERAL INFORMATION

Item Parameter Units Feas. Study

Spacecraft Name

Manufacturer

Isometric view, launch configuration

Isometric view, on-orbit configuration

Estimated launch date


Page C2-2

C2. INPUT DOCUMENTS


Spacecraft DynamicModel

Per specification LKES-9704-0207

Spacecraft ThermalModel

Per specification LKET 9704-0206

Spacecraft PhysicalModel

Per specification LKEB 9801-039

Launch Operations Plan Reference Section 5.3.5 of this Mission Planner’s Guide

Environmental TestPlan

Reference Section 5.3.3 of this Mission Planner’s Guide

SC Interface ControlDrawing

Drawing based on physical model above which shows:

Locations of all critical surfaces (surfaces closer than 50 mm fromuseable volume envelope). Show on drawing and in table in SCcoordinates x,y and z.

mm

Major SC dimensions while in launch configuration mm

Complete description of any points OUTSIDE of the useable volume

Complete dimensions of interface ring with structural definition up to125 mm above interface plane including stiffness characteristics

Umbilical connector definition including position relative toseparation plane

Pneumatic connector interfaces (if applicable)

Access envelopes for accessing SC and description of how access is tobe made

Drawing to show handling fixtures and how attached to SC (hoists,slings, crossbars)


Page C3-3

C3. INTERFACES

C3.1 MECHANICAL INTERFACES


Adapter system Specify which standard adapter system. Reference Appendix D of thisMission Planner’s Guide.

# pushoff springs (if known). Reference Appendix D of this MissionPlanner’s Guide.

Specific requirements for mechanical interfaces not provided by abovestandard adapter system

Coordinate system Drawing showing spacecraft coordinate system

Drawing showing desired orientation of spacecraft coordinate systemrelative to adapter coordinate system

Description of constraints on spacecraft orientation

SC Logo Provide design

SC interface flange Cross section propertiesIxx

Iyy

Ls = 25 mm

S

mm4

mm4

mm

mm2

Scribe mark location

SC stiffness Minimum fundamental lateral and axial mode frequencies (must begreater than 10 and 25 Hz respectively

Hz

SC interface loads Confirm SC lifting device and structure can lift SC+adapter mass =200 kg

Yes or no

SC mass properties Fill in tables in attached Table C3.1-1 and C3.1-2.

Fairing access doors Location required for access SCcoordinates

Method of access required

Time when access required

Fitcheck/ shocktest Confirm fitcheck and shock test requirements

SC Pendulum model Provide pendulum model of SC during powered flight perFigure C3.1-1

SC Slosh model Provide slosh model of SC during ballistic flight and at separation byproviding parameters in Figure C3.1-2.

Propellant tank Provide general propellant tank geometry per Figure C3.1-3


Page C3-4

C3.2 ELECTRICAL INTERFACES


Electrical connector Confirm type of connectorsLVIJ1LVIJ2Confirm location and keyingLVIJ1LVIJ2Confirm quantities to be suppliedLVIJ1LVIJ2LVIP1LVIP2Provide pin locations of SC separation jumper loopsLVIJ1 pins:LVIJ2 pins:Current flow 20 seconds or less prior to separation. mA (must

be lessthan 100)

Dry loop commands Dry loop commands required? Yes/noCharacteristics if yes:# commandsDesired pin locationsLVIJ1 pins:LVIJ2 pins:CurrentVoltagePulse durationTime during flight

mAVmssec priortoseparation

SC telemetry processing LV processing of SC telemetry required? Yes/noCharacteristics of if yes:Desired pin locationsLVIJ1 pins:LVIJ2 pins:VoltageCurrentData rate

VmAHz

Current throughumbilicals at liftoff

Current flow 5 minutes or less prior to liftoff. mA (mustbe lessthan 100)

SC RF characteristics Fill in table in attached Table C3.2-1.SC omni antenna Provide location of omni antennas to be used on pad, including:

Antenna pattern showing antenna origin in SC coordinates, -3dbbeamwidth

Umbilical wiring Fill out Table C3.2-2 with pin assignments and desired linecharacteristics


Page C3-5

C3.3 ENVIRONMENTAL INTERFACES


Thermal requirements Provide any particular ground thermal requirements

Provide any particular thermal requirements during ascent or parkingorbit

Provide any particular thermal requirements during transfer orbitincluding:

sun angle vs. time

Provide any particular thermal requirements during final injectionorbit

Venting analysis Provide archimedes volume of SC in launch configuration (forventing analysis)

SC Testing Provide confirmation of compliance with Planner’s Guide testrequirements for acoustic, sine, static testing and for thefitcheck/shocktest

EMC Provide confirmation of compliance with EM Susceptibility curve inLV in Planner’s Guide

Provide confirmation of acceptability of LV radiated emissions curvein Planner’s Guide

Humidity Provide humidity requirements for ground transportation

Provide humidity requirements for processing facility


Page C3-6

C3.4 FLIGHT DESIGN


Parking Orbit Define thermal conditioning maneuvers required

Define sun angle constraints

Transfer Orbit Define thermal conditioning maneuvers required

Define sun angle constraints

Injection Orbit Target orbit for performance calculation:

Injection mass

Perigee

Apogee

Inclination

Argument of Perigee

Longitude of Ascending Node

kg

km

km

deg.

deg.

deg.

Launch Window Define constraints on launch window

Provide target launch date

Provide launch window at perigee passage for one year covering thecontractual launch date. Include open and close times in GMT foreach day.

Separation Define type of separation 3 axisstabilized,spinningortransversespin

Define desired separation attitude with a diagram showing thepointing vector in SC coordinates and the pointing vector relative torelative right ascension and declination as defined in the Planner’sGuide.

Define desired spin rate about each SC axis.

Define desired spin rate tolerance about each SC axis

deg.

deg.

Define desired separation velocity m/sec.


Page C3-7

C3.5 OPERATIONS


EGSE Fill out table in Table C3.5-1 for all EGSE to be used at Launch Site

Fluid and gases Fill out table in Table C3.5-2 for quantities and types of fluids andgases

Contamination Control Provide any special requirements

Campaign Support Provide list of support required in each area (if nonstandard):

Area 92A-50

Hall 102

Hall 101

Hall 103A

Control Rm

Hall 103

Offices

Area 31

Rm 100C

Rm 100B

Rm 100A

Rm 119

Offices

Launch Complex 81

Bunker

Vault

Pad

Bldg 92-1


Page C3-8

C3.6 OPERATIONS (Continued)


Transportation Provide description of all items including propellant to be shippedincluding:

Item nameQuantityWeightTiedown methodStorage requirements

kg

Handling Provide description of items which require physical handling at thelaunch site including:Equipment name/locationDimensionsWeightHandling method

mkg

Communications Define number and type of international lines required. Up to 3 64kbps lines can be provided.

If multiplexer provided, provide characteristics of and desiredlocation.

Provide locations of all intersite data transmissions including data typeand rate and interface (analog modem, V.35 or RS232 interface)

Provide requirements for hardline data transmission between Area 81Vault and Bunker, Bldg 44 Rm 58 and Bldg 40D Rm 120 and/orbetween Bldg 40D Rm 301 and Room 120.

Ground ElectricalInterfaces

Provide block diagrams of desired umbilical interfaces between SCand EGSE while being processed and while on the pad

Feedthroughs Provide description of feedthroughs required between Control Rm of92A-50 and Hall 103A or between Rm 58 of Bldg 44 and Hall 1,including:

Feedthrough designation

Cable designation

Cable connector dia.

Cable dia.

mm

mm


Page C3-9

Table C3.1-1: SC Mass properties

SC Mass Properties are shown with a normal distribution. Table C3.1-1c provides the dry spacecraft mass properties tobe used for analysis of separation dynamics taking into account fluid sloshing effects.

Approximately TBD kg of helium gas pressurant is included in the full-up spacecraft mass.

Table C3.1-1a: SC Mass Properties Near 0g

Mass (kg) Center of Gravity location(spacecraft coordinates, mm)

Moments of inertia relative to spacecraft

c.g.(kg-M2)

CGX CGY CGZ IXX IYY IZZ PXY PXZ PYZ

Nominal TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD

+/- + TBD + TBD + TBD + TBD TBD% TBD% TBD% TBD% TBD% TBD%

Table C3.1-1b: SC Mass Properties Near 1g



c.g.(kg-M2)

CGX CGY CGZ IXX IYY IZZ PXY PXZ PYZ


+/- + TBD + TBD + TBD + TBD TBD% TBD% TBD% TBD% TBD% TBD%

Note:

a) Maximum to minimum inertia ratio is greater than or equal to 1.02

b) Z coordinate relative to separation plane

c) Maximum required tolerance on the final weight before launch =+0/- TBD kg. and will be based on the SC manufacturersfinal mass properties report

d) Above data based on the SC manufacturers Mass Properties Report dated TBD.


Page C3-10

Table C3.1-1c: SC Mass Properties(Dry Spacecraft)



c.g.(kg-M2)

CGX CGY CGZ IXX IYY IZZ IXY IXZ IYZ


+/- + TBD + TBD +TBD + TBD TBD % TBD % TBD % TBD% TBD% TBD%

Notes:

a) Z coordinate relative to separation plane

b) Above data based on the SC manufacturers Mass Properties Report dated TBD.


Page C3-11

Table C3.1-2: Description Of Liquid Masses

These tables provide the mass properties for the individual tanks for the nominal propellant load of TBD% fill fraction(full tanks) being assumed, in a near 0 g field in a 1 g field and during transportation.

The associated tank geometry is shown in Figure C3.1-3.

a) near 0 g (0.125 g)

Mass(kg)

Center of Gravity location(spacecraft coordinates, mm)

Moments of inertia relative to propellant c.g.

(kg-M2)

CGX CGY CGZ* IXX IYY IZZ IXY IXZ IYZ

Ox TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD

+/- TBD TBD TBD TBD TBD% TBD % TBD % TBD % TBD % TBD %

Fuel TBD TBD TBD TBD TBD TBD TBD TBD TBD TBD

+/- TBD TBD TBD TBD TBD% TBD % TBD % TBD % TBD % TBD %

b) 1g

Mass(kg)



(kg-M2)



+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD %



Note: Z coordinate relative to separation plane

c) 1g (During Transportation)

Mass(kg)



(kg-M2)






Note: Z coordinate relative to separation plane


Page C3-12

Figure C3.1-1: SC Pendulum Model

L

ms

as

xs

Pendulum Model Parameters:

1. 0g

Tank ms, Kg L, m xs, m as, m , % m0, Kg x0, m

F

O

2. 1g

Tank ms, Kg L, m xs, m as, m , % m0, Kg x0, m

F

O

Filling Parameters: (t = 20OC)

Tank Mass, Kg Level, m Filling, % Density, Kg/m3 Kinematic viscosity, m2/s

F

O

as - Pivot point of the pendulum from the tank bottom, mL - Length of the pendulum, m;xs - Slosh mass location from the tank bottom, m;ms - Mass of the pendulum, Kg; - Damping, %.m0 - Mass of fixed liquid, kg;x0 - Coordinate of mass m0 from the tank bottom, m.


Page C3-13

Figure C3.1-2: SC Slosh Properties During Ballistic Flight and at Separation

gI dt

dω + ×CM g

2

2

dt

ldT - ×ω ( )ωgI ( )

iCToii

FR ⋅×∑

- ×CM g dt

dω+ gM

2

2

dt

ld = F - ×ω ( )CM g×ω -

iCTi

F ⋅∑

2

2

dt

ad i = -

dt

dω ia× - 2 ×ω

dt

ad i - ×ω ( ×ω ia ) - iiCTii m/FAG ⋅+

iA = 2

2

dt

ld+

dt

dωoi

r× + ×ω ( )oi

r×ω

sg MM = + ( )iii

G1m −∑

CM g = ss rM + ( )ii

i

G1m −∑ oi

r

sg II = - ( )iii

G1m −∑ ( )oi

rS ( )oi

rS

( )oi

rS = S

iziyix

=

−

−

−

0ixiyix0iziyiz0

i

i

oioi G

arR +=

Gi = )1(1

)1(2

i

i

K

K

−+−

iCTF ⋅ =

−−−tishearpidemipr VkVkdk

0

ii

ii

Rr

Rr

≥

<

ri=( )( )31 11

1

4

3− ÷−

iRiH

iK( )a

a

MINIMUMiβ

]5.0;5.0[ +−∈• ( )

+

−−=

0

0

1

aHK1

Kaa i

i

ii

iβ

( )( )a

add

i

ii0i

ββ

−= ( )

+−

−−= −

3

i

ii

i

iiii0

1R4

H31K1

K

K1rRd

i

ii

ii

Pi ddddt

add

V ••

•

= I Pii

t Vdt

adV i −=

mNk shear/170229= )( sec/16256 mNk shear

= )( sec/3.354 mNk shear=


Page C3-14

NOMENCLATURE

ω the spacecraft angular rate vector;

l the vector determining the position of the coordinate origin fixed in the spacecraft body in the inertialcoordinate system;

C the vector determining the spacecraft mass center position;

T the sum of external moments with respect to the selected origin of coordinates;

F- the sum of external forces acting on the spacecraft;

ai position vector of fluid center of mass of i-th tank about center of this tank;

Vpi,Vtithe radial and tangential components of velocity

dt

dai of the liquid mass center in the i - tank;

Ms, Is the mass and matrix of inertia of the dry spacecraft (without liquid components);

Ri the i - tank radius;

rsthe vector determining the dry spacecraft mass center position in the i - tank with respect to the origin of

coordinates;

mi the liquid mass in the i - tank;

roithe vector drawn from the origin of coordinates to the i - tank center;

Roi the vector determining the liquid mass center position in the i - tank with respect to the origin of coordinates;

Aithe vector of inertial acceleration of the i - tank;

Ki the i - tank filling factor

Ki=V filling

V total

where: V filling is the volume of poured fuel component;V total is the total volume of the tank.

Hi the i - tank cylinder length

Hi=0 corresponds to spherical tank


Page C3-15

Figure C3.1-3: Propellant Tank Geometry Required Data

TBD

r T B D

-Xsc

T B D

T B D

TBDSC/Upper Stage Separation Plane

TBD

Liquid Level


Page C3-16

Table C3.2-1: SC RF Characteristics

Transmitter 1 Transmitter 2 Receiver 1 Receiver 2

Description TM Transmitter 1 TM Transmitter 2 Command Receiver1A

Command Receiver1B

Carrier Frequency (Ghz) TBD TBD TBD TBD

3 db Bandwidth (Mhz) TBD TBD TBD TBD

20 db Bandwidth (Mhz) [TBD] [TBD] TBS TBS

60 db Bandwidth (Mhz) TBD TBD TBD TBD

Modulation type andcharacteristics

TBD TBD TBD TBD

Intermediate frequency (Mhz) TBD TBD

Local Oscillator frequencies(Ghz)

TBD TBD

Transmit antenna output powerEIRP (dbW)

Max (on-orbit setting)

Max (normal on pad setting)

Nom

Min

TBD

TBD

TBD

TBD

TBD

TBD

TBD

TBD

Receive antenna receive flux

density (dbW/M2)

Max

Nom

Min

TBD

TBD

TBD

TBD

TBD

TBD

Antenna description, polarization,pattern

TBD(c) TBD(c) TBD(c) TBD(c)

Antenna location

Operating on pad? TBD (e) TBD (e) TBD TBD

Operating in flight? TBD TBD TBD TBD

Ground equipment output power(dbm)

Max

Nom

Min

TBD

TBD

TBD

TBD

TBD

TBD

Ground equipment receive power(dbm)

Max (see note f)

Nom

Min

TBD

TBD

TBD

TBD

TBD

TBD


Page C3-17

Notes: [Sample]

a) TBD RF command links and TBD RF telemetry links are required. The SC Checkout Station will have 1 physical commandinterface and 1 physical telemetry interface.

b) Transmit and receive frequencies at the RF Console are the same as those indicated in the above table.

c) Vertical polarization (E-Field parallel to spacecraft Y axis), H: horizontal polarization (E-Field parallel to spacecraft X axis).

d) SC-TLM amplifiers are TBD.

e) SC Telemetry Transmitters are on and amplifiers are off during all on-pad operations.

f) Ground equipment can accommodate TBD dBm power levels without damage.

g) SC GSE RF interface impedance shall be 50 ohm

h) Uninterrupted operation of RF devices shall not exceed 8 hours, with a 30 minute break before the next 8 hour session.

i) The SC manufacturer shall provide to Khrunichev the measured coefficient values for TT &C signals via the RF windowobtained during the RF channel checkup in the integration facility following the Ascent Unit encapsulation.

j) Prior to installation of the LV+Ascent Unit on the pad and following the delivery of the STE to the bunker, the SCmanufacturer shall verify continuity between command RF link and STE and issue to Khrunichev the Certificate of LaunchPad Readiness to accommodate the LV and Ascent Unit. The spectrum analyzer to be provided by the SC manufacturer shallbe adapted to 220 V 50 Hz.

k) After installation of LV+Ascent Unit and prior to the roll-up of service tower, the SC manufacturer, in conjunction withKhrunichev, shall checkup the RF link between the Ascent Unit and STE. Such checkup shall be performed 20 minutes afterthe mating of the LV aft section. Loral to confirm functionality of the RF link within 45 minutes.

l) At L-6 months, the SC manufacturer shall provide to Khrunichev two connectors for installation by Khrunichev on theexisting RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructions on cable dressingand cable performances. (If required)


Page C3-18

Table C3.2-2a: LVIJ1 Umbilical Pin Assignments

ConnectorPin

Function Max.Voltage(V)*

Max.Current(A)

Max.Voltage atLiftoff (V)

Max.Current atLiftoff (A)

LineResistance(ohms) (4)

Time ofUsage

123456789101112131415161718192021222324252627282930313233343536373839404142434445


Page C3-19

ConnectorPin


Max.Current(A)




Time ofUsage

46474849505152535455565758596061

Notes:

a) For the bus power lines, the maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.*Maximum voltage measured at the M&C interface at the vault

b) SC CONTRACTOR equipment provides protection against exceeding 100 V at spacecraft umbilical interface. It also providescontinuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power throughthe umbilicals will be automatically shutoff within 0.2 sec if max current is exceeded by 50%.

c) Separation jumpers are configured as follows:

1) Pins TBD and TBD jumpered on LV side


3) Pins TBD and TBD jumpered on SC side

d) Indicated resistance values are from LVIP1/P2 IFD connection to the KhSC/SC CONTRACTOR EGSE interface in thevault room (or on the Mobile Service Tower for designated circuits).

e) Some circuits (e.g. battery charging circuits) may be terminated at the LV Mobile Service Tower if necessary


Page C3-20

Table C3.2.2b: LVIJ2 Umbilical Pin Assignments

ConnectorPin


Max.Current(A)




Time ofUsage

123456789101112131415161718192021222324252627282930313233343536373839404142434445


Page C3-21

ConnectorPin


Max.Current(A)




Time ofUsage

46474849505152535455565758596061

Notes:

a) For the bus power lines, the maximum voltage at the P1 and P2 spacecraft umbilical connectors is 100 V.* Maximum voltage measured at the M&C interface in the vault

b) SC Contractor equipment provides protection against exceeding 100 V at spacecraft umbilical interface. It also providescontinuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power throughthe umbilicals will be automatically shutoff within 0.2 sec if max current is exceeded by 50%.

c) Separation jumpers are configured as follows:



3) Pins TBD and TBD jumpered on SC side

d) Indicated resistance values are from LVIP1/P2 IFD connection to the KhSC/SC CONTRACTOR EGSE interface in thevault room (or on the Mobile Service Tower for designated circuits).

e) Some circuits (e.g. battery charging circuits) may be terminated at the LV Mobile Service Tower if necessary.


Page C3-22

Table C3.5-1: EGSE Description

Connectors

Equipment Power Source Power Required Heat Output Equipment Plug side Facility side

BLDG. 92A-50

Hall 101

Hall 103

Hall 103A

Control Room

Office Areas


Page C3-23

Table C3.5.1-1: SC Contractor Electrical Ground Support Equipment (Continued)

Connectors


AREA 31

Room 100A

Room 100B

Room 119

Room 120

Room 121

Offices

Bldg 44 Hall 1


Page C3-24

Table C3.5.1-1: SC Contractor Electrical Ground Support Equipment (Continued)

Connectors


Bldg 44 Rm 58

BLDG. 92-1

LAUNCHCOMPLEX 81

Bunker (Rm. 250)

Bunker (Rm 251)

Bunker (Rm 244)

Vault (Rm. 64 or 76))


Page C3-25

Table C3.5-2: Fluids/Gases Requirements

Name Conditions Supplied By Location of Use

Compressed Air for Breathing(SCAPE)

See Launch Campaign Guide for max available KhSC/ILS Fueling Hall

distilled water See Launch Campaign Guide for max available. KhSC/ILS Processing HallFueling Hall

demineralized water See Launch Campaign Guide for max available KhSC/ILS DecontaminationArea

Nitrogen GOST-92-93-74,technical grade 1

See Launch Campaign Guide for max available. KhSC/ILS DecontaminationArea

Nitrogen GOST-92-93-74,technical grade 1

See Launch Campaign Guide for max available KhSC/ILS Fueling Hall

Liquid Nitrogen (LN2) TBD liters KhSC/ILS Fueling Hall

He (Ghe) per spec Mil-p-27407Type 1, Grade A

TBD K-bottles high pres (400 bar) TBD K-bottles low pres (135 bar)

SC contractor Fueling Hall

MMH TBD Cylinders - Tot weight ea TBD kg max,TBD kg max prop weight


Nitrogen TBD SC contractor Fueling Hall

Nitrogen Tetroxide TBD Cylinders - Tot weight ea TBD kg max,TBD kg max prop weight


Shop Air See Launch Campaign Guide for max available KhSC/ILS Fueling Hall

Ethyl Alcohol See Launch Campaign Guide for max available. KhSC/ILS Fueling Hall

Service Water See Launch Campaign Guide for max available KhSC/ILS Fueling Hall

Grade “Extra” or Highest GradeGOST 18300-87 Ethyl Alcohol

See Launch Campaign Guide for max available. KhSC/ILS Fueling Hall

Freon TBD SC contractor

Argon TBD SC contractor

IPA TBD SC contractor


Page D1-1

D1. 1194AX-500 Adapter System

D1.1 INTRODUCTION

The 1194AX-500 adapter system is comprised of the 1194 AX clamp band system, separation springs, a payloadadapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interfacecharacteristics, structural capability, usable volume, and accelerometer installation.

D1.2 MECHANICAL INTERFACE AND STRUCTURAL CAPABILITY

The 1194AX-500 adapter is a one piece conic frustum of monocoque construction with a height of 500 mm andfabricated from aluminum alloy. Mechanical drawings for the 1194AX-500 adapter are presented in Section D1.7 ofthis Appendix.

D1.2.1 Interface Ring Characteristics

The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a loadpath between the spacecraft and adapter during ground operations and flight. The outboard features of the combinedcross section are designed to interface with the separation system clamp band. The cross section and material propertycharacteristics for the spacecraft and adapter interface ring are presented in Table D1.2.1-1 and FigureD1.2.1-1. The dimensions of the spacecraft interface are presented in Section D1.7 of this Appendix.

Table D1.2.1-1: Spacecraft and Adapter Interface Ring Characteristics

Ring Characteristics Spacecraft Ring Adapter Ring

Height Of Effective Cross Section (L) 25 mm 25 mm

Cross Section Area (A) 460 mm2 ± 15% 558 mm2

Moment Of Inertia (Ixx) 52000 mm4 ± 15% 72000 mm4

Moment Of Inertia (Iyy) 13400 mm4 ± 15% 30000 mm4

Young Modulus (E) 69 x 103 MPa 70 x 103 MPa


Page D1-2

Figure D1.2.1-1: Spacecraft and Adapter Interface Ring Cross Section

y

S e p a r a t i o n P l a n e


LA d a p t e r R i n g

y

yy

x

x

x

x

S p a c e c r a f t R i n g


Page D1-3

Scribe marks on the spacecraft interface ring verify proper alignment of the spacecraft relative to the launch vehicle.The attributes and location of the scribe mark for the spacecraft and adapter ring are presented in Section D1.7 of thisAppendix.

D1.2.2 Structural Capability

The structural capability of each adapter system is based on the allowable line load obtained from testing. Therelationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1194AX-500 adapter system ispresented in Figure D1.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffnesscharacteristics presented in Table D1.2-1, the geometry presented in Section D1.7 of this Appendix and the quasi-static design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculatedby classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reducethe allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as pointloads through the spacecraft structure close to the interface.

The structural capability presented should only be used as a guideline for assessment of interface structuralcompatibility. Coupled loads analysis performed early in the mission integration will verify margins for structuralloading of the adapter and separation system.


Page D1-4

Figure D1.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC)

SC Mass(kg) 40kN Tension 26.6kN Tension2000 3.07 2.302200 2.79 2.092400 2.56 1.922600 2.37 1.782800 2.20 1.653000 2.06 1.553200 1.93 1.453400 1.82 1.373600 1.72 1.293800 1.63 1.234000 1.55 1.174200 1.48 1.114400 1.42 1.07 Adapter system capability based on the following data:

4600 1.36 1.02 Allowable Limit Line Loads

4800 1.30 0.98 Tension (26.6 kN clamp band tension) Nt = 91 N/mm

5000 1.25 0.94 Tension (40 kN clamp band tension) Nt = 122 N/mm

5200 1.19 0.89 Compression Nc = 156 N/mm

5400 1.14 0.84 Quasi Static Loads

5600 1.08 0.79 Per Figure 3.4.1.2-1 of this Mission Planner's Guide

Allowable C.G. Offset (m)

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600

SC Mass (kg)

Long

itudi

nal O

ffset

of S

C C

.G. f

rom

Sep

arat

ion

Pla

ne

Preliminary

40kN Clamp Band Tension

26.6kN Clamp Band Tension


Page D1-5

D1.3 USABLE VOLUME

The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-500 adapter system ispresented in Figures D1.3-1a through D1.3-1e. The spacecraft static envelope (maximum dimensions of unloadedspacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond theuseable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacementsdue to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DMmay protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements willnot exceed 50 mm. This must be verified by coupled loads analysis.


Page D1-6

Figure D1.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter


Page D1-7

Figure D1.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 1 of 4)


Page D1-8

Figure D1.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 2 of 4)


Page D1-9

Figure D1.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 3 of 4)


Page D1-10

Figure D1.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 4 of 4)


Page D1-11

D1.4 SEPARATION SYSTEM

The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp bandare preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preloadis 40.0 kN for the1194AX adapter system. The separation system is released when these bolts are severed by boltcutters. Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and theretention device moves the clamp band away from the interface ring. The retention device also secures the band to theadapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft andadapter.

The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planner’s Guide.

To ensure proper separation velocity, a matched set of separation springs are provided. Each separation springs has aninitial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring requiredto provide the desired separation velocity and separation rates. Table D1.4-1 defines the maximum and minimumspring stroke range with the associated spring force characteristics. Spring sets can include any number between twoand twelve.

Table D1.4-1: Separation Spring Characteristics

Stroke

(mm)

Initial Force

(N)

Final Force

(N)

Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20

Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20

The separation event is affected by interface hardware that impart force during separation. This hardware consists ofelectrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electricaldisconnects and grounding connectors are provided as shown in Section D1.7 of this Appendix. This symmetricalarrangement is provided to minimize overturning moments at separation. The force profile for each electricaldisconnects is shown in Figure D1.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resistsseparation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D1.4-2.


Page D1-12

Figure D1.4-1: Electrical Disconnect Force Profile for a Single 37 Pin Electrical Connector

Force Displacement(N) (mm)178 0156 3.4-115 3.4-115 6.7

0 6.7

-150

-100

-50

0

50

100

150

200

0 1 2 3 4 5 6 7 8 9 10

Displacement (mm)

Fo

rce

(N)

Note: A positive force assists separation.


Page D1-13

Figure D1.4-2: Purge Connector Force Diagram (Typical)

0

20

40

60

80

100

0.00 0.10 0.20 0.30 0.40 0.50 0.60

Axial Travel (mm)

Fo

rce

(N)


Page D1-14

D1.5 ELECTRICAL INTERFACE

In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. Theseelectrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment andthe launch vehicle. The standard adapter system includes two diametrically opposed electrical rise off disconnects.There are two standard configurations for the electrical interface. Details for each of the two standard electricalinterfaces configurations are presented in Section D1.7 of this Appendix. Connectors conforming to MIL-C-81703have been used for both adapter system configurations. Part numbers for the standard 37 pin connectors are presentedin Table D1.5-1.

Table D1.5-1: Standard Electrical Connectors

Connector ID No. of Pins Deutsch Part No.

Spacecraft Side LVIP1 37 MS3446E37-50P

LVIP2 37 MS3446E37-50P

Launch Vehicle Side LVIJ1 37 MS3464E37-50S

LVIJ2 37 MS3464E37-50S

For each electrical connector, two pins and a loop back on the spacecraft side of the interface are required for launchvehicle separation indicators. Loop backs on the launch vehicle side of the interface for indication of separation for thespacecraft can be provided as required.

Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electricalconnectors and GSE. This information includes available wire types, shielding, voltage requirements, currentrequirement, and resistance requirements.

The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuityacross the separation interface. Electrical continuity across the separation interface is provided by two diametricallyopposed electrical grounding connectors as presented in Section D1.7 of this Appendix.

D1.6 INSTRUMENTATION

Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. Thestandard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented tomonitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D1.7 ofthis Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels forthe Proton mission are presented in Section 4.2.1.7 of the Mission Planner’s Guide.


Page D1-15

D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS


Page D1-16

1194AX-500 Adapter Mechanical Drawings (Continued 2 of 9)

TBS


Page D1-17


TBS


Page D1-18


TBS


Page D1-19


TBS


Page D1-20


TBS


Page D1-21


TBS


Page D1-22


TBS


Page D1-23


TBS


Page D2-1

D2. 1194AX-625 ADAPTER SYSTEM

D2.1 INTRODUCTION

The 1194AX-625 adapter system is comprised of the 1194 AX clamp band system, separation springs, a payloadadapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interfacecharacteristics, structural capability, usable volume, and accelerometer installation.


The 1194AX-625 adapter is a one piece conic frustum of monocoque construction 625 mm in height and fabricatedfrom aluminum alloy. Mechanical drawings for the 1194AX-625 adapter are presented in Section D2.7 of thisAppendix.





Height Of Effective Cross Section(L) 25 mm 25 mm






Page D2-2


y




y

yy

x

x

x

x



Page D2-3



The structural capability of each adapter system is based on the allowable line load obtained from testing. Therelationship between spacecraft mass and longitudinal center of gravity (C.G.) for the 1194AX-625 adapter system ispresented in Figure D2.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffnesscharacteristics presented in Table D2.2-1, the geometry presented in Section D2.7 of this Appendix and the quasi-static design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculatedby classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reducethe allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as pointloads through the spacecraft structure close to the interface.



Page D2-4

Figure D2.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC)

SC Mass(kg) 40kN Tension 26.6kN Tension2000 3.07 2.302200 2.79 2.092400 2.56 1.922600 2.37 1.782800 2.20 1.653000 2.06 1.553200 1.93 1.453400 1.82 1.373600 1.72 1.293800 1.63 1.234000 1.55 1.174200 1.48 1.114400 1.42 1.07 Adapter system capability based on the following data:

4600 1.36 1.02 Allowable Limit Line Loads

4800 1.30 0.98 Tension (26.6 kN clamp band tension) Nt = 91 N/mm

5000 1.25 0.94 Tension (40 kN clamp band tension) Nt = 122 N/mm

5200 1.19 0.89 Compression Nc = 156 N/mm

5400 1.14 0.84 Quasi Static Loads

5600 1.08 0.79 Per Figure 3.4.1.2-1 of this Mission Planner's Guide

Allowable C.G. Offset (m)

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600

SC Mass (kg)

Long

itudi

nal O

ffset

of S

C C

.G. f

rom

Sep

arat

ion

Pla

ne

Preliminary

40kN Clamp Band Tension

26.6kN Clamp Band Tension


Page D2-5

D2.3 USABLE VOLUME

The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-625 adapter system ispresented in Figures D2.3-1a through D2.3-3e for the Block DM and Breeze M versions of the LV. The spacecraftstatic envelope (maximum dimensions of unloaded spacecraft, including manufacturing tolerances and expansion ofthermal blankets) must not protrude beyond the useable volume, except where it is mutually agreed upon by ILS andKhrunichev. Spacecraft dynamic displacements due to ground or flight loads and deviations caused by an imperfectinstallation of the spacecraft on the Fourth Stage may protrude beyond the boundaries of this useable volume. It isassumed that spacecraft dynamic displacements will not exceed 50 mm. This must be verified by coupled loadsanalysis.


Page D2-6

Figure D2.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter


Page D2-7

Figure D2.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4)


Page D2-8

Figure D2.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4)


Page D2-9

Figure D2.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4)


Page D2-10

Figure D2.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4)


Page D2-11

Figure D2.3-2a: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter


Page D2-12

Figure D2.3-2b: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC)


Page D2-13

Figure D2.3-2c: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC)


Page D2-14

Figure D2.3-2d: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC)


Page D2-15

Figure D2.3-2e: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC)

Proton Mission Planners Guide, LKEB-9812-1990Issue 1, Revision 4, March 1, 1999

Page D2-16

Figure D2.3-3a: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter


Page D2-17

Figure D2.3-3b: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC)


Page D2-18

Figure D2.3-3c: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC)


Page D2-19

Figure D2.3-3d: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC)


Page D2-20

Figure D2.3-3e: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC)


Page D2-21


The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp bandare preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preloadis 26.6 or 40.0 kN for the 1194AX adapter system. The separation system is released when these bolts are severed bybolt cutters. Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and theretention device moves the clamp band away from the interface ring. The retention device also secures the band to theadapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft andadapter.

The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planners Guide.



Stroke

(mm)

Initial Force

(N)

Final Force

(N)

Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20

Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20

The separation event is affected by interface hardware that impart force during separation. This hardware consists ofelectrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electricaldisconnects and grounding connectors are provided as shown in Section D2.7 of this Appendix. This symmetricalarrangement is provided to minimize overturning moments at separation. The force profile for each electricaldisconnects is shown in Figure D2.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resistsseparation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D2.4-2.


Page D2-22


Force Displacement(N) (mm)178 0156 3.4-115 3.4-115 6.7

0 6.7

-150

-100

-50

0

50

100

150

200

0 1 2 3 4 5 6 7 8 9 10

Displacement (mm)

Fo

rce

(N)



Page D2-23

Figure D2.4-2: Purge Connector Force Diagram (Typical)

0

20

40

60

80

100

0.00 0.10 0.20 0.30 0.40 0.50 0.60

Axial Travel (mm)

Fo

rce

(N)


Page D2-24





Spacecraft Side LVIP1 37 MS3446E37-50P

LVIP2 37 MS3446E37-50P

Launch Vehicle Side LVIJ1 37 MS3464E37-50S

LVIJ2 37 MS3464E37-50S


Refer to Section 4.2.2 of the Mission Planners Guide for information on the electrical wiring between the electricalconnectors and GSE. This information includes available wire types, shielding, voltage requirements, currentrequirement, and resistance requirements.

The requirement for maximum resistance across the separation interface is of 10 milliohms to ensure electricalcontinuity across the separation interface. Electrical continuity across the separation interface is provided by twodiametrically opposed electrical grounding connector as presented in Section D2.7 of this Appendix.


Accelerometers are included in the standard adapter system to monitor spacecraft mechanical environments. Thestandard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented tomonitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D2.7 ofthis Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels forthe Proton mission are presented in Section 4.2.1.7 of the Mission Planners Guide.


Page D2-25

D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS


Page D2-26

1194AX-625 Adapter Mechanical Drawings(Sheet 2 of 9)

TBS


Page D2-27


TBS


Page D2-28


TBS


Page D2-29


TBS


Page D2-30


TBS


Page D2-31


TBS


Page D2-32


TBS


Page D2-33


TBS


Page D3-1

D3. 1666V-1000 ADAPTER SYSTEM

D3.1 INTRODUCTION

The 1666V-1000 adapter system is comprised of the 1666 V clamp band system, separation springs, a payload adapterand electrical rise off disconnects. This appendix defines the mechanical and electrical interface characteristics,structural capability, usable volume, and accelerometer installation.


The 1666V-1000 adapter is a two piece conic frustum of monocoque construction 1000 mm in height and fabricatedfrom aluminum alloy. Mechanical drawings for the 1666V-1000 adapter are presented in Section D3.7 of thisAppendix.











Page D3-2


y




y

yy

x

x

x

x



Page D3-3



The structural capability of each adapter system is based on the allowable line load obtained from testing. Therelationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666V-1000 adapter system ispresented in Figure D3.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffnesscharacteristics presented in Table D3.2-1, the geometry presented in Section D4.7 of this Appendix and the quasi-static design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculatedby classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reducethe allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as pointloads through the spacecraft structure close to the interface.



Page D3-4

Figure D3.2.2-1: Capability of 1666V Adapter System - SC Mass vs Longitudinal C.G. (TBC)

Allowable C.G.SC Mass Offset (m)

(kg) 30.0kN Tension2000 3.392200 3.092400 2.832600 2.622800 2.443000 2.283200 2.143400 2.023600 1.893800 1.774000 1.66 Note: A positive force assists separation.

4200 1.564400 1.474600 1.39 Adapter system capability based on the following data:

4800 1.32 Allowable Limit Line Loads

5000 1.25 Tension Nt = 69 N/mm

5200 1.18 Compression Nc = 86 N/mm

5400 1.12 Quasi Static Loads

5600 1.07 Per Figure 3.4.1.2-1 of this Mission Planner's Guide

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600

SC Mass (kg)

Long

itudi

nal O

ffset

of S

C C

.G. f

rom

Sep

arat

ion

Plan

e

Preliminary


Page D3-5

D3.3 USABLE VOLUME

The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666V-1000 adapter system ispresented in Figures D3.3-1a through D3.3-1e. The spacecraft static envelope (maximum dimensions of unloadedspacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond theuseable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacementsdue to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DMmay protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements willnot exceed 50 mm. This must be verified by coupled loads analysis.


Page D3-6

Figure D3.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter


Page D3-7

Figure D3.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 1 of 4)


Page D3-8

Figure D3.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 2 of 4)


Page D3-9

Figure D3.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 3 of 4)


Page D3-10

Figure D3.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 4 of 4)


Page D3-11


The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp bandare preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preloadis 30.0 kN for the1666 V adapter system. The separation system is released when these bolts are severed by bolt cutters.Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the retentiondevice moves the clamp band away from the interface ring. The retention device also secures the band to the adapterand prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter.




Stroke

(mm)

Initial Force

(N)

Final Force

(N)

Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20

Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20

The separation event is affected by interface hardware that impart force during separation. This hardware consists ofelectrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects and groundingconnectors are provided pairs as shown in Section D3.7 of this Appendix. This symmetrical arrangement is provided tominimize overturning moments at separation. The force profile for each electrical disconnects is shown in FigureD3.4-1.


Page D3-12


Force Displacement(N) (mm)300 0281 3.4

-156 3.4-156 6.7

0 6.7

-200

-100

0

100

200

300

400

0 1 2 3 4 5 6 7 8 9 10

Displacement (mm)

For

ce (N

)



Page D3-13


In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. Theseelectrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment andthe launch vehicle. The standard adapter system includes two diametrically opposed electrical rise off disconnects.There are two standard configurations for the electrical interface. Details for each of the two standard electricalinterfaces configurations are presented in Section D3.7 of this Appendix. Connectors conforming to MIL-C-81703 areused for both adapter system configurations. Part numbers for the standard 61 pin connectors are presented in TableD3.5-1.



Spacecraft Side LVIJ1 61 MS3424E61-50S

LVIJ2 61 MS3424E61-50S

Launch Vehicle Side LVIP1 61 MS3446E61-50P

LVIP2 61 MS3446E61-50P


Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electricalconnectors and GSE. This information includes available wire types, shielding, voltage requirements, currentrequirement and resistance requirements.

The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuityacross the separation interface. Electrical continuity across the separation interface is provided by conductive coatingson both the spacecraft and adapter interface flanges.




Page D3-14

D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS


Page D3-15

1666V-1000 Adapter Mechanical Drawings(Sheet 2 of 9)

TBS


Page D3-16


TBS


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TBS


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TBS


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TBS


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TBS


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TBS


Page D3-22


TBS


Page D4-1

D4. 1666A-1150 ADAPTER SYSTEM

D4.1 INTRODUCTION

The 1666A-1150 adapter system is comprised of the 1666 A clamp band system, separation springs, a payload adapterand electrical rise off disconnects. This appendix defines the mechanical and electrical interface characteristics,structural capability, usable volume, and accelerometer installation.


The 1666A-1150adapter is a two piece conic frustum of monocoque construction 1150 mm in height and fabricatedfrom aluminum alloy. Mechanical drawings for the 1666A-1150adapter are presented in Section D4.7 of thisAppendix.


The spacecraft and adapter interface ring in conjunction with the separation system are designed to a provide a loadpath between the spacecraft and adapter during ground operations and flight. The outboard features of the combinedcross section are designed to interface with the separation system clamp band. The cross section and material propertycharacteristics for the spacecraft and adapter interface ring are presented in Table D4.2.1-1 and Figure D4.2.1-1. Thedimensions of the spacecraft interface are presented in Section D4.7 of this Appendix.









Page D4-2


y




y

yy

x

x

x

x



Page D4-3



The structural capability of each adapter system is based on the allowable line load obtained from testing. Therelationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666A-1150 adapter system ispresented in Figure D4.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffnesscharacteristics presented in Table D4.2-1, the geometry presented in Section D1.7 of this Appendix and the quasi-static design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculatedby classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reducethe allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as pointloads through the spacecraft structure close to the interface.



Page D4-4

Figure D4.2.2-1: Capability of 1666A Adapter System - SC Mass vs Longitudinal C.G. (TBC)

Allowable C.G.SC Mass Offset (m)

(kg) 30.0kN Tension2000 3.392200 3.092400 2.832600 2.622800 2.443000 2.283200 2.143400 2.023600 1.893800 1.774000 1.66 Note: A positive force assists separation.

4200 1.564400 1.474600 1.39 Adapter system capability based on the following data:

4800 1.32 Allowable Limit Line Loads

5000 1.25 Tension Nt = 69 N/mm

5200 1.18 Compression Nc = 86 N/mm

5400 1.12 Quasi Static Loads

5600 1.07 Per Figure 3.4.1.2-1 of this Mission Planner's Guide

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600

SC Mass (kg)

Long

itudi

nal O

ffset

of S

C C

.G. f

rom

Sep

arat

ion

Plan

e

Preliminary


Page D4-5

D4.3 USABLE VOLUME

The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666A-1150adapter system ispresented in Figures D4.3-1a through D4.3-1e. The spacecraft static envelope (maximum dimensions of unloadedspacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond theuseable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacementsdue to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DMmay protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements willnot exceed 50 mm. This must be verified by coupled loads analysis.


Page D4-6

Figure D4.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter


Page D4-7

Figure D4.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 1 of 4)


Page D4-8

Figure D4.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 2 of 4)


Page D4-9

Figure D4.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 3 of 4)


Page D4-10

Figure D4.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 4 of 4)


Page D4-11


The spacecraft is secured to the adapter forward ring by a Marmon type clamp band. The two halves of the clamp bandare preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preloadis 29.5 kN for the1666 A adapter system. The separation system is released when these bolts are severed by bolt cutters.Each bolt cutter is activated by redundant pyros. At separation, the preload energy in the system and the retentiondevice moves the clamp band away from the interface ring. The retention device also secures the band to the adapterand prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter.




Stroke

(mm)

Initial Force

(N)

Final Force

(N)

Minimum Stroke 7.5 + 0.3 1500 + 20 1365 + 20

Maximum Stroke 77.7 + 0.3 1500 + 20 100 + 20

The separation event is affected by interface hardware that impart force during separation. This hardware consists ofelectrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects connectors areprovided pairs as shown in Section D4.7 of this Appendix. This symmetrical arrangement is provided to minimizeoverturning moments at separation. The force profile for each electrical disconnect is shown in Figure D4.4-1.


Page D4-12


Force Displacement(N) (mm)300 0281 3.4

-156 3.4-156 6.7

0 6.7

-200

-100

0

100

200

300

400

0 1 2 3 4 5 6 7 8 9 10

Displacement (mm)

For

ce (N

)



Page D4-13





Spacecraft Side LVIJ1 61 MS3424E61-50S

LVIJ2 61 MS3424E61-50S

Launch Vehicle Side LVIP1 61 MS3446E61-50P

LVIP2 61 MS3446E61-50P


Refer to Section 4.2.2 of the Mission Planner’s Guide for information on the electrical wiring between the electricalconnectors and GSE. This information includes available wire types, shielding, voltage requirements, currentrequirement and resistance requirements.

The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuityacross the separation interface. Electrical continuity across the separation interface is provided by conductive coatingson the spacecraft and adapter interface flanges.




Page D4-14

D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS


Page D4-15

1666A-1150 Adapter Mechanical Drawings(Sheet 2 of 9)

TBS


Page D4-16


TBS


Page D4-17


TBS


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TBS


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TBS


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TBS


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TBS


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TBS

Documents

Proton Rev.4