proton launch vehicle mission planner's guide

Revision 4 March 1999 

PROTON LAUNCH VEHICLE 

MISSION PLANNER’S GUIDE 

REVISION NOTICE 

This document supersedes the Proton Launch Vehicle User’s 

Guide - Revision 3, Issue 1 dated February 1997 

DISCLOSURE OF DATA LEGEND 

These technical data are exempt from the licensing requirement 

of International Traffic in Arms Regulation (ITAR) under ITAR 

Part 125.4(b)(13) and are cleared for public release. 

LKEB-9812-1990 

© 1999 International Launch Services 

International Launch Services 

101 West Broadway, Suite 2000 

San Diego, California 92101 USA

Revision 4 March 1999 

Anatoli I. Kiselev 

General Director 

Khrunichev State Research and 

Production Space Center 

Anatoli K. Nedaivoda 

General Designer 

Salyut Design Bureau 

Khrunichev State Research and 

Production Space Center 

PROTON LAUNCH VEHICLE 

MISSION PLANNER’S GUIDE 

Wilbur C. Trafton 

President 


Eric F. Laursen 

Chief Engineer 




San Diego, California 92101 USA 

Page i

Proton Mission Planner’s Guide, LKEB-9812-1990 

Issue 1, Revision 4, March 1, 1999 

PREFACE 

The Proton Mission Planner’s Guide is intended to provide information to potential Customers and spacecraft 

suppliers, concerning spacecraft design criteria, Proton launch capability, available mission analysis and custom 

engineering support, documentation availability and requirements, and program planning. It is intended to serve as an 

aid 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), with 

English units given in parenthesis and all identified dimensions shown should be considered as approximate. In the 

event that one or more dimensions are critical to a specific payload intregration or processing operation, the SC 

Customer should obtain accurate dimensions from International Launch Services (ILS). 

This Guide will be updated or revised periodically. All comments and suggestions for additional information are 

hereby solicited and will be greatly appreciated. 

Those Customers wishing to receive revisions and updates to this manual, or who wish to submit comments or 

suggestions, are asked to kindly contact: 



San Diego, California 92101 USA 

Telephone: (619) 645-6400 

Facsimile: (619) 645-6500 

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

Page ii



REVISION HISTORY 

Revision Date Revision No. Change Description Approval 

15 December 1993 1, Issue 1 Eric Laursen 

Chief Engineer, LKEI 

December 1995 2, Issue 1 Eric Laursen 


Proton 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 

March 1999 4, Issue 1 • Complete rewrite/update of document to 

reflect flight measured environments, 

interfaces and performance 

• Addition of Proton M/Breeze M vehicle 

data 

• Discussion of Baikonur payload processing 

and Launch operations facilities 

Eric Laursen 


Proton Division, ILS 

Eric Laursen 

Proton Chief Engineer, 

ILS 

Rich Waterman 

Manager Mission 

Development, ILS 

Page iii



FOREWORD 

International Launch Services (ILS), is pleased to offer one of the most capable commercial launch vehicles, and the 

most comprehensive launch services, available today. The Proton’s services are now available to worldwide Customers 

at a most competitive price. 

ILS is the exclusive marketing agent for commercial sales of the Proton launch vehicle worldwide, and is supported in 

its operations by full access to the incomparable technological expertise of its parent companies; Lockheed Martin 

Corporation (LMC), Khrunichev State Research and Production Space Center (KhSC), and Russian Space Complex 

Energia ( 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 members 

of the ILS team are committed to providing the most cost-effective launch services available in the world-from initial 

program planning to successful spacecraft launch. 

This document provides performance, environments and interfaces for Proton vehicles that have already been qualified 

to manage payloads up to 4.8 mt in mass. For SC exceeding 4.8 mt, parameters referred to in this document should be 

considered as PRELIMINARY and will be further defined/refined considering the geometry, mass properties and 

other physical characteristics of a particular spacecraft. 

Page iv



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........................................................................................................... 1-2 

1.2 ILS CONSTITUENT COMPANY EXPERTISE......................................................................................... 1-3 

1.3 ADVANTAGES OF USING THE PROTON LAUNCH VEHICLE .......................................................... 1-4 

1.4 VEHICLE DESCRIPTION ......................................................................................................................... 1-5 

1.5 GENERAL DESCRIPTION OF THE PROTON FAMILY ....................................................................... 1-7 

1.6 PROTON THREE - STAGE BOOSTER..................................................................................................... 1-8 

1.6.1 Proton First Stage ................................................................................................................................. 1-10 

1.6.2 Proton Second Stage.............................................................................................................................. 1-10 

1.6.3 Proton Third Stage................................................................................................................................ 1-10 

1.6.4 Proton Flight Control System ............................................................................................................... 1-11 

1.7 BLOCK DM FOURTH STAGE ................................................................................................................ 1-11 

1.8 BREEZE M FOURTH STAGE................................................................................................................. 1-15 

1.9 PAYLOAD FAIRINGS .............................................................................................................................. 1-18 

1.10 BAIKONUR INFRASTRUCTURE AND SERVICES ........................................................................... 1-23 

1.10.1 Baikonur Infrastructure ....................................................................................................................... 1-23 

1.10.2 Proton Launch Campaign.................................................................................................................... 1-25 

1.11 PLANNED ENHANCEMENTS ............................................................................................................. 1-27 

1.12 PROTON PRODUCTION AND OPERATIONS-SLIDES.................................................................... 1-27 

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

2.1 OVERVIEW .................................................................................................................................................. 2-1 

2.2 PROTON LAUNCH SYSTEM CAPABILITIES......................................................................................... 2-1 

2.2.1 The Baikonur Launch Site....................................................................................................................... 2-2 

2.2.2 Launch Availability ................................................................................................................................. 2-2 

2.2.3 Payload Fairings and Adapters ................................................................................................................ 2-3 

2.2.4 Upper Stage Capabilities ......................................................................................................................... 2-3 

2.3 PROTON ASCENT PROFILE .................................................................................................................... 2-4 

2.3.1 Proton Booster Ascent ............................................................................................................................. 2-4 

2.3.2 Block DM Trajectory and Sequence ........................................................................................................ 2-8 

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2.3.3 Breeze M Trajectory Sequence ...............................................................................................................2-12 

2.3.4 Collision and Contamination Avoidance Maneuver............................................................................... 2-13 

2.4 PERFORMANCE GROUNDRULES....................................................................................................... 2-14 

2.4.1 Payload Systems Mass Definition.......................................................................................................... 2-14 

2.4.2 Payload Fairings.................................................................................................................................... 2-15 

2.4.3 Mission Analysis Groundrules ............................................................................................................... 2-15 

2.4.4 Performance Confidence Levels ............................................................................................................ 2-15 

2.5 DIRECT INJECTION LEO MISSIONS................................................................................................... 2-15 

2.6 GEOSYNCHRONOUS TRANSFER MISSIONS..................................................................................... 2-16 

2.6.1 Launch to Geosynchronous Transfer Orbit ............................................................................................ 2-16 

2.7 ORBIT INJECTION ACCURACY.............................................................................................................2-23 

2.8 SPACECRAFT ORIENTATION AND SEPARATION.............................................................................2-24 

2.9 LAUNCH VEHICLE TELEMETRY DATA ............................................................................................. 2-25 

2.10 MISSION OPTIMIZATION/ PERFORMANCE ENHANCEMENTS..................................................2-29 

2.10.1 Nonstandard Mission Designs ..............................................................................................................2-29 

2.10.2 Subsynchronous Transfer......................................................................................................................2-29 

2.10.3 Super Synchronous Transfer .................................................................................................................2-29 

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

3.1 THERMAL/HUMIDITY ............................................................................................................................ 3-1 

3.1.1 Ground Thermal Environment................................................................................................................ 3-1 

3.1.2 Ascent ..................................................................................................................................................... 3-6 

3.1.3 Orbit ....................................................................................................................................................... 3-7 

3.1.4 Humidity ................................................................................................................................................ 3-7 

3.1.5 Air Impingement Velocity....................................................................................................................... 3-7 

3.2 CONTAMINATION ENVIRONMENT...................................................................................................... 3-8 

3.2.1 Ground Contamination Control.............................................................................................................. 3-8 

3.2.2 In Flight Contamination Control ............................................................................................................ 3-8 

3.3 PRESSURE .................................................................................................................................................. 3-9 

3.3.1 Payload Compartment Venting ............................................................................................................... 3-9 

3.4 MECHANICAL LOADS............................................................................................................................ 3-10 

3.4.1 Quasi-Static Loads................................................................................................................................ 3-10 

3.4.2 Sine and Random Vibration Loads ........................................................................................................ 3-14 

3.4.3 Acoustic Loads ...................................................................................................................................... 3-19 

3.4.4 Shock Loads.......................................................................................................................................... 3-21 

3.4.5 Environmental Test Requirements ........................................................................................................ 3-23 

3.5 ELECTROMAGNETIC COMPATIBILITY ............................................................................................ 3-24 

3.5.1 EMI Safety Margin (EMISM) ............................................................................................................. 3-24 

3.5.2 Radiated Emissions............................................................................................................................... 3-24 

3.5.3 RF Transmitter/Receiver Systems EMC............................................................................................... 3-29 

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

4.1 MECHANICAL INTERFACES .................................................................................................................. 4-1 

4.1.1 Structural Interfaces ................................................................................................................................ 4-1 

4.1.2 General SC Structural and Load Requirements ....................................................................................... 4-1 

4.1.3 Fairing Interfaces .................................................................................................................................... 4-2 

4.1.4 GN2/Dry Air Purge Option .................................................................................................................. 4-23 

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4.2 ELECTRICAL INTERFACES .................................................................................................................. 4-24 

4.2.1 Airborne Interfaces................................................................................................................................ 4-24 

4.2.2 Launch Pad EGSE Interfaces................................................................................................................ 4-40 

4.2.3 Telemetry/Command Links.................................................................................................................. 4-41 

4.2.4 Electrical Grounding............................................................................................................................. 4-47 

4.2.5 Electrical Bonding ................................................................................................................................ 4-47 

4.2.6 SC/LV Lightning Protection ................................................................................................................. 4-47 

4.2.7 Static Discharge .................................................................................................................................... 4-47 

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

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

5.1 MANAGEMENT PROVISIONS................................................................................................................. 5-1 

5.1.1 Key Personnel ......................................................................................................................................... 5-1 

5.1.2 Interface Control Document (ICD) ........................................................................................................ 5-1 

5.1.3 Schedule Monitoring............................................................................................................................... 5-1 

5.1.4 Documentation Control and Delivery ..................................................................................................... 5-4 

5.1.5 Meetings and Reviews ............................................................................................................................. 5-4 

5.1.6 DTRA Oversight ..................................................................................................................................... 5-8 

5.1.7 Quarterly Report..................................................................................................................................... 5-8 

5.1.8 Quality Provisions ................................................................................................................................... 5-8 

5.1.9 Launch License And Permits................................................................................................................... 5-8 

5.2 ILS DELIVERABLES.................................................................................................................................. 5-8 

5.2.1 ICD Development................................................................................................................................... 5-9 

5.2.2 Preliminary and Critical Design.............................................................................................................. 5-9 

5.2.3 Spacecraft Testing - Fitcheck/ Shock Test Plan, and Report ................................................................. 5-11 

5.2.4 Safety.................................................................................................................................................... 5-11 

5.2.5 Launch Campaign and Launch.............................................................................................................. 5-11 

5.2.6 Management and Reports...................................................................................................................... 5-11 

5.3 CUSTOMER DELIVERABLES................................................................................................................ 5-11 

5.3.1 ICD Development ................................................................................................................................ 5-13 

5.3.2 Preliminary and Critical Design............................................................................................................ 5-13 

5.3.3 Spacecraft Testing ................................................................................................................................. 5-13 

5.3.4 Required Safety Data and Certificates................................................................................................... 5-13 

5.3.5 Launch Campaign and Launch.............................................................................................................. 5-14 

5.4 SPECIFIC CUSTOMER RESPONSIBILITIES....................................................................................... 5-14 

5.4.1 Campaign Duration .............................................................................................................................. 5-14 

5.4.2 Spacecraft And Associated Ground Equipment ..................................................................................... 5-14 

5.4.3 Final Spacecraft Data............................................................................................................................ 5-15 

5.4.4 Spacecraft Readiness ............................................................................................................................. 5-15 

5.4.5 Removal of SC Support Equipment ...................................................................................................... 5-15 

5.4.6 Evaluation Of Launch Vehicle And Associated Services......................................................................... 5-15 

5.4.7 Spacecraft Propellants........................................................................................................................... 5-15 

5.4.8 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..................................................................................................................................... 6-1 

6.1.2 Building 92A-50 ..................................................................................................................................... 6-3 

6.1.3 Area 31 Facilities .................................................................................................................................... 6-3 

6.1.4 Building 92-1 and the Proton Launch Zone (Area 81) ............................................................................ 6-4 

6.1.5 Hotels ..................................................................................................................................................... 6-4 

6.2 SPACECRAFT PROCESSING FACILITIES............................................................................................. 6-4 

6.2.1 Facility 92A-50....................................................................................................................................... 6-5 

6.2.2 Area 31, Buildings 40/40D General Description ................................................................................... 6-14 

6.2.3 Area 31, Bldg 44 - General Description ................................................................................................ 6-20 

6.2.4 Building 254 (TBD).............................................................................................................................. 6-23 

6.2.5 Area 92, Building 92-1 - General Description....................................................................................... 6-23 

6.3 LAUNCH COMPLEX FACILITIES......................................................................................................... 6-24 

6.3.1 Area 81, Launch Pad 23 - General Description ..................................................................................... 6-24 

6.3.2 Facility Layout & Area Designations..................................................................................................... 6-26 

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

7.1 ORGANIZATIONAL RESPONSIBILITIES.............................................................................................. 7-1 

7.1.1 Khrunichev.............................................................................................................................................. 7-1 

7.1.2 Strategic Rocket Forces ........................................................................................................................... 7-1 

7.1.3 Energia (Block DM launches only) ......................................................................................................... 7-1 

7.1.4 ILS.......................................................................................................................................................... 7-2 

7.1.5 Spacecraft Customer ............................................................................................................................... 7-2 

7.2 CAMPAIGN ORGANIZATION ................................................................................................................. 7-2 

7.2.1 Contractual and Planning Organization................................................................................................... 7-2 

7.2.2 Organization During Combined Operations ............................................................................................ 7-3 

7.2.3 Planning Meetings................................................................................................................................... 7-4 

7.3 COUNTDOWN ORGANIZATION ............................................................................................................ 7-5 

7.4 ABORT CAPABILITY ................................................................................................................................. 7-6 

7.5 LAUNCH OPERATIONS OVERVIEW....................................................................................................... 7-7 

7.5.1 Launch Vehicle Processing .................................................................................................................... 7-12 

7.5.2 Spacecraft Preparations Through Arrival ............................................................................................... 7-12 

7.5.3 Area 31 (Buildings 40, 40D, 44) - Spacecraft Testing, Fueling, and Ascent Unit integration ................ 7-12 

7.5.4 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.... 7-15 

7.7 LAUNCH PAD OPERATIONS................................................................................................................. 7-16 

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

8.1 HIGH VOLUME PAYLOAD FAIRINGS................................................................................................... 8-3 

8.2 TANDEM LAUNCH SYSTEM ................................................................................................................... 8-5 

8.3 AVIONICS SYSTEM MASS UPGRADES.................................................................................................. 8-7 

8.4 NEW CYROGENIC UPPER STAGE ......................................................................................................... 8-7 

8.5 SUMMARY.................................................................................................................................................. 8-7 

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

A.1 BACKGROUND AND HISTORY..............................................................................................................A-1 

A.2 PROTON FLIGHT HISTORY ...................................................................................................................A-2 

A.3 DETAILED FLIGHT HISTORY ...............................................................................................................A-4 

A.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 .............................................................................................................C3-3 

C3.2 ELECTRICAL INTERFACES................................................................................................................C3-4 

C3.3 ENVIRONMENTAL INTERFACES......................................................................................................C3-5 

C3.4 FLIGHT DESIGN ..................................................................................................................................C3-6 

C3.5 OPERATIONS.........................................................................................................................................C3-7 

C3.6 OPERATIONS (CONTINUED).............................................................................................................C3-8 

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

D1.1 INTRODUCTION................................................................................................................................. D1-1 

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

D1.2.1 Interface Ring Characteristics............................................................................................................ D1-1 

D1.2.2 Structural Capability ......................................................................................................................... D1-3 

D1.3 USABLE VOLUME ............................................................................................................................... D1-5 

D1.4 SEPARATION SYSTEM ..................................................................................................................... D1-11 

D1.5 ELECTRICAL INTERFACE .............................................................................................................. D1-14 

D1.6 INSTRUMENTATION ....................................................................................................................... D1-14 

D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS..................................................................... D1-15 

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

D2.1 INTRODUCTION ................................................................................................................................. D2-1 

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







D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS..................................................................... D2-25 

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









D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS...................................................................... D3-14 

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









D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS ..................................................................... D4-14 

Page x



LIST OF FIGURES 

FIGURE 1-1: ILS LAUNCH SERVICES CHARTER ...................................................................................................... 1-1 

FIGURE 1.1-1: ILS CORPORATE PARENTAGE.......................................................................................................... 1-2 

FIGURE 1.1-2: ILS/LKE TASK PARTITIONS............................................................................................................ 1-3 

FIGURE 1.2-1: LAUNCH EXPERIENCE ..................................................................................................................... 1-4 

FIGURE 1.3-1: BENEFITS TO THE SPACECRAFT DESIGNER AND OWNER................................................................ 1-5 

FIGURE 1.4-1: PROTON LAUNCH VEHICLES FLIGHT PROVEN HARDWARE.............................................................. 1-6 

FIGURE 1.6-1: PROTON K AND M MAJOR HARDWARE ELEMENTS ......................................................................... 1-9 

FIGURE 1.7-1: PROTON K/BLOCK DM MAJOR HARDWARE ELEMENTS................................................................ 1-12 

FIGURE 1.7-2: PROTON DM UPPER STAGE WITHIN ITS AERODYNAMIC SHROUDS .............................................. 1-13 

FIGURE 1.7-3: BLOCK DM AS IT APPEARS IN FLIGHT.......................................................................................... 1-14 

FIGURE 1.8-1: PROTON M/BREEZE M MAJOR HARDWARE ELEMENTS ................................................................ 1-16 

FIGURE 1.8-2: BREEZE M (WITH AUXILIARY PROPELLANT TANK) ....................................................................... 1-17 

FIGURE 1.8-3: THE BREEZE M AND BREEZE M CORE AS THEY APPEAR IN FLIGHT............................................ 1-17 

FIGURE 1.9-1: PROTON K AND M PAYLOAD FAIRINGS ......................................................................................... 1-19 

FIGURE 1.9-2: BLOCK DM PAYLOAD FAIRING FOR SINGLE SPACECRAFT ............................................................ 1-20 

FIGURE 1.9-3: BREEZE-M PAYLOAD FAIRING (STANDARD) FOR SINGLE SPACECRAFT....................................... 1-21 

FIGURE 1.9-4: BREEZE M PAYLOAD FAIRING (LONG) ..........................................................................................1-22 

FIGURE 1.10.1-1: OVERALL LAYOUT OF THE BAIKONUR COSMODROME.............................................................. 1-23 

FIGURE 1.10.1.2-1: BUILDING 92A-50 SPACE VEHICLE PROCESSING FACILITY ................................................... 1-24 

FIGURE 1.10.2-1: PROTON LAUNCH CAMPAIGN FOR 92A-50 ............................................................................... 1-26 

FIGURE S-1: TANK COMPONENT FABRICATION AT KHRUNICHEV USES AUTOMATED MACHINING CENTERS........ S-1 

FIGURE S-2: FIRST STAGE PROPELLANT TANK AUTOMATED WELD-UP................................................................. S-1 

FIGURE S-3: FIRST, SECOND, AND THIRD STAGE SUB-ASSEMBLIES AWAITING INTEGRATION.............................. S-2 

FIGURE S-4: PROTON INTERSTAGE COMPONENT FABRICATION............................................................................. S-2 

FIGURE S-5: FIRST STAGE BUILD UP ON PROTON FIXTURE................................................................................... S-3 

FIGURE S-6: INTERSTAGE JOINING SECOND AND THIRD STAGES........................................................................... S-3 

FIGURE S-7: RD-253 HIGH-PRESSURE ENGINE ON THE FIRST STAGE EXTERNAL FUEL TANK............................S-5 

FIGURE S-8: PROTON FINAL ASSEMBLY HALL AT KHRUNICHEV ............................................................................S-5 

FIGURE S-9: ASSEMBLED FIRST STAGE SHOWING HOLD DOWN POINTS AND AFT END SERVICES 

CONNECTORS .................................................................................................................................................. S-6 

FIGURE S-10: END-TO END TEST OF ASSEMBLED PROTON AT KHRUNICHEV FACTORY ....................................... S-6 

FIGURE S-11: THE BLOCK DM UNDERGOING FINAL ASSEMBLY AND TESTING AT THE ENERGIA PLANT IN 

KOROLEV, NEAR MOSCOW............................................................................................................................... S-7 

FIGURE S-12: THE COMPLETED BLOCK DM STAGE, BEFORE ATTACHMENT OF THE EXTERNAL SHROUD ........... S-7 

FIGURE S-13: FINISHED BLOCK DM STAGES IN THEIR AERODYNAMIC SHROUDS, AWAITING SHIPMENT TO 

BAIKONUR ....................................................................................................................................................... S-8 

FIGURE S-14: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING EXTERNAL VIEW .......................................... S-8 

FIGURE S-15: PROTON STANDARD COMMERCIAL PAYLOAD FAIRING INTERNAL VIEW........................................... S-9 

FIGURE S-16: BREEZE M STAGE STRUCTURAL COMPONENTS IN MANUFACTURE ................................................ S-9 

FIGURE S-17: BREEZE M CORE STAGE MANUFACTURING .................................................................................. S-10 

FIGURE S-18: BREEZE M AVIONICS BAY ASSEMBLY ............................................................................................. S-10 

FIGURE S-19: BREEZE M STAGE FINAL ASSEMBLY .............................................................................................. S-11 

FIGURE S-20: SPACECRAFT ARRIVAL AT BAIKONUR.............................................................................................. S-11 

Page xi



FIGURE S-21: SPACECRAFT PROCESSING IN AREA 92 (SVPF).............................................................................. S-12 

FIGURE S-22: SPACECRAFT PROCESSING AND ENCAPSULATION IN AREA 31........................................................ S-13 

FIGURE S-23: THE COMPLETE PROTON LAUNCH VEHICLE BEING LIFTED ONTO ITS 

TRANSPORTER/ERECTOR .............................................................................................................................. S-13 

FIGURE S-24: LAUNCH VEHICLE INTERFACES AT A PROTON LAUNCH PAD, BAIKONUR....................................... S-14 

FIGURE S-25: PROTON ERECTION AT THE PAD ..................................................................................................... S-14 

FIGURE S-26: MOBILE SERVICE TOWER (MST).................................................................................................... S-15 

FIGURE S-27: LIFT-OFF OF THE PROTON K/BLOCK DM CARRYING A COMMERCIAL COMMUNICATIONS SATELLITE 

...................................................................................................................................................................... S-16 

FIGURE 2.2.1-1: THE BAIKONUR LAUNCH SITE, SHOWING AVAILABLE DIRECT INJECTION INCLINATIONS........... 2-2 

FIGURE 2.3.1-1: TYPICAL PROTON BOOSTER ASCENT ............................................................................................ 2-6 

FIGURE 2.3.1-2: TYPICAL PROTON BOOSTER ASCENT GROUND TRACK AND VACUUM IMPACT POINTS................. 2-6 

FIGURE 2.3.1-3: TYPICAL GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO THE 

SUPPORT ORBIT............................................................................................................................................... 2-7 

FIGURE 2.3.1-4: TYPICAL PROTON LOWER ASCENT ALTITUDE, INERTIAL VELOCITY, ACCELERATION, AND 

DYNAMIC PRESSURE ....................................................................................................................................... 2-7 

FIGURE 2.3.2-1: PROTON/BLOCK DM UPPER ASCENT.......................................................................................... 2-9 

FIGURE 2.3.2-2: BLOCK DM TWO-IMPULSE TRANSFER TO GEOSYNCHRONOUS TRANSFER ORBIT ..................... 2-10 

FIGURE 2.3.2-3: PROTON K/BLOCK DM UPPER ASCENT GROUND TRACK TO GSO ........................................... 2-11 

FIGURE 2.3.2-4: GROUND TRACKER ACQUISITION TIMES FOR PROTON ASCENT TO A GSO ................................ 2-11 

FIGURE 2.3.3-1: TYPICAL BREEZE M FLIGHT PROFILE TO GEOSYNCHRONOUS TRANSFER ORBIT ...................... 2-13 

FIGURE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS 

TRANSFER ORBITS......................................................................................................................................... 2-17 

FIGURE 2.6.1-2: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS 

TRANSFER ORBITS......................................................................................................................................... 2-19 

FIGURE 2.8-1: SC SEPARATION ATTITUDE........................................................................................................... 2-25 

FIGURE 3.1.1.1-1A: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS 

TIMELINE (BLOCK DM) .................................................................................................................................. 3-4 

FIGURE 3.1.1.1-1B: FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS 

TIMELINE (BREEZE M) ................................................................................................................................... 3-5 

FIGURE 3.1.1.2-1: SUPPLEMENTAL FAIRING AIR CONDITIONING SCHEMATIC (REPRESENTATIVE; DETAILED DESIGN 

CONDUCTED PER CUSTOMER REQUEST) .......................................................................................................... 3-6 

FIGURE 3.3.1-1: TYPICAL VENTING PROFILE DURING ASCENT ............................................................................. 3-9 

FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS.................................................................................. 3-11 

FIGURE 3.4.1.2-1: QUASI-STATIC DESIGN LOAD FACTORS (CONTINUED)........................................................... 3-12 

FIGURE 3.4.1.2-2: FLIGHT LIMIT ACCELERATIONS AT THE SC INTERFACE.......................................................... 3-13 

FIGURE 3.4.2-1: EQUIVALENT SINE LEVELS AT SPACECRAFT INTERFACE - FLIGHT ENVIRONMENT.................... 3-15 

FIGURE 3.4.2-2: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC IN CONTAINER 

AND SC ATTACHED TO ASCENT UNIT .......................................................................................................... 3-16 

FIGURE 3.4.2-3: TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER, TRANSPORTATION OF SC IN 

KHSC CONTAINER, TRANSPORTATION OF ASCENT UNIT .............................................................................. 3-17 

FIGURE 3.4.2-4: RANDOM VIBRATION LEVELS-GROUND TRANSPORTATION BY RAIL, SC AND ASCENT 

UNIT ATTACHED TO L/V................................................................................................................................ 3-18 

FIGURE 3.4.2-4: TRANSPORTATION OF INTEGRATED PROTON LV........................................................................ 3-19 

FIGURE 3.4.3-1: MAX EXPECTED ACOUSTIC ENVIRONMENT (THIRD OCTAVE).................................................... 3-20 

Page xii



FIGURE 3.4.4-1: PYROSHOCK SPECTRUM AT ADAPTER/PAYLOAD INTERFACE......................................................3-22 

FIGURE 3.5.2-1A: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS 

(PROTON K/BLOCK DM) (TBC)................................................................................................................... 3-27 

FIGURE 3.5.2-1B: LAUNCH VEHICLE AND LAUNCH BASE PAD NARROWBAND RADIATED EMISSIONS 

(PROTON M/BREEZE M) (TBC) ....................................................................................................................3-28 

FIGURE 3.5.2-2: LAUNCH VEHICLE AND LAUNCH PAD RADIATED SUSCEPTIBILITY LIMITS................................. 3-29 

FIGURE 4.1.1-1: LV COORDINATE SYSTEM ............................................................................................................ 4-1 

FIGURE 4.1.3.1-1A: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 3) .................. 4-4 

FIGURE 4.1.3.1-1B: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 3) .................. 4-5 

FIGURE 4.1.3.1-1C: PROTON/BLOCK DM COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 3).................. 4-6 

FIGURE 4.1.3.1-2A: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME 

DIMENSIONS (SHEET 1 OF 3) ........................................................................................................................... 4-7 

FIGURE 4.1.3.1-2B: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME 


FIGURE 4.1.3.1-2C: GENERIC PROTON/BLOCK DM COMMERCIAL FAIRING - USEABLE VOLUME 


FIGURE 4.1.3.1-3A: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT 

(SHEET 1 OF 3)............................................................................................................................................... 4-10 

FIGURE 4.1.3.1-3B: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT 

(SHEET 2 OF 3)............................................................................................................................................... 4-11 

FIGURE 4.1.3.1-3C: PROTON/BREEZE M STANDARD COMMERCIAL FAIRING GENERAL LAYOUT 

(SHEET 3 OF 3)............................................................................................................................................... 4-12 

FIGURE 4.1.3.1-4A: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME 

DIMENSIONS (SHEET 1 OF 3) ......................................................................................................................... 4-13 

FIGURE 4.1.3.1-4B: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME 


FIGURE 4.1.3.1-4C: GENERIC PROTON/BREEZE M STANDARD COMMERCIAL FAIRING - USEABLE VOLUME 


FIGURE 4.1.3.1-5A: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 1 OF 4) ....... 4-16 

FIGURE 4.1.3.1-5B: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 2 OF 4) ....... 4-17 

FIGURE 4.1.3.1-5C: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 3 OF 4)....... 4-18 

FIGURE 4.1.3.1-5D: PROTON/BREEZE M LONG COMMERCIAL FAIRING GENERAL LAYOUT (SHEET 4 OF 4)....... 4-19 

FIGURE 4.1.3.1-6A: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS 

(SHEET 1 OF 3)............................................................................................................................................... 4-20 

FIGURE 4.1.3.1-6B: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS 

(SHEET 2 OF 3)............................................................................................................................................... 4-21 

FIGURE 4.1.3.1-6C: GENERIC PROTON/BREEZE M LONG COMMERCIAL FAIRING - USEABLE VOLUME DIMENSIONS 

(SHEET 3 OF 3)................................................................................................................................................4-22 

FIGURE 4.2.1.6-1: DRY LOOP FUNCTIONAL SCHEMATIC ..................................................................................... 4-25 

FIGURE 4.2.1.7-1A: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BLOCK DM FAIRING (TYPICAL)........ 4-32 

FIGURE 4.2.1.7-1B: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON BREEZE M FAIRING (TYPICAL) ......... 4-33 

FIGURE 4.2.1.7-2: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1666 ADAPTER SYSTEM (TYPICAL)....... 4-34 

FIGURE 4.2.1.7-3: LOCATIONS OF MEASUREMENT SYSTEM SENSORS ON 1194 ADAPTER SYSTEM (TYPICAL) ...... 4-35 

FIGURE 4.2.1.7-4: INSTRUMENTATION DURING TRANSPORTATION OF SC IN CONTRACTOR’S CONTAINER......... 4-36 

FIGURE 4.2.1.7-5: INSTRUMENTATION DURING TRANSPORTATION OF SC IN KHSC CONTAINER....................... 4-36 

FIGURE 4.2.1.7-6: INSTRUMENTATION DURING TRANSPORTATION OF ASCENT UNIT......................................... 4-37 

Page xiii



FIGURE 4.2.1.7-7: INSTRUMENTATION DURING INTEGRATION OF ASCENT UNIT TO LV..................................... 4-37 

FIGURE 4.2.1.7-8: INSTRUMENTATION DURING TRANSPORTATION OF INTEGRATED PROTON LV....................... 4-38 

FIGURE 4.2.1.7-9: INSTRUMENTATION DURING ON-PAD OPERATIONS .............................................................. 4-39 

FIGURE 4.2.2-1: SPACECRAFT TO LAUNCH VEHICLE AND GROUND SYSTEMS ELECTRICAL INTERFACES............. 4-40 

FIGURE 4.2.3-1: SC TO BUNKER RF/ELECTRICAL INTERFACE BLOCK DIAGRAM ............................................... 4-46 

FIGURE 5.1.3-1A: BASELINE INTEGRATION SCHEDULE (NON-RECURRING PROGRAM)........................................ 5-2 

FIGURE 5.1.3-1B: BASELINE INTEGRATION SCHEDULE (RECURRING PROGRAM)................................................. 5-3 

FIGURE 6.1-1: BAIKONUR FACILITIES MAP............................................................................................................ 6-2 

FIGURE 6.2.1.1-1: BUILDING 92A-50 GENERAL ARRANGEMENT.......................................................................... 6-7 

FIGURE 6.2.1.4-1: BUILDING 92A-50 SPACECRAFT PROCESSING AND FUELING AREA ....................................... 6-10 

FIGURE 6.2.2-1: SPACECRAFT PROCESSING ZONE (AREA 31) .............................................................................. 6-15 

FIGURE 6.2.2.2-1: BUILDING 40, HALL 100 (COMMON HALL) LAYOUT............................................................... 6-16 

FIGURE 6.2.2.3-1: BUILDING 40D, FIRST FLOOR LAYOUT................................................................................... 6-17 

FIGURE 6.2.2.4-1: BUILDING 40D, SECOND FLOOR LAYOUT............................................................................... 6-19 

FIGURE 6.2.2.4-2: BUILDING 40D, THIRD FLOOR LAYOUT .................................................................................. 6-19 

FIGURE 6.2.3.1-1: BUILDING 44 (FILLING HALL) FLOOR PLAN .......................................................................... 6-21 

FIGURE 6.5.2.1-1: BUILDING 92-1....................................................................................................................... 6-24 

FIGURE 6.3.1-1: PROTON LAUNCH ZONE, AREA 81.............................................................................................. 6-25 

FIGURE 6.3.2.2-1: PROTON MOBILE SERVICE TOWER .......................................................................................... 6-27 

FIGURE 7.2.1-1: ORGANIZATION DURING LAUNCH CAMPAIGN ............................................................................ 7-3 

FIGURE 7.3-1: COUNTDOWN ORGANIZATION ........................................................................................................ 7-5 

FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50................................. 7-8 

FIGURE 7.5-1: TYPICAL SC CAMPAIGN OPERATIONS ASSUMING USE OF BUILDING 92A-50 (CONTINUED) ......... 7-9 

FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 .................................................. 7-10 

FIGURE 7.5-2: TYPICAL SC LAUNCH OPERATIONS ASSUMING USE OF AREA 31 (CONTINUED)........................... 7-11 

FIGURE 7.7-1: TYPICAL LAUNCH PAD OPERATIONS TIMELINE............................................................................ 7-17 

FIGURE 8-1: PROTON EVOLUTION OPTIONS........................................................................................................... 8-2 

FIGURE 8.1-1: PROTON LARGER PAYLOAD FAIRING CONCEPTS ............................................................................ 8-4 

FIGURE 8.2-1: BREEZE-M LAUNCH CONFIGURATION WITH TANDEM LAUNCH SYSTEMS (TLS).......................... 8-6 

FIGURE A.1-1: PROTON LAUNCH VEHICLE FAMILY...............................................................................................A-1 

FIGURE C3.1-1: SC PENDULUM MODEL...........................................................................................................C3-12 

FIGURE C3.1-2: SC SLOSH PROPERTIES DURING BALLISTIC FLIGHT AND AT SEPARATION ...............................C3-13 

FIGURE C3.1-3: PROPELLANT TANK GEOMETRY REQUIRED DATA ...................................................................C3-15 

FIGURE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CROSS SECTION ............................................ D1-2 

FIGURE D1.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D1-4 

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

ADAPTER ...................................................................................................................................................... D1-6 

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

ADAPTER (SHEET 1 OF 4).............................................................................................................................. D1-7 

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


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


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

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

Page xiv



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

CONNECTOR ............................................................................................................................................... D1-12 

FIGURE D1.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D1-13 


FIGURE D2.2.2-1: CAPABILITY OF 1194AX ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC)......... D2-4 

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

ADAPTER ...................................................................................................................................................... D2-6 

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


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


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


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


FIGURE D2.3-2A: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH 

1194AX X 625MM ADAPTER ....................................................................................................................... D2-11 

FIGURE D2.3-2B: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH 

1194AX X 625MM ADAPTER (SHEET 1 OF 4) (TBC) ................................................................................... D2-12 

FIGURE D2.3-2C: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH 


FIGURE D2.3-2D: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH 


FIGURE D2.3-2E: USABLE VOLUME - PROTON/BREEZE M STANDARD COMMERCIAL FAIRING WITH 


FIGURE D2.3-3A: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH 

1194AX X 625MM ADAPTER ....................................................................................................................... D2-16 

FIGURE D2.3-3B: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH 


FIGURE D2.3-3C: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH 


FIGURE D2.3-3D: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH 


FIGURE D2.3-3E: USABLE VOLUME - PROTON/BREEZE M LONG COMMERCIAL FAIRING WITH 



CONNECTOR ................................................................................................................................................D2-22 

FIGURE D2.4-2: PURGE CONNECTOR FORCE DIAGRAM (TYPICAL).................................................................. D2-23 


FIGURE D3.2.2-1: CAPABILITY OF 1666V ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D3-4 

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

ADAPTER ...................................................................................................................................................... D3-6 

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


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


Page xv



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


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



CONNECTOR ............................................................................................................................................... D3-12 


FIGURE D4.2.2-1: CAPABILITY OF 1666A ADAPTER SYSTEM - SC MASS VS LONGITUDINAL C.G. (TBC) ........... D4-4 

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

ADAPTER ...................................................................................................................................................... D4-6 

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


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


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


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



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

Page xvi



LIST OF TABLES 

TABLE 1.6-1: PROTON K AND PROTON M CHARACTERISTICS COMPARISON........................................................... 1-8 

TABLE 2.2-1: SUMMARY PROTON PERFORMANCE (PSM) TO REFERENCE ORBITS ................................................ 2-1 

TABLE 2.2.2-1: ALLOWABLE LAUNCH ENVIRONMENT CONSTRAINTS ..................................................................... 2-3 

TABLE 2.3.1-1: TYPICAL BOOSTER ASCENT EVENT TIMES....................................................................................... 2-5 

TABLE 2.3.2-1: TYPICAL BLOCK DM ATTITUDE MANEUVERS FOR GEOSYNCHRONOUS MISSION (90 O EAST 

LONGITUDE) ..................................................................................................................................................2-12 

TABLE 2.4.1-1: LAUNCH VEHICLE MISSION PECULIAR HARDWARE ..................................................................... 2-14 

TABLE 2.5-1: PROTON BOOSTER PERFORMANCE TO LOW EARTH ORBITS (DIRECT INJECTION, NO 

UPPER STAGE)................................................................................................................................................ 2-15 

TABLE 2.6.1-1: PROTON K/BLOCK DM PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER 

ORBITS............................................................................................................................................................2-18 

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

GEOSYNCHRONOUS TRANSFER ORBITS ..........................................................................................................2-18 

TABLE 2.6.1-3: PROTON K/BLOCK DM PARAMETRIC GEOSYNCHRONOUS TRANSFER PERFORMANCE DATE 

(2 BURN MISSION) .........................................................................................................................................2-20 

TABLE 2.6.1-4: PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS TRANSFER ORBITS 

.......................................................................................................................................................................2-21 

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

(CONFIGURATION 3).......................................................................................................................................2-22 

TABLE 2.7-1: BLOCK DM UPPER STAGE ORBIT INJECTION ACCURACIES.............................................................2-23 

TABLE 2.7-2: BREEZE M UPPER STAGE ORBIT INJECTION ACCURACIES ..............................................................2-23 

TABLE 2.8-1: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION I................................................................2-24 

TABLE 2.8-2: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION II ...............................................................2-24 

TABLE 2.8-3: UPPER STAGE ORBIT INJECTION ACCURACIES, OPTION III..............................................................2-24 

TABLE 2.9-1: FORMAT I - PRELIMINARY STATE VECTOR DATA PROVIDED FOLLOWING UPPER STAGE 

1 ST BURN.........................................................................................................................................................2-26 

TABLE 2.9-2: FORMAT II - TRANSFER ORBIT PARAMETERS FOLLOWING UPPER STAGE 1 ST BURN........................2-26 

TABLE 2.9-3: FORMAT III - PRELIMINARY VECTOR DATA AT SEPARATION EPOCH...............................................2-27 

TABLE 2.9-4: FORMAT IV - VECTOR DATA AT SEPARATION EPOCH.......................................................................2-27 

TABLE 2.9-5: FORMAT V - STATE VECTOR DATA AT SEPARATION EPOCH .............................................................2-28 

TABLE 2.9-6: SPACECRAFT SUPPLIED SEPARATION DATA .....................................................................................2-28 

TABLE 3.1.1-1: 3σ AMBIENT TEMPERATURES AT THE BAIKONUR COSMODROME................................................... 3-1 

TABLE 3.1.1-2: SPACECRAFT THERMAL ENVIRONMENT......................................................................................... 3-2 

TABLE 3.1.4-1: GROUND HUMIDITY ENVIRONMENT ............................................................................................. 3-7 

TABLE 3.2.1-1: GROUND CONTAMINATION ENVIRONMENT................................................................................... 3-8 

TABLE 3.4.1.1-1: GROUND LIMIT QUASISTATIC LOAD ENVIRONMENT-TRANSPORTATION AND HANDLING 

OPERATIONS .................................................................................................................................................. 3-10 

TABLE 3.4.4-1: SHOCK LOADS DURING TRANSPORTATION IN SC CONTAINER..................................................... 3-21 

TABLE 3.4.5.1-1: ACOUSTIC TEST REQUIREMENTS .............................................................................................. 3-23 

TABLE 3.4.5.2-1: SINE TEST REQUIREMENTS....................................................................................................... 3-23 

TABLE 3.5.2-1: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON K/BLOCK DM (TBC)............................ 3-25 

TABLE 3.5.2-2: LAUNCH VEHICLE RF CHARACTERISTICS FOR PROTON M (TBC)............................................... 3-25 

TABLE 3.5.2-3: LAUNCH VEHICLE RF CHARACTERISTICS FOR BREEZE M (TBC) ............................................... 3-26 

TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS ........................ 4-27 

Page xvii



TABLE 4.2.1.7-1: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR GROUND OPERATIONS (CONTINUED) ..4-28 

TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (TYPICAL).................. 4-29 

TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-30 

TABLE 4.2.1.7-2: INSTRUMENTATION QUANTITIES AND LOCATIONS FOR FLIGHT EVENTS (CONTINUED)............ 4-31 

TABLE 4.2.3-1A: C-BAND RF LINK CHARACTERISTICS......................................................................................... 4-42 

TABLE 4.2.3-1B: KU-BAND RF LINK 1 CHARACTERISTICS .................................................................................. 4-43 

TABLE 4.2.3-1C: K-BAND RF LINK 2 CHARACTERISTICS..................................................................................... 4-44 

TABLE 4.2.3-1D: KU-BAND RF LINK 3 CHARACTERISTICS.................................................................................. 4-45 

TABLE 5.1.5.1-1A: BASELINE MEETING SCHEDULE FOR NON-RECURRING PROGRAM ......................................... 5-5 

TABLE 5.1.5.1-1B: BASELINE MEETING SCHEDULE FOR RECURRING PROGRAM .................................................. 5-6 

TABLE 5.2-1: ILS DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM................... 5-9 

TABLE 5.2.2-1: DESIGN REVIEW ANALYSES.......................................................................................................... 5-10 

TABLE 5.3-1: CUSTOMER DELIVERABLE SCHEDULE FOR A RECURRING AND A NON-RECURRING PROGRAM..... 5-12 

TABLE A.1-2: PROTON 50-LAUNCH MOVING AVERAGE.........................................................................................A-2 

TABLE A.2-1: PROTON LAUNCH RECORD SUMMARY (1970-1998).........................................................................A-3 

TABLE A.3-1: PROTON LAUNCH HISTORY..............................................................................................................A-4 

TABLE A.3-1A: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-5 

TABLE A.3-1B: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-6 

TABLE A.3-1C: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-7 

TABLE A.3-1D: PROTON LAUNCH HISTORY (CONTINUED) ....................................................................................A-8 

TABLE A.3-1E: PROTON LAUNCH HISTORY (CONTINUED).....................................................................................A-9 

TABLE C3.1-1: SC MASS PROPERTIES ..................................................................................................................C3-9 

TABLE C3.1-1A: SC MASS PROPERTIES NEAR 0G.................................................................................................C3-9 

TABLE C3.1-1B: SC MASS PROPERTIES NEAR 1G.................................................................................................C3-9 

TABLE C3.1-1C: SC MASS PROPERTIES(DRY SPACECRAFT) ..............................................................................C3-10 

TABLE C3.1-2: DESCRIPTION OF LIQUID MASSES..............................................................................................C3-11 

TABLE C3.2-1: SC RF CHARACTERISTICS..........................................................................................................C3-16 

TABLE C3.2-2A: LVIJ1 UMBILICAL PIN ASSIGNMENTS......................................................................................C3-18 

TABLE C3.2.2B: LVIJ2 UMBILICAL PIN ASSIGNMENTS.......................................................................................C3-20 

TABLE C3.5-1: EGSE DESCRIPTION..................................................................................................................C3-22 

TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-23 

TABLE C3.5.1-1: SC CONTRACTOR ELECTRICAL GROUND SUPPORT EQUIPMENT (CONTINUED) .....................C3-24 

TABLE C3.5-2: FLUIDS/GASES REQUIREMENTS.................................................................................................C3-25 

TABLE D1.2.1-1: SPACECRAFT AND ADAPTER INTERFACE RING CHARACTERISTICS .......................................... D1-1 

TABLE D1.4-1: SEPARATION SPRING CHARACTERISTICS................................................................................... D1-11 

TABLE D1.5-1: STANDARD ELECTRICAL CONNECTORS .................................................................................... D1-14 



TABLE D2.5-1: STANDARD ELECTRICAL CONNECTORS..................................................................................... D2-24 







Page xviii



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 Information 

Interchange 

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 D 

AWG 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 

O C 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 Avoidance 

Manuever 

CCSCS Central Control Station of Communication 

System 

CCTV Closed-Circuit Television E 

DTSA Defense Technology Security Administration 

DUF Dynamic Uncertainty Factor 

CDR 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 xix



EIA Electronics Industry Association ft Foot; Feet 

ELAN Electronic Library for Analysis of 

Nonconformances 

FTP Filt Transfer Protocol 

EM Electromagnetic FTS Flight Termination System 

EMC Electromagnetic Compatibility G 

EMI 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 Modeling 

Code 

EUTELSAT European Telecommunications Satellite 

Organization 

G&N Guidance and Navigation 

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 

O F 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 H 

FDLC 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 I 

FPA Flight Plan Approval ICD Interface Control Document 

FPR Flight Performance Reserve ICWG Interface Control Working Group 

FSO Flight Safety Officer I/F Interface 

Page xx



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 Launch 

Services 

in. Inch(es) LN 2 Liquid Nitrogen 

INTELSAT International Telecommunications Satellite 

Organization 

LO 2 

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 

I SP Specific Impulse LV Launch Vehicle 

ITA Intregrated Thermal Analysis LVMP Launch Vehicle Mission Peculiar 

ITAR International Traffic in Arms Regulations M 

K m Meter 

kg Kilogram(s) mA Milliamps 

KhSC Khrunichev State Research and Production 

Space Center 

MDRD Module-Level Design Requirements 

Document 

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 

lb m 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 

LH 2 Liquid Hydrogen MR Mixture Ratio 

Page xxi



MRB Material Review Board P 

MRR 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 Requirements 

Document 

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 

N 2H 4 Hydrazine PMPCB Parts, Materials, and Processes Control Board 

NIOSH National Institute of Occupational Safety and 

Health 

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 

N 2O 4 Nitrogen Tetroxide psi/psf Pound(s) per Square Inch/ Pound(s) per 

Square Foot 

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

NVR Nonvolatile Residue PSM Payload Systems Mass 

O PSS Payload Separation System 

O 2 Oxygen PST Product Support Team 

OASPL Overall Sound Pressure Level PSW Payload Systems Weight 

OCC/CS 

S 

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 

ω P Argument of Perigee P&W Pratt & Whitney 

Page xxii



Q SIU Servo-Inverter Unit 

q Dynamic Pressure SLC Space Launch Complex 

Spacecraft 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 T 

s 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 Protective 

Ensemble 

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 xxiii



TVCF Transportable Vehicle Checkout Facility 

TWG Technical Working Group 

3-D Three-Dimensional 

U 

UDMH Unsymmetrical Dimethylhydrazine 

UHF Ultra-High Frequency 

UPS Uninterruptable Power Systems 

U.S. United States 

US Upper Stage-Proton Fourth Stage 

UT Umbilical Tower 

V 

V 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/M 2 Watts Per Square Meter 

WSR Weather Service Radar 

X 

Xdcr Transducer 

Page xxiv



1. PROTON LAUNCH SERVICES 

International Launch Service’s (ILS) launch support services equal or exceed in comprehensiveness and quality those 

services 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, ground 

and air transport, Government export license assistance, launch site spacecraft integration and testing, spacecraft and 

launch vehicle integration, launch site support and security services, launch of the spacecraft, and post launch mission 

support (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 with 

the supply of the launch vehicle and associated spacecraft launch services, including all required liaison with various 

United States, Russian, and other government organizations and agencies, as well as accommodation of any special 

Customer requirements. ILS will support all spacecraft preparation activities, oversee the integration of the spacecraft 

with the Proton, and conduct the spacecraft launch. The Customer need interact solely with ILS for full support in all 

aspects of the launch. 

Figure 1-1: ILS Launch Services Charter 

Lockheed 

Martin 

Corporation 

Khrunichev 

SRPSC 

RSC 

Energia 

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 1-1



1.1 CONSTITUENT ILS COMPANIES 

In 1992, the Space Systems Division of Lockheed Missiles & Space Co., Inc. established the Lockheed Commercial 

Space Company, Inc. (U.S.). This company, along with the Russian organizations now known as the Khrunichev State 

Research and Production Space Center and the Rocket Space Complex Energia, signed an agreement to form a joint 

venture, Lockheed-Khrunichev-Energia International, Inc. (LKE). This venture, which has since become the Proton 

Division of ILS, which in turn reports to the Astronautics Division of Lockheed Martin Corporation (LMC), has the 

charter to provide commercial launch services based on the Proton launch vehicle to customers worldwide. The 

Corporate parentage of ILS is shown in Figure 1.1-1. 

Figure 1.1-1: ILS Corporate Parentage 

LMC Khrunichev Energia 

LM 

Astronautics 

ILS 

This joint venture has been approved by both the U.S. and Russian governments. Lockheed Martin is the United States 

leading company in all aspects of spacecraft/launch vehicle integration, launch vehicle design, development and 

manufacture, and launch site operations. Khrunichev (along with its subsidiary, the Salyut Design Bureau) is the 

Russian designer and manufacturer of the first stage, second stage, and third stage of the Proton space launch vehicle as 

well as the Breeze-M and other upper stages. Energia is the Russian designer and manufacturer of the Block DM 

fourth stage of the Proton, and is the largest producer of launch vehicles in Russia. ILS has access to all resources of the 

constituent companies required to fulfill spacecraft launch campaign requirements. 

Lockheed Martin, Khrunichev, and Energia all function as subcontractors, reporting to ILS, during the execution of a 

Customer’s launch services contract. A memorandum of understanding (MOU) that delineates the responsibilities of 

each of the companies is in place. (Figure 1.1-2) Constituent senior management of each of the companies have 

approved 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 ready 

for immediate activation, as needed. 

Page 1-2



ILS personnel are responsible for all Customer interface activities, and for the coordination of all activities of the 

constituent companies so that contract objectives are met. They are also responsible for all program system 

engineering, custom engineering, mission analysis, program integration and program management, liaison with all 

government agencies, and liaison with the world’s financial and insurance markets. Khrunichev is responsible for 

manufacturing 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's 

Block DM fourth stage, providing the Block DM timeline, dynamics, and other data for mission analysis, and 

providing Baikonur services. 

Figure 1.1-2: ILS/LKE Task Partitions 

Lockheed Martin 

Corporation 

Launch program integration 

System management 

Insurance liaison 

Customer reporting 

Mission analysis 

Third party liaison 

Post launch services 

Performance data liaison 

Interface data liaison 

Custom engineering services 

Launch processing liaison 

1.2 ILS CONSTITUENT COMPANY EXPERTISE 

ILS/LKE 

Khrunichev 

(KhSC) 

First, second, and third stage production 

Preparation of stages for launch 

Proton-induced environment design criteria 

Interface services 

Launch vehicle GSE, consumables,and launch 

services 

Coordination of Baikonur Cosmodrone services 

CIS government liaison 

Subcontract management 

Breeze M production 

Customerliaison/report 

Subcontract direction 

U.S. Government liaison 

RSC Energia 

Block DM production 

Block DM preparation services 

for launch 

Proton spacecraft interface 

design criteria 

Proton-induced environment design 

criteria 

Russian subcontract management 

Interface services 

Launch vehicle GSE, 

consumables, and launch services 

Subcontract management 

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 and 

Energia, can provide all necessary resources required to support a spacecraft launch. ILS provides spacecraft Customers 

with a single point of contact for all mission engineering and launch support tasks using the Proton launch vehicle. 

Page 1-3



The constituent companies of ILS have more spacecraft and launch vehicle expertise than any other organizations 

providing launch services today. (Figure 1.2-1) Lockheed Martin has built more than half of all Western satellites 

flown, and has designed, built and launched the Atlas and Titan launch vehicles, as well as the Agena and Centaur 

upper stages, and the Polaris, Poseidon and the Trident strategic defense missiles. Khrunichev, in addition to designing 

and 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, and 

Almaz orbital stations, and several different orbital return capsule designs. Energia has provided all of the fourth stages 

for the Proton from the inception of the program until 1999, and has also been responsible for significant subcontract 

management. Energia is the designer and builder of the R-7 series of space launch vehicles, including the Soyuz and 

Molniya launchers, of which more than 1300 have been launched to date. During the 1980’s they developed both the 

Energia heavy lift launch vehicle and the Buran space shuttle, in addition to significant portions of the Mir space 

station hardware. The constituent companies of ILS have the necessary expertise to successfully and efficiently support 

any launch campaign. 

Figure 1.2-1: Launch Experience 

1950s 

1950s 

1960s 1970s 1980s 1990s 2000s 

Lockheed Martin 

Titan I, II, III, , and IV launches 

Atlas launches 

Agena flights 

Polaris/Poseidon/Trident launches 

Space Shuttle launches 

Mission analysis and system engineering 

Khrunichev 

200+ Proton launches 

Mission analysis and system engineering 

Spacecraft/launch vehicle integration 

RSC Energia 

Spunik, Vostok, Voskhod, Soyuz, Molniya launches 

Proton fourth stage production 

Spacecraft/launch vehicle integration 

1960s 1970s 1980s 1990s 2000s 

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 revenue 

generating 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 these 

procedures, 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 fourth 

stages, provides the Customer with mission flexibility and maximized payload capacity to orbit. This results in unique 

mission design options, including delivery of spacecraft directly to geostationary orbit by the Proton. Spacecraft apogee 

fuel may be dedicated to extended mission life, because inclination reduction and orbit circularization are 

accomplished by the Proton’s 4th stage. 

Page 1-4



The ability of ILS to meet the full range of Customer requirements guarantees each launch campaign will be 

conducted to the Customer’s satisfaction. This approach to satisfying key requirements is shown in Figure 1.3-1. ILS’s 

overriding concern in providing commercial launch services is the careful coordination of our company’s resources and 

capabilities with the Customer’s detailed requirements, so as to allow ILS to tailor each launch campaign to meet the 

individual Customer’s unique needs. This ensures that each campaign will proceed in an efficient manner toward a 

successful, 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 State 

Research and Production Space Center (KhSC) and Rocket Space Complex (RSC) Energia, both of Russia. Based on 

vehicles 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 Baikonur 

Cosmodrome in Kazakhstan. 

The Proton vehicle used for commercial launches can be provided in either a 3-stage or 4-stage configuration to meet 

the needs of Low Earth Orbit and High-Energy launch missions. The 3-stage versions of Proton, designated Proton-K 

and 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 injection 

requirements such as those for the Geosynchronous Transfer Mission. Multiple payload fairing and adapter systems are 

available 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 the 

Proton-K/Block DM launch system and the Proton M/Breeze M launch system. 

Page 1-5



Figure 1.4-1: Proton Launch Vehicles Flight Proven Hardware 

Page 1-6



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/ Block 

DM , Proton K/Breeze M, Proton M/Block DM, and Proton M/Breeze M models. A variety of supplemental orbital 

propulsion units, in a range of capabilities, can be used with either the three-stage or the four-stage Proton. In 

addition, there are multiple fairing designs presently qualified for flight, including "Standard Commercial Payload 

Fairings" developed specifically to meet the needs of Western customers. Tandem launch and multi-spacecraft 

dispenser capabilities are in development. Later sections provide more detailed information on these hardware 

elements 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 DM 

fourth stage is carried out by Russian Space Complex (RSC) Energia, also in Moscow. Production capacity for the 

commercial Proton is approximately twelve vehicles per year. 

General specifications for both versions of the Proton are similar. Overall height of the vehicle in either configuration 

is approximately 61 m (200 ft), while the diameter of the second and third stages, and of the first stage core tank, is 4.1 

m (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, when 

present, has a diameter of 4.1 m (13.5 ft). Total mass of the Proton at launch is approximately 651,500 kg (1,524,000 

lbm). 

Maximum performance capability (Payload Systems Mass, or PSM) of the Proton in the configurations of principal 

interest 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. Specific 

performance 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 1-7



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 use 

nitrogen tetroxide (N 2O 4) and unsymmetrical dimethylhydrazine (UDMH) as propellants. A summary comparison of 

the 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 in 

Table 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, mt 

hcirc.= 200 km, I = 51,6 O 

Geostationary Orbit Payload Mass, mt 

hcirc.= 36000 km, I = 0 O 

Geostationary Transfer Orbit Payload Mass, t 

ha = 36000 km, I = 7…51,6 O 

20.7 22.0 

Max. 1.9-2.1 Max. 3.0 

3.5…6.5 5.0…7.8 

Payload Bay Volume, m 3 65 (Std PLF) 100 (Std PLF) 

Launch Vehicle Structure Mass, mt 

First Stage 

Second Stage 

Third Stage 

Upper (fourth) Stage 

Propulsion System Performance 

(Maximum Vacuum Thrust), (mt) 

First Stage* 

Second Stage 

Third Stage 

Upper (fourth) Stage 

31.0 

11.75 

4.15 

3.13 

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

1070 

237 

62 

8.5 

30.6 

11.4 

3.7 

2.37 

1070 

237 

62 

2.0 

Page 1-8



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

Payload 

fairing 

Payload 

adapter 

Third stage 

4.1 m ∅ 

Second stage 

4.1 m ∅ 

First stage 

20.2m long 

7.4 m dia. ∅ 

4 4 .3 m 

lo n g 

(1) RD - 0210 

(4) - RD - 0210 

Core 

oxidizer 

tank 

Strap-on 

fuel tank 

1.7 m ∅ 

(6)RD-253 

The third stage is equipped with one 

fixed single-chamber liquid propellant 

rocket engine developing 0.6 MN thrust 

and one liquid propellant control rocket 

engine, with four gimbaled nozzles, 

developing 30 kN thrust 

The second stage is equipped with four 

gimbaled single-chamber liquid 

propellant rocket engines developing a 

total thrust of 2.3 MN 

The first stage is equipped with six gimbaled 

single-chamber liquid propellant rocket 

engines developing a total thrust of 9 MN at 

lift-off 

Page 1-9



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 tank 

during first stage flight. Each fuel tank also carries one of the six RD-253 engines that provide first stage power. Total 

first stage sea-level thrust is approximately 9.5 MN (2.1 x 10 6 lbf). Total first stage dry mass is approximately 31,000 

kg (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. Mass 

optimization 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, rated 

thrust on the engines of the first stage of Proton M is being increased by 7%. This enhancement is accomplished with 

only a minor modification to the propellant flow control valves. This modification first flew as a mission-unique 

enhancement on the Proton K that delivered the MIR space station core module in 1986. Engines incorporating this 

change have undergone extensive additional qualification firings since then, in order to approve them for use in 

standard production vehicles. Shipsets have already been produced and incorporated onto production Proton K 

vehicles. 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 the 

increased thrust engines. 

The propellant feed systems of the first and second stages have been simplified and redesigned in order to reduce 

propellant residuals in these stages by 50%, and a propellant purge system has been added to dump all residuals from 

the 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 increased 

utilization of propellants is to reduce the environmental effects of the impact of the first and second stages in the 

downrange "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 thrust 

of 2.3 MN, or 5.24 x 10 5 1bf. Total second-stage dry mass is approximately 11,715 kg (25,827 lbm); total second stage 

propellant 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 to 

carry 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 10 5 Lbf) thrust, and a four-nozzle vernier 

engine that produces 31 kN (6.9 x 10 3 lbf) thrust. Total third-stage dry mass is approximately 4,185 kg (9,226 lbm); 

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 to 

carry greater payload and aerodynamic loads, and minor structural weight reductions. 

Page 1-10



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 triple 

redundant closed-loop analog avionics system mounted principally in the Proton's third stage, using data collected 

from a distributed set of inertial reference platforms. This system also provides for flight termination in the event of a 

major malfunction during ascent. 

For Proton M, the principal modification to the Proton K's first three stages is the incorporation of a digital flight 

control system based on modern avionics technology. This digital system replaces the analog flight control hardware of 

the Proton K, which, although it has achieved a high demonstrated reliability, is based on obsolete 1960's era 

electronic designs. The new system eliminates the need to maintain a unique and limited production capability for 

Proton K avionics, and allows for simplified control algorithm loading and test. It also enables greater ascent program 

design flexibility with respect to vehicle pitch profile and other parameters. The system is being developed by the Mars 

and Proton Control Systems Companies. 

The self-contained inertial control system uses a precision three-axis gyro-stabilizer and a three-channel voting onboard 

digital computer. The digital control computer resides on the Proton M third stage and controls the flight of the 

first, 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 several 

different kerosene-type fuels. The configuration of the Proton K/Block DM is depicted in Figure 1.7-1. The Block 

DM is shown in cutaway, within its aerodynamic shrouds ("middle and lower adapters") in Figure 1.7-2, and as it 

appears in flight in Figure 1.7-3. The Block DM and its predecessor, the Block D, have together flown more than 200 

times. 

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 10 4 lbf), is gimbaled to provide three axis control during powered flight 

operations, and can be restarted as many as seven times during flight. The stage is 3.7 m (12.1 ft) in diameter, 6.28 m 

(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) attitude 

control system, comprised of two "SOZ" (or "micro") thruster units located at the base of the Block DM. The fourth 

stage 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 of 

the stage's gyro platform; it cannot undergo continuous rotation. Guidance, navigation, and control of the fourth stage 

are 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. The 

interface between the stage and its payload adapter is 2000 mm in diameter, allowing the Block DM to accommodate 

large diameter payload adapters and a static bending moment about this interface of 11,000 kg-m. The Block DM 

payload fairing is a two piece composite structure developed specifically to meet the volume and environmental 

requirements of commercial spacecraft. Payload fairing usable volume geometry is provided in Appendix D of this 

Mission Planner’s Guide. 

Page 1-11



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

Adapter 

Fourth Stage 

Block - DM 

Forward fourth 

stage shroud 

Aft fourth 

stage shroud 

Third stage 

4.1 m ∅ 

Second stage 

4.1 m ∅ 

First stage 

20.2m long 

7.4 dia m ∅ 

Payload 

Fairing 

Payload 

(1) 11D58M 

4 4 .3 m 

lon g 

(1) RD - 0210 

(4) RD - 0210 

Core 

oxidizer 

tank 

Strap-on 

fuel tank 

1.7 m∅ 

(6) RD-253 

The fourth stage is equipped with 

one liquid propellant rocket engine 

developing 86 kN thrust, and two 

"micro" engine clusters for attitude 

control and ullage maneuvers 



rocket engine developing 0.6 MN 

thrust and one liquid propellant control 

rocket engine, with four gimbaled 

nozzles, developing 30 kN thrust 







engines developing a total thrust of 9 MN 

Page 1-12



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

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) 

6280 mm 

4900 mm 

1 

1 

2 

3 

4 

5 

6 

φ2000 mm 

φ3700 mm 

φ4100 mm 

Plane of spacecraft adapter/ 

Block DM joint 

1 

6 

2 

3 

4 

5 

6 

Page 1-13



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

Page 1-14



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 improved 

payload performance and operational capabilities over the Block DM flown on the current Proton K. The Breeze M 

program was initiated in 1994 by the Khrunichev Space Center and has the full support of the Russian government. An 

isometric 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 a 

total 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, the 

propulsion system, and the avionics equipment bay, and 

2) a toroidal auxiliary propellant tank that surrounds the core section, and which is jettisoned in flight following 

depletion. 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 1-15



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

Adapter 

Fourth Stage 

Block - DM 

Forward fourth 

stage shroud 

Aft fourth 

stage shroud 

Third stage 

4.1 m ∅ 

Second stage 

4.1 m ∅ 

First stage 

20.2m long 

7.4 dia m ∅ 

Payload 

Fairing 

Payload 

(1) 11D58M 

4 4 .3 m 

lon g 

(1) RD - 0210 

(4) RD - 0210 

Core 

oxidizer 

tank 

Strap-on 

fuel tank 

1.7 m∅ 

(6) RD-253 

The fourth stage is equipped with 

one liquid propellant rocket engine 

developing 86 kN thrust, and two 

"micro" engine clusters for attitude 

control and ullage maneuvers 



rocket engine developing 0.6 MN 

thrust and one liquid propellant control 

rocket engine, with four gimbaled 

nozzles, developing 30 kN thrust 







engines developing a total thrust of 9 MN 

Page 1-16



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

1. Equipment Bay 

2. Core stage 

3. Auxiliary propellant tank 

4. Oxidizer tank 

5. Fuel tank 

6. Auxiliary propulsion for attitude control 

7. Main engine 

2610 mm 

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

? 2490 mm 

? 4000 mm 

The Breeze M uses nitrogen tetroxide (N 2O 4) 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 12 

attitude control thrusters of 13.3 N thrust each. The main engine can relight up to eight times per mission, and is 

equipped with a backup restart system that can fire the engine in the event of a primary ignition sequence failure. 

1 

2 

3 

4 

5 

6 

7 

Page 1-17



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 engine 

for the Breeze M, eleven main engines have been ground tested, some up to 6,000 seconds total burn duration. The 

Breeze M attitude control thrusters were previously used on the Kvant, Kristall, Spektr, and Priroda modules of the 

MIR space station, and are used on the Russian FGB and Service Module components of the International Space 

Station. 

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 preprogrammed 

maneuvers about all axes during parking orbit and transfer orbit coast. The stage is normally three axis stabilized 

during coast, but can be rotated at up to 2 degrees per second for thermal control. During powered flight the upper 

stage attitude is determined by mission specific pitch, yaw and roll programs. Breeze M can perform separation of a 

customer'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 exceed 

0.5 degrees per second, and the spatial attitude error in relation to the inertial coordinate system will not exceed 1 

degree, or 

2) spin-up mode, during which the stage can achieve a maximum angular rate with respect to the upper stage 

longitudinal axis of 30 degrees per second, and the spin axis deflection from the upper stage longitudinal axis will 

not 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. The 

interface between the stage and its payload adapter is 2490-mm in diameter, allowing the Breeze M to accommodate 

large diameter payload adapters and a static bending moment about this interface of 18,000 kg-m. The Breeze M stage 

is encapsulated within the payload fairing, along with the customer's spacecraft, allowing loads from the payload 

fairing to be borne by a short (600-mm) spacer ring attached to the Proton third stage equipment section, rather than 

by the Breeze M. The Breeze M payload fairing is a derivative of the payload fairing currently in use with commercial 

spacecraft on the Proton K/Block DM. Modifications consist of a redefined attachment geometry at the aft end of the 

fairing; the attachment and separation hardware, however, is essentially unchanged from the current design. Payload 

fairing 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 fairings 

typically enclose both the payload of the Proton and any supplemental orbital propulsion units employed by the 

payload. The payload fairing is jettisoned after second stage separation, with the exact time determined by spacecraft 

free-molecular heating and impact zone constraints. Normally payload fairing separation occurs at approximately 344 

seconds 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 the 

Proton K or the Proton M. 

Page 1-18



Figure 1.9-3 shows the standard commercial payload fairing available for use with the Breeze M upper stage. The 

usable volume of this fairing is comparable to that of the payload fairing for the Block DM. Figure 1.9-4 gives 

dimensions for the long version of the Breeze M payload fairing. Use of this fairing results in a decrease in 

performance to GTO of approximately 100kg. 

Figure 1.9-1: Proton K and M Payload Fairings 

250 mm 

150 mm 

250 mm 

250 mm 

100 mm 

30° 

9047.5 mm 

L = 2245 mm 

11970 mm 

12650 mm 

300 mm 

300 mm 

150 mm 

1600 mm 

2840 mm 

100 mm 

φ4100 mm φ4100 mm 

2150 mm 

15882 mm 

1000 mm 

3000 mm 

3752 mm 

14562 mm 

Page 1-19



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

Page 1-20



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

Page 1-21



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

Page 1-22



1.10 BAIKONUR INFRASTRUCTURE AND SERVICES 

All Proton launches occur from the Baikonur Cosmodrome, which is located in the Republic of Kazakhstan in central 

Asia. Baikonur is approximately 2,000 km (1,300 miles) southeast of Moscow. Located in a semiarid zone, annual 

temperature and climate vary considerably. The Proton launch system is designed to reliably operate from just such a 

climatic range. 

1.10.1 Baikonur Infrastructure 

All primary launch systems infrastructure at Baikonur are being maintained in accordance with standard Khrunichev 

Space Center (KhSC) and Russian Space Agency (formerly Russian Space Forces) policies concerning operational 

efficiency and system safety. A number of facilities in use to support the Proton and satellite mission launch campaign 

process have been upgraded to support “Western-style” launch campaign processes. Of primary importance are 

upgrades to Launch Complex 24 at Baikonur and the addition of new, integrated payload processing facilities in close 

proximity to the launch complex. Figure 1.10.1-1 shows the locations of the major facilities at the Baikonur 

Cosmodrome. 

Figure 1.10.1-1: Overall Layout Of The Baikonur Cosmodrome 

SC Processing Area 

(Facility 92A-50) 

Area 92 

Proton LV 

Processing Area 

6 

7 

Area 95 

Living Quarters 

Kranyi Airfield 

Proton LV 

Launch Complex 81 

(Pads 23 and 24) 

9 

4 

5 

90 km 

DM Upper Stage 

Processing Area 

2 

1 

Yubileini Airfield 

Sir-Darya River 

Baikonur Town 

(Leninsk) 

Area 31 

3 

DM Upper Stage 

Fueling Area 

Legend: 

Railroad 

Road 

1-2 24 km 

2-3 18 km 

2-4 18 km 

4-5 16 km 

4-6 23.5 km 

6-7 1.25 km 

7-8 9 km 

8-9 7 km 

N 

75 km 

Page 1-23



1.10.1.1 Yubileini Airport 

All launch campaign personnel, ground support equipment, and the satellite arrive at Baikonur by air via the Yubileini 

airport. Yubileini operates as a Category 3 airport facility capable of supporting all large transport aircraft. Airport 

instrumentation and support services have been brought inline with western requirements with regard to aircraft 

operations 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 a 

properly configured series of rail cars for transportation to the payload processing facility. The aircraft offload/rail car 

onload process can take up to 8 hours. ILS offers an environmentally controlled rail car for thermal conditioning of the 

satellite 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 is 

the facility in which Proton vehicles are integrated and tested before mating with the fourth-stage/payload “ascent 

unit.” Building 92A-50 houses the new SVPF (Figure 1.10.1.2-1). The Space Vehicle Processing Facility is covered in 

greater detail in Section 6 of this Mission Planner’s Guide. Area 92A-50 payload processing facilities are 

approximately 20 km from airport facilities. 

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

Page 1-24



1.10.1.3 Area 31 Payload Processing 

Area 31 was configured to enable nonhazardous and hazardous processing of Western communications satellites to 

support the first commercial Proton launches. All ILS launches through the end of 1997 were processed in Area 31 

facilities, 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/or 

processing “surge” commercial requirements. ICD’s for commercial spacecraft should continue to be written to 

support this facility. Area 31 will continue to support Russian government payloads such as the SOYUZ manned 

system. 

1.10.1.4 Area 81 Launch Complex 

Area 81 contains both active commercial Proton launch pads. Launch Pad 23 is presently launching commercial 

Proton K/Block DM vehicles in support of ILS. Pad 24 has been refurbished to support launch operations for the 

Proton K/Block DM system and the Proton M/Breeze M vehicle. Launch operations on Complex 24 will commence 

in mid-1999 after refurbishment and upgrades are complete. Complex 23 will support the ILS Proton business as a 

backup 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-related 

facilities. Area 95 houses the three primary hotel facilities used by ILS and western launch campaign teams. Hotels 

Fili, 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 utilities 

infrastructure 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 been 

implemented. 

ILS typically arranges for the accommodation of up to 30 customer and satellite contractor personnel per spacecraft 

during 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 purposes 

of initial discussions with potential Proton launch system users. As shown in Figure 1.10.2-1, a 30-day campaign is 

based on the general philosophy that the satellite is shipped from the satellite manufacturer’s facility in a near “shipand-shoot”condition. 

For most missions in which ILS is under contract, the satellite is shipped to Baikonur in a 

configuration that requires minimal nonhazardous checkout. ILS launch campaign process goals reflect the desire to 

minimize launch campaign duration, without jeopardizing mission success. The 30-day cycle is consistent with this 

philosophy. 

Page 1-25



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

L-30 days 

ILS & SC Technical Support 

Team Onsite 

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

SC Loaded 

into SC Container 

at Factory 

Ground Support 

Equipment L-34 days 

GSE Loaded in Cargo 

Containers & on 

Pallets at Factory 

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

Propellant Shipment to 

Russia Port-of-Entry 

L-11 » L-10 days L-16 » L-14 days 

• 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 

Move SC Container to Airport 

via 5-Wheel Tow; Load SC 

Container into Transport Aircraft 

Move GSE to Embarkation 

Facility for Shipment to Russia 

L-9 » L-7 days 

Rail Shipment to 

Baikonur 

• Final Operations on 

Spacecraft 

• Fuel Launch Vehicle 

Fourth Stage 

Transport Payload 

Spacehead 

To Launch Vehicle 

Assembly Area 

L-13 » L-12 days 

Checkout 

Spacecraft 

Pyrotechnics 

SC & GSE 

Arrives at Airport 

Near Baikonur 

Launch Site 

Temporary Storage 

Move Fuel Charts 

to Fuel Hall 

L-30 » L-19 days 

L-17 days L-18 days 

Fuel SC at Fueling Hall with 

SC Fuel Carts & Return to 

Technical Assembly Hall 

L-6 days L-5 » L-4 days L-3 days 

• Prepare & Mate Payload 

Section with Launch 

Vehicle 

• Checkout Electrical 

Connections 

• Transfer Launch Vehicle 

with Payload Section 

onto Erector Transporter 

Transport SC 

to Technical 

Assembly Hall 

• 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 

Install SC 

in Cleanroom 

Connect Electrical 

GSE to SC 

• Test TC&C Systems 

• Test Solar Arrays 

• Pneumatic & 

Pressure Tests 

• Test Deployment 

Mechanisms 

L-2 days 

Prepare SC 

for Fueling & 

Transfer to 

Fueling Hall 

Integrated SC 

& Launch 

Vehicle Test 

Flight 

Readiness 

Review 

L-1 » L-0 days 

Countdown 

& Launch 

Page 1-26



1.11 PLANNED ENHANCEMENTS 

Beyond the Proton M/Breeze M, our Khrunichev partners have planned a number of possible upgrades to the Proton 

launch system. Payload fairings that are longer in length and/or wider in diameter have been assessed. A tandem 

launch system concept has been studied. 

Additionally, a new liquid oxygen/liquid hydrogen high energy upper stage, to be used in place of the current 

LOX/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 of 

Moscow. Figures S-11 to S-13 show the Block DM assembly area at Energia. Figures S-14 and S-15 show external 

and internal views of the Proton Standard Commercial Payload Fairing. Figures S-16 to S-19 show the Breeze M 

being manufactured at Khrunichev. Figure S-20 shows the arrival of a commercial spacecraft, in its shipping 

container, at the Yubileni airport at Baikonur. Figures S-21 and S-22 show spacecraft processing operations in Area 92 

and Area 31, respectively. Figure S-23 shows the complete Proton launch vehicle, after integration of the Ascent Unit 

onto the first three stages, being lifted onto its transporter/erector. Figure S-24 illustrates the launch vehicle interfaces 

at the base of a Proton launch pad. Figure S-25 shows the Proton being erected onto the pad. Figure S-26 shows the 

Proton’s Mobile Service Tower. Figure S-27 shows the Proton at the moment of lift-off. 

Page 1-27



PROTON 

PRODUCTION AND OPERATIONS 

SLIDES 

S-1



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

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

S-2



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

Figure S-4: Proton Interstage Component Fabrication 

S-3



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

Figure S-6: Interstage Joining Second and Third Stages 

S-4



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-5



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-6



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-7



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-8



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

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

S-9



Figure S-17: Breeze M Core Stage Manufacturing 

Figure S-18: Breeze M Avionics Bay Assembly 

S-10



Figure S-19: Breeze M Stage Final Assembly 

Figure S-20: Spacecraft Arrival at Baikonur 

S-11



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

S-12



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-13



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

Figure S-25: Proton Erection At The Launch Pad 

S-14



Figure S-26: Mobile Service Tower (MST) 

S-15



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

S-16



2. VEHICLE PERFORMANCE 

2.1 OVERVIEW 

This section provides the information needed to make preliminary performance estimates for several variants of the 

Proton family, into a variety of mission orbits. It is organized so as to provide the user with essential background 

mission 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 the 

Block 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 out 

more than 246 launches as of December 1998. Over the last 50 launches, the Proton has achieved a success rate in 

excess of 92%. The Proton launch vehicle is presently available in three variants: the three-stage Proton booster the K 

and M variants offer similar performance in this configuration and the four-stage Proton K/Block DM and Proton 

M/Breeze M variants. Proton K/Breeze M and Proton M/Block DM variants may also be implemented, should this 

be desireable. 

The production Proton K 4-stage vehicle is equipped with a large upper stage known as the Block DM. The Block DM 

can 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 Proton 

K, it incorporates improvements to the avionics and structures of the first three stages. It also incorporates improved 

engines that have been flying since 1996. The Proton M is capable of delivering 21.0 metric tons into a 200 km 

circular, 51.6° inclination orbit. The Proton M/Breeze M will become available for commercial use in 2000. The 

Breeze M storable propellant upper stage offers enhanced performance and operational capability and is described in 

Section 1. Table 2.2-1 summarizes performance for Proton K, Proton M, Proton K/Block DM, and Proton 

M/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 2-1



2.2.1 The Baikonur Launch Site 

The Proton launch complex, consisting of spacecraft and launch vehicle processing and integration facilities and four 

launch 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 of 

Kazakhstan. The Baikonur Cosmodrome measures approximately 90 km east-to-west, and 75 km north-to-south, and 

supports many other launch vehicles, including the Soyuz, Vostok, Molniya, Zenit and Energia. Temperatures range 

from -40°C to 45°C during the year. 

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

80 o 

60 o 

40 o 

20 o 

0o 20o 0 o 

20 o 

BLACK SEA 

MEDI TERRANEAN SEA 

2.2.2 Launch Availability 

MOSCOW 

RED SEA 

40 o 

CA SPIAN SEA 

60 o 

FLIGHT AZIMUTH: 22.5 o 

INCLINATION: 72.7 o 

BAIKONUR 

COSMODROME 

(45.6 o N) 

(TYURATAM) 

80 o 

100 o 

120 o 





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 can 

be as short as 25 days. Proton has demonstrated a launch rate of four per month from multiple launch pads, and a long 

term average launch rate of approximately twelve per year. The capability of the Proton system to launch in severe 

environmental conditions decreases launch delays and ensures that payloads reach orbit as scheduled to begin revenue 

generating activities. The short turnaround time between launches can ensure that spacecraft constellations are 

deployed quickly, minimizing the time required to enter service. 

140 o 

160 o 

180 o 

Page 2-2



Table 2.2.2-1: Allowable Launch Environment Constraints 

Temperature -40 O C to 45 O C 

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 of 

this 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. For 

Proton K/Block DM and initial Proton M/Breeze M flights, fairing jettison occurs at approximately 342 to 344 

seconds (121-125 kilometers altitude) into the flight. Free molecular heating is below 1,135 W/m 2 for these cases. 

2.2.3.2 Payload Adapters 

Available Proton payload adapters are shown in Appendix D. These adapters provide mechanical and electrical 

interfaces compatible with established commercial launch industry standards and can accommodate a wide range of 

payload 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 lift 

capability of the Proton’s first three stages, coupled with the Block DM’s high performance, allows the placement of 

spacecraft directly into their final orbits, reducing the size and complexity of spacecraft propulsion systems. The Block 

DM’s multiple start capability, designed for up to 7 starts in flight (with 5 demonstrated to date), and a minimum 24hr 

orbital lifetime, can increase mission utility through unique mission design. Block DM orbital lifetime can be 

extended 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 the 

Breeze M can be restarted up to 8 times in flight and allows the stage to offer high precision placement of spacecraft 

into orbit. With storable propellant, Breeze M orbital lifetime is limited only by available on board battery power and 

is currently in excess of 24 hours. The jettisonable, toroidal propellant tanks offer significant mission design flexibility 

and enable launch services to be offered for low and high energy delivery requirements. 

Page 2-3



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 stage 

and/or payload) into a 200 km (108 nmi) circular parking orbit inclined at either 51.6°, 64.8°, or 72.7°. A standard 

ascent trajectory is required to meet jettisoned stage and payload fairing impact point constraints. The use of a standard 

ascent 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 its 

final 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 time 

of occurrence for major ascent events for a typical launch. The six Stage 1 RD-253 engines are ignited at 

approximately 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 are 

functioning nominally before being committed to launch. The launch vehicle executes a roll maneuver beginning at 

T+10 sec. to align the flight azimuth to the desired direction. The vehicle incurs its maximum dynamic pressure of 

3,890 N/m 2 (800 psf) approximately 65 sec. into flight. Stage 2’s four RD-0210 engines begin their ignition sequence 

at 122 sec. and are commanded to full thrust when Stage 1 is jettisoned at 126 sec. Payload fairing jettison typically 

occurs at 344.2 sec into flight, depending on spacecraft heating constraints. Stage 3’s vernier engines are ignited at 332 

sec. followed by Stage 2 shutdown at 334 sec. Stage 2 separation occurs after six small, solid retro-fire motors are 

ignited at 335 sec. into flight. Stage 3’s single RD-0210 engine is ignited at 339 sec. and burns until shutdown at 567.1 

sec. The four vernier engines burn for an additional 10 sec. and are shutdown at 585 sec. After a 10 second coast, the 

Stage 3 retro-fire motors are ignited and Stage 3 is separated from the upper stage or spacecraft. Figure 2.3.1-2 shows 

the ascent vacuum instantaneous impact points and ground track. Figure 2.3.1-3 shows the ascent telemetry coverage 

provided 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 2-4



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 2-5



Figure 2.3.1-1: Typical Proton Booster Ascent 

Separation 

of 1st stage 

T=126.73 

V=1649.4 

H=43.5 

n x =3.6/0.4 

q=3350 

1750 x 6 = 10500 kN 

q max 

T=65 

V=125.1 

H=10.5 

n x=1.97 

q=35265 

583 x 4 = 2332 kN 

L = 310 

Separation 

of 2nd stage 

T=335.12 

V=4331.9 

H=148.11 

n x =2.8/0.04 

q=0 

Retro 

Separation of payload fairing 

T=344.2 

V=4340.5 

H=157.7 

n x=0.9 

q=0 

L = 1985 

583 + 31 = 614 kN 

Separation 

of 3rd stage 

T=589.0 

V=7477.4 

H=228.0 

n x =0.13/0 

q=0 

Ignition No. 1 

of 4th stage 

n x=0.44 

Burnout 

n x=0.9 

Ignition No. 2 

of 4th stage 

n x=0.9 

Burnout and 

separation 

n x=1.7/0 

T = time (seconds) 

V = velocity (meters per second) 

H = altitude (kilometers) 

L = horizontal distance (kilometers) 

q = dynamic pressure (Pa) 

n x = axial acceleration at end of prior 

Retro 

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

L a ti t u d e , d e g r e e s 

90 

75 

60 

45 

30 

15 

0 

-15 

-30 

-45 

-60 

-75 

Orbit Injection 

Vacuum Impact Points 

Moscow 

Stage 1 Jettison PLF Jett Option 1 

stage/beginning of next stage 

83.5 kN 83.5 kN 

Stage 2 Jettison PLF Jett Option 2 

-90 

-180 -150 -120 -90 -75 -60 -45 -30 -15 0 15 30 45 60 75 90 105 135 165 

-165 -135 -105 120 150 

Longitude, degrees 

180 

Page 2-6



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

Ussurick (USK) 

Kolpashevo (KLP) 

Dshusali (DZhS) 

0 100 200 300 400 500 600 

Time from Liftoff (seconds) 

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

Velo cit y ( m/s ec) 

Accel erat ion ( g) 

Alti tud e (k m) 

Dynami c pres sure (N/m 2 ) 

Page 2-7



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 the 

Block 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 DM 

is delivered to the 200 km circular parking orbit with a 51.6-deg inclination. Once on orbit and separated from the 

Proton third stage, the Block DM executes a 15-min duration maneuver to properly align its longitudinal axis for the 

first burn. After the alignment maneuver, the Block DM enters into a stabilized flight mode. Twenty-five min after the 

longitudinal alignment maneuver, the Block DM executes a 180-deg turn about the roll axis to compensate for 

possible gyroscopic drift. This maneuver can also help with spacecraft thermal management. Forty min after the roll 

maneuver, the Block DM reaches the first ascending node, and the two SOZ unit’s four 2.5-kg axial loading engines 

begin a 300-sec burn to settle the stage’s propellants. After this settling burn, the main engine ignites, raising the 

transfer orbit apogee to geosynchronous altitude. The first main engine burn lasts approximately 450 sec. After engine 

cutoff, the Block DM executes another alignment maneuver to place the longitudinal axis in the correct orientation for 

the second burn. The Block DM then enters stabilized flight for the approximately 5 hr and 15 min transfer required 

to reach transfer orbit apogee. At approximately 2.5 hr into this transfer ellipse, another 180-deg rotation about the 

roll axis is executed. After reaching the transfer orbit apogee, the Block DM initiates another 300-sec propellant 

settling burn followed by a main engine burn of approximately 230 sec. duration to circularize the orbit and reduce the 

inclination to 0 deg. The Block DM then commands spacecraft separation 14.8 ±0.05 sec after the end of the final 

burn. Figure 2.3.2-3 depicts the upper ascent ground track and Figure 2.3.2-4 shows available tracker coverage 

provided by the CIS ground stations. Telemetry (including state vector) data can also be collected from the Block DM 

during transfer orbit flight by Russian geostationary spacecraft. Table 2.3.2-1 gives the times and values of the Block 

DM attitude maneuvers. These maneuvers may be modified to assist in spacecraft thermal management; the Block 

DM can perform a maximum of 11 such maneuvers between the 1 st and 2 nd burn during a normal mission. with a 

mission maximum of 14. The standard mission scenario for commercial spacecraft involves retention of the SOZ units 

to provide attitude control following main engine shutdown. This allows spacecraft separation to be delayed until after 

the 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 in 

order to enable larger masses to be delivered to low earth orbit. The larger mass is made up of the heavier payload and 

more fuel on the Block DM stage for orbital maneuvers. The “enhanced” Block DM performance level is enabled 

with the suborbital burn. 

Normally, injection into final orbit occurs at 90 O however, longitude placement for geosynchronous orbits can be 

controlled with an increment of 12.25 deg by allowing the Block DM and payload to coast in the standard parking 

orbit for the required number of revolutions, and initiating the first transfer orbit burn on either the ascending or 

descending node. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parking 

orbit revolutions are within the lifetime constraints of the Block DM, giving complete 360-deg coverage. 

Page 2-8



Figure 2.3.2-1: Proton/Block DM Upper Ascent 

180° rotation 

about roll axis 

Begin attitude 

maneuver for 

block DM 1st burn 

Begin propellant 

settling burn 

(MES1 - 300sec) 

Block DM 

1st burn 

(MES1) 

Event description 

(block DM upper ascent) 

MECO1 

Block 4 upper adapter Jettison 

SOZ unit first setting burn (SOZ1) 

Block 4 DM first burn 

SOZ unit engine shutdown 

SOZ unit second burn (SOZ2) 

Block 4 DM second burn 

SOZ unit engine shutdown 

Block 4 DM separation 

Begin altitude 

maneuver for 

block DM 

2nd burn 

Event time 

(sec) 

637 

5437 

5732 

5737 

24,800 

25,105 

25,105 

25, 345 

180° rotation 

about roll 

axis 

Begin propellant 

settling burn 

(MES1 - 300sec) 

MECO2 

Block DM 

2nd burn 

(MES 2) 

Block DM/ 

payload 

separation 

MECO2 

+14.8sec 

Page 2-9



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

SC Separation 

Target Orbit 

H a = 36,000 km 

H p = variable km 

7 

DV 3 

Parking Orbit 

1 - Launch from Baikonur Cosmodrome and booster ascent 

6 

5 

Intermediate 

Transfer Orbit 

2 - (DV 1 ) For sub-orbital burn cases, first Block DM main engine fairing 

3 - Coast to ascending node of parking orbit 

4 - (DV 2 ) Block DM main engine firing to raise orbit from LEO to GTO 

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

6 - (DV 3 ) Block DM main engine firing to optimally raise perigee and reduce inclination 

7 - Spacecraft separation from Block DM 

3 

DV 1 

2 

1 

4 

DV 2 

i ref = 51.6 o 

GSO 

i t= variable 

Page 2-10



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

Booster 

Burn 3 

K O 

AY1 

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 

0 

10,600 15,60 

Time from liftoff 

(seconds) 

0 

Burn 2 

20,600 25,600 

Page 2-11



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

Start Time of Attitude 

Maneuver (hr.min.sec) 

Change in 

Pitch Angle 

(deg) 

Change in 

Yaw Angle 

(deg) 

Change in 

Roll Angle 

(deg) 

Duration of 

Maneuver 

(sec) 

00.10.44 43.00 - - 86 

Comments 

00.10.59.9 33.02 7.54 -13.25 269 Attitude alignment for First 

Block 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 the 

third stage, the Breeze upper stage performs a small propulsive maneuver to deliver itself and the attached satellite to a 

standard low earth parking orbit. After a coast of approximately 45 minutes, the Breeze stage performs the second of 

four propulsive maneuvers. This second main burn is used to begin the process of raising the apogee of the transfer orbit 

to 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 - 7,000 

km; the actual value is mission and satellite mass dependent. After one full revolution in this intermediate transfer 

orbit (approximately 2.5 hours), the third main burn of the Breeze M main engine occurs raising apogee altitude to 

geosynchronous. Perigee altitude and inclination are adjusted somewhat based on the trajectory optimization process 

that 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 compatible 

with the thermal and solar incidence requirements for the satellite. Roll maneuvers can be programmed to ensure even 

heating/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 fourth 

burn, the satellite is oriented and separated from the upper stage. Total elapsed time from launch for a typical Breeze 

M mission is approxiamately 10 hours. 

Page 2-12



Normally, injection into final orbit occurs at 90 O however, longitude placement for geosynchronous orbits can be 

controlled with an increment of 12.25 deg by allowing the Breeze M and payload to coast in the standard parking orbit 

for the required number of revolutions, and initiating the first transfer orbit burn on either the ascending or descending 

node. Each parking orbit revolution will change the final longitude by 22.5 deg. Fifteen standard parking orbit 

revolutions 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 geosychronous 

transfer 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 or 

contamination of a customer’s spacecraft. The separation event provides a typical relative velocity between the 

spacecraft and the upper stage of 0.3 meter/sec. 

Page 2-13



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 the 

relative separation velocity between the upper stage and the spacecraft. This is followed by a burn of opposing pairs of 

SOZ thrusters to SOZ propellant depletion (with zero net delta-velocity). Finally the LOZ tank residuals are vented 

through a pair of forward facing nozzles located on either side of the stage near the bottom of the LOZ tank. These 

nozzles are canted at 15 degrees from the vertical in such a way as to spin up the stage whole simultaneously adding a 

relative 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 reorient 

the stage. A small propulsive maneuver is made to increase relative velocity between the upper stage and 

spacecraft. After completion of the maneuver, the upper stage propellant tanks are depressurized and the stage is made 

inert. Relative velocity increments between the spacecraft and upper stage are similar to those for Block DM are 

achieved. 

2.4 PERFORMANCE GROUNDRULES 

A number of standard mission groundrules have been used to develop the reference Proton K and Proton M 

performance 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. Payload 

Systems Mass (PSM) is defined as the total mass delivered to the target orbit, including the separated spacecraft, the 

spacecraft-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 for 

estimating the performance effects of various mission-peculiar hardware requirements. As a note, the performance 

effects shown are approximate. 

Table 2.4.1-1: Launch Vehicle Mission Peculiar Hardware 

∅1666 mm 

two piece adapter 

∅1194 mm adapter 

500 mm height 

∅1194 mm adapter 

625 mm height 

Block DM Breeze M 

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

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

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

Page 2-14



2.4.2 Payload Fairings 

All performance quotes in this document are based on use of standard payload fairing options. For Proton K/Block 

DM the standard commercial payload fairing is used. For Proton M/Breeze M the baseline “short” payload fairing 

option that enables a usable volume similar to Block DM is used. For Proton M/Breeze M, a performance 

degradation 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 



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 Proton 

launches. Variable mass satellites are delivered to a geosynchronous transfer orbit with a 35,786 km apogee altitude, a 0 

degree argument of perigee, and a variable orbit perigee and inclination consistent with the liftoff mass of the satellite 

and the delivered launch vehicle performance. From this point, the satellite will perform the remaining perigee raising 

and inclination reduction to reach geosynchronous orbit. 

For reference purposes, ILS has established a reference goesynchronous transfer mission performance quotation based 

on an orbit that is 1,500 m/sec delta-velocity from geosynchronous orbit. This reference mission is indicative of the 

geosynchronous transfer missions used by vehicles launched from low inclination launch sites. The reference orbit 

assumes a 5,500 km perigee altitude, a 35,786 km apogee altitude, a 25.0 degree orbit inclination, and a 0.0 degree 

argument of perigee. 

Proton K/Block DM Performance - Performance to elliptical transfer orbits with GSO apogees is shown in Figure 

2.6.1-1 for the Proton K/Block DM. Data is shown that represents launch vehicle performance as a function of 

payload systems and residual delta-velocity from targeted transfer orbit to geosynchronous orbit. Analyses have been 

conducted to determine the near optimum orbit that can be achieved with Proton given a spacecraft mass and the 

various performance variants of the Proton launch system. Given a payload mass and launch from the Baikonur 

Cosmodrome, the upper stage of Proton delivers the satellite to a high-energy GTO that results in the minimum deltavelocity 

remaining to reach GSO. The derived perigee altitudes and orbit inclinations are provided in Table 2.6.1-1 for 

the standard Block DM 2-Burn mission profile. Table 2.6.1-3 provides parametric GTO performance data for a full 

range 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 the 

Block DM upper stage enabling a higher upper stage/payload mass to be delivered to low earth orbit. The 3-Burn 

mission is executed by the Proton K booster delivering a 23,000 kg (50,700 lb) orbit unit “stack” to a high-energy 

suborbital state. The Block DM, after separation from the third stage of Proton, performs a main engine firing to 

achieve a low earth parking orbit. The Block DM mission then progresses similarly to the two burn profile. In this 

instance, 3 firings of the Block DM main engine result in a higher performance capability results as the Block DM 

reaches LEO with a higher propellant load than if it were offloaded and delivered directly to LEO by the booster 

statges. 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 to 

4,600 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. Figure 

2.6.1-2 plots optimum GTO payload versus Delta-Velocity to GEO for these three performance variants of Proton 

M/Breeze M. The first configuration (Configuration 1) represents the performance capability for Proton K/Breeze M 

for initial flights which are to commence in early 1999. In this configuration, the Proton K booster delivers the Breeze 

M to orbit. A payload systems mass of 4800 kg can be delivered to a reference GTO in this case. With introduction of 

the Proton M booster, vehicle performance to GTO increases to 5200 kg PSM. Flights 5, 6 and 7 of the Breeze M, to 

be launched during the second half of 2000, will deliver this performance capability to GTO and are represented by the 

Configuration 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 missions 

for the mature flight configuration, Configuration 3, of Proton M/Breeze M. 

Page 2-16



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

Payload System Mass (kg) 

5500 

5000 

4500 

4000 

3500 

3000 

2500 

Enhanced 

Standard 

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

Delta-V to GSO (m/s) 

Page 2-17



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/m 2 ; 

Launch Mission Duration Approximately 6.5 hours. 

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

Orbits 

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

(m/s) (deg) Hp (km) 

1343.6 19.83 5600 4600 

1375.0 20.52 5310 4666 

1400.0 21.08 5082 4719 

1425.0 21.64 4859 4772 

1450.0 22.21 4640 4825 

1475.0 22.79 4425 4879 

1500.0 23.37 4217 4932 

1525.0 23.96 4016 4985 

1500.0 25.00 5500 4910 

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



Page 2-18



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

Payload Systems Mass (kg) 

6500 

6000 

5500 

5000 

4500 

4000 

3500 

3000 

Config 1 

Config 2 

Config 3 

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

Delta-V to GSO (m/s) 

Page 2-19



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; 



Page 2-20



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; 



Configuration 1: Performance for Proton K/Breeze M Flights (Flights 1, 2, and 3) and contrained performance 

for Proton M/Breeze M (Proton/Block DM backup) 

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

Configuration 3: Performance for Proton M/Breeze M Flights 8 and on 

Proton Vehicle Variant Performance (kg) 

Page 2-21



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; 


Launch Mission Duration Approximately 10 hours. 

TO BE SUPPLIED 

AT A LATER TIME 

Page 2-22



2.7 ORBIT INJECTION ACCURACY 

Table 2.7-1 shows Block DM 3-sigma accuracy predictions for various missions. The accuracy predictions are 

enveloping values and mission-unique analysis will be performed to verify that payload accuracy requirements are 

satisfied. 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 

200 km Circular 

Support Orbit 


Orbit 

5500 km X 36000 km 

@ 25.0 deg GTO 

Perigee Apogee Inclination Arg of Perigee Longitude of 

Ascending Node 

Period 

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

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

± 400 km ± 150 km ± 0.5 O ± 0.25 O ± 0.5 O ± 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 


Support Orbit 

Perigee Apogee Inclination Arg of Perigee Longitude of 

Ascending Node 

Period 

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

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


Orbit 

5500 km x 36000 km 

@ 25.0 deg GTO 

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

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

Eccentricity Longitude Inclination Period 

Geostationary 0.009 ± 1 O 0.75 O ± 20 min 

Page 2-23



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 a 

longitudinal spinup, or with a transverse spinup enabled with either asymmetric separation springs or via upper stage 

maneuvering. Tables 2.8-1 through 2.8-3 show the separation and pointing accuracies for the Block DM and Breeze M 

upper 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 verify 

that payload requirements are satisfied. The Proton upper stagges can accommodate Customer requirements for spin 

rates 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 Z sc axis -6 deg/s ± 1% 

SC tip-off rate about Y sc and Z sc 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 


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

SC tip-off rate about Y sc and Z sc axes ± 0.7 deg/ s 



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



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


Page 2-24



Figure 2.8-1: SC Separation Attitude 

Up p e r S t ag e 

L o n g it u d i na l a x i s 

Sp a c e c ra ft P o s i t i o n 

V e c t o r 

2.9 LAUNCH VEHICLE TELEMETRY DATA 

T ra ns v e r s e 

Sp i n 

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

X 

Y 

Su n L i ne 

ILS will provide (in Formats I-V , Tables 2.9-1 through 2.9-5 respectively) Launch Vehicle state vector data following 

stage 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 Control 

Center. 

ILS has adopted standard formats regarding orbital state vector data that are provided to the launch services customer 

during and after the launch mission. These standard formats enable the satellite operator to properly determine orbital 

conditions at various times during the mission. The standard data is transmitted to the spacecraft Mission Control 

Center 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 2-25



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 1 st Burn 

Item Units 

Magnitude of �V for main engine 1 st firing m/sec 

Upper stage 1 st 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 1 st Burn 

Item Units 

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

Upper stage 1 st 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 is 

the estimated time of the Block DM 1 st burn cutoff. 

Page 2-26



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

Item Units 

Magnitude of �V for main engine 2 nd firing m/sec 

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

Actual time of Spacecraft separation from 4 th 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)) 


Eccentricity --- 



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 


the estimated moment of spacecraft separation. 

Page 2-27



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

Item Units 

Fourth Stage third ignition time(UTC) for deorbit 

maneuver 

Fourth Stage third burn cutoff time (UTC) for deorbit 

maneuver 

De-orbit �V m/s 

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

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


the estimated moment of spacecraft separation. 

Within two days following separation, the SC contractor will provide spacecraft derived state vector data to ILS as 

shown 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 until 

15 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 


Eccentricity -- 



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 2-28



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 and 

Block DM and Breeze M Upper Stages. ILS can take advantage of the unique capabilities of the Proton launch system 

to 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. This 

maneuver consists of separating the spacecraft in a transfer orbit whose apogee is less than its final desired value. The 

spacecraft then performs one or more burns to raise apogee, as well as a final burn to raise perigee. This maneuver is 

generally not used with the Proton launch system due to the Proton’s high throw weight capability, but it can be 

investigated if desired. 

2.10.3 Super Synchronous Transfer 

Use of a super synchronous transfer trajectory can increase performance into 24-hr orbits. The super synchronous 

transfer trajectory takes advantage of the increased efficiency with which the inclination change can be performed at a 

higher altitude. A number of vehicle hardware and satellite operational constraints interact with missions utilizing 

supersynchronous transfer. ILS is able to asses supersynchronous missions by specific request and with detailed satellite 

configuration data. 

The long coast life and multiple restart capabilities of the Block DM and Breeze M can also assist in constellation 

phasing, thereby reducing spacecraft propellant usage. 

ILS also has the resources available to develop special hardware items, such as dispensers for multiple spacecraft, to 

meet unique mission requirements. 

Page 2-29



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 the 

Yubeleini Airport through launch base processing, launch and separation. SC component temperatures to be used for 

assessing 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 transportation 

temperatures are provided in Table 3.1.1-2. During transport, the spacecraft is air-conditioned either by Customer 

provided air-conditioning equipment or by a railcar mounted thermal control unit. While on the pad, thermal control 

is provided by a pad air conditioning system and/or a liquid thermal control system (refer to Section 3.1.1.1 for a 

description of the on-pad systems). 

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

Month Max ( O C) Min ( O C) 

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 3-1



Table 3.1.1-2: Spacecraft Thermal Environment 

Location/event Item Temperature 

(deg. C) 

Temperature Control 

Min. Max. 

Spacecraft Container Inlet air to 15 30 External air-conditioning unit provided by ILS/Khrunichev. 

Fueling Hall 

container 

Flow rate of 2000 to 8000 cubic m/hr. Monitoring of the 

temperature environment in S/C container is the SC 

contractor’s responsibility. * 

92A-50 

Spacecraft 15 25 Building air-conditioning* 

Bldg 44 

Payload Processing Area 

ambient air 17 27 

92A-50 


Area 31 rm 119 

Fourth Stage Integration Area 


92A-50 


Area 31 rm 100A 


Upper Stage fueling station/ Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev. 

(Breeze M only) 

ambient air 

Maximum flow rate to 4000 m3 /hr. * 

Bldg 92-1 Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev. 

ambient air 

Maximum flow rate to 4000 m3 /hr. * 

Encapsulated in fairing during Spacecraft 15 25 External air-conditioning unit provided by ILS/Khrunichev. 

transportation 

ambient air 

Preconditioning will be established by Khrunichev (taking 

into account SC contractor recommendations) prior to 

transport to pad to guarantee temperature range during 

erection. Maximum flow rate to 6000 m3 /hr. * 

Erection 

On-pad with air-conditioning 

through umbilical 

Spacecraft 

ambient air 

10 30 No active temperature control is provided. 

Block DM 

Spacecraft 13 27 Air-conditioning through umbilical, flow rates adjustable 

Breeze M 

ambient air 13 23 from 5000 to 13000 M 3 

/hr.* 

On-pad following removal of Spacecraft 10 30 Temperature control panels in fairing. 

air-conditioning umbilical ambient air 

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 + 2 O C. 

Page 3-2



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 insulation 

during the erection of the Launch Vehicle on the pad when there is no active air conditioning. During transportation to 

the pad, conditioned air is provided to the spacecraft from the Thermal Control Railcar. At the pad, the air 

conditioning is disconnected and the Launch Vehicle is erected. Following erection, the Mobile Service Tower is 

brought 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 air 

conditioning to the connection of on pad air conditioning is 4 hours maximum. A thermal analysis is performed to 

verify that under worst case ambient conditions, the spacecraft temperature will not exceed allowable temperature 

limits during the erection process. 

The on-pad air conditioning system remains active 24 hours a day until approximately 1.5 hours prior to launch when 

preparations are begun for Mobile Service Tower rollback. To provide thermal conditioning of the fairing after Mobile 

Service 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 to 

ethylene glycol filled pipes which run to a thermal control system in the launch pad complex. This system is activated 3 

hours prior to launch and purged with dry nitrogen 5 minutes prior to launch to insure that the lines are free of liquid 

prior to liftoff. Should the launch be aborted, the liquid system can be quickly reactivated and the Mobile Service 

Tower will be brought up to renew air-conditioning within 2 hours. A schematic of both the liquid and air thermal 

control systems is shown in Figures 3.1.1.1-1a and 3.1.1.1-1b along with an approximate operational timeline, for 

both the Block DM and the Breeze M upper stages. 

Page 3-3



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

Liquid flow rate = 0.250 - 0.9 m3/h 

T = -10 to +80oC deg. C 

27 

22 

17 

12 

Thermal 

cover 

mounting 

Launch Vehicle installed 

on launch pad 

Winter 

Summer 

Preliminary 

preparation 

(12-28 hr) 

Transportation 

92-1 to 

launch pad 

Ascent 

Unit 

H 

~H/2 

LSTR 

Liquid system, 

thermal regulation 

SC 

Air outlet 

emissivity =



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

Liquid flow rate = 0.250 - 0.9 m3/h 

de g. C 

27 

22 

17 

12 

Thermal 

cover 

mounting 

T = -10 to +80 o C 

Launch Vehicle installed 

on launch pad 

Winter 

Summer 

Preliminary 

preparation 

Air out le t 

emissivity =



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. Up 

to 480 m 3 /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 in 

existing 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 per 

customer request) 

3.1.2 Ascent 

35 o 

30 

Fairing/Block DM Interface Plane 

1200 

2100 

7.5 o 

Battery Radiator Cooling System Nozzle Location 

840 

160 

260 

30 o 

300 

Fairing Access Door 

During Ascent, the launch vehicle will be exposed to aerodynamic heating. Following fairing jettison, the space craft 

will be exposed to solar radiation and free molecular heat flux. A thermal analysis will be performed using the 

Customer supplied SC thermal model to predict spacecraft temperatures during this phase of the mission. The heat 

flux density radiated upon the spacecraft by the internal surfaces of the NF should not exceed 500 W/M 2 from the time 

of launch until NF jettison. For commercial missions, the fairing is jettisoned at 342 - 344 seconds (121 – 125 km 

alt.) into flight and the free molecular heat flux shall not exceed 1135 W/M 2 at any time following fairing jettison. 

Page 3-6



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 provide 

desired sun angles for maintaining thermal control. An integrated thermal analysis is performed to determine 

spacecraft 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) % 

Inside Spacecraft 

Container 

Min. Max. 

Humidity Control 

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 fairing 

during transportation 

Fairing conditioning in 

Bldg 92-1 

Air inside fairing 35 60 ILS/Khrunichev provided air conditioning 

Air inside fairing 35 60 ILS/Khrunichev provided air conditioning 

Erection Air inside fairing 35 60 None 

On-pad with airconditioning 

through 

umbilical 

On-pad following removal 

of air-conditioning 

umbilical 

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

Air inside fairing 0.5 60 None 

adjustable from 5000 to 13000 M 3 /hr. 

Max dewpoint approx 0 deg. C. 

Note: Should the relative humidity drop below 30%, SC contractor will be consulted prior to any work beginning in the vicinity of 

the 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 following 

encapsulation through launch. 

Page 3-7



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 and 

strict control of material cleanliness of flight hardware in proximity to the spacecraft. During transportation using ILS 

provided air-conditioning systems and while on the pad, air is filtered to provide better than Class 100 000 particle 

content. Air cleanliness is monitored regularly in all areas where the SC is present to ensure particle count levels are 

maintained within specification. In addition, witness plates can be mounted inside the fairing following encapsulation 

to monitor particle fallout inside the fairing up until the day prior to launch. The ground contamination environment 

around the spacecraft meets the cleanliness levels specified in Table 3.2.1-1. 

Table 3.2.1-1: Ground Contamination Environment 

Location/event Cleanliness Level 

Required* 

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 during 

transportation and battery charging 

100 000 Payload encapsulated, filtered air provided 

Erection 100 000 Payload compartment sealed 

On-pad with air-conditioning through 

umbilical 

On-pad following removal of airconditioning 

umbilical 

* Per FED Std 209E 

3.2.2 In Flight Contamination Control 

100 000 Filtered air provided 

100 000 Payload compartment sealed 

The Launch Vehicle systems are designed to preclude in-flight contamination of the SC. The Launch Vehicle 

pyrotechnic devices near the spacecraft used for fairing jettison and SC separation have sealed gas chambers and do not 

release significant contamination to the outside environment. The fairing liquid thermal control system pipes are 

sealed by automatic valves which close at fairing jettison. The third stage retro rocket plume does not result in any 

significant particle contact with the spacecraft due to their position on the aft end of the third stage and the orientation 

of the retro rocket plume 15 degrees away from the Launch Vehicle longitudinal axis. The thrusters are located on the 

aft section of the Fourth Stage and oriented perpendicular to the attitude control longitudinal axis (worst case) such 

that the plume does not contact the spacecraft while the spacecraft is attached to the Fourth Stage. Following 

separation, the Fourth Stage is attitude controlled to prevent reorientation relative to the spacecraft until the spacecraft 

is a safe distance away from the Fourth Stage. Finally, an evaporative cooling system in the Fourth Stage ejects a 20 

percent 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 3-8



3.3 PRESSURE 

3.3.1 Payload Compartment Venting 

During ascent, the payload compartment is vented through 4 venting orifices distributed equally about the cylindrical 

portion of the fairing. Maximum rate of pressure drop in the fairing will not exceed 3.5 kPa/s. A representative 

pressure 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 the 

Customer. 

Figure 3.3.1-1: Typical Venting Profile During Ascent 

Pascals 

1.20E+05 

1.00E+05 

8.00E+04 

6.00E+04 

4.00E+04 

2.00E+04 

0.00E+00 

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

Time, seconds 

Time, sec Pressure (Pascals) 

Atmosphere Inside Fairing 

0 100000 100000 

10 98000 99000 

20 91000 93000 

30 80000 83000 

40 65000 70000 

50 49000 55000 

60 32500 38500 

70 19300 26500 

80 10000 17000 

90 5000 11000 

100 2500 7500 

110 1000 5000 

120 0 3500 

130 0 2000 

Atmosphere 

Inside Fairing 

Page 3-9



3.4 MECHANICAL LOADS 

The spacecraft is subject to various types of mechanical loads due to transportation and handling during the launch 

campaign as well as due to the various flight events following liftoff. This section breaks down these events by type of 

excitation: 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 primary 

structure and/or evaluation of compatibility for existing spacecraft with the Proton launch vehicle. The load factors 

were derived for application at the center of gravity of a rigid spacecraft and generate a conservative estimate of flight 

maximum interface axial, shear, and bending moments. The actual interface responses experienced during flight are a 

function of the static and dynamic characteristics of the spacecraft, but these load factors have generally proven 

conservative for spacecraft in the weight range of 2000 - 4800 kg with c.g. height (above the interface) between 1.0 and 

m − z 

2.0 meters and where > 0. 

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

I 

spacecraft cantilevered fundamental frequencies are assumed to be a minimum of 10 Hz lateral and 25 Hz axial to 

insure applicability of both the ground Transport and the Flight load factors. Spacecraft that do not meet these criteria 

will require preliminary analyses to generate loads environments for assessing compatibility with Proton. The Proton 

load factors are conservative, but do not include uncertainty factors. It is recommended that the spacecraft analyst use 

uncertainty factors in preliminary sizing of primary structure if the new design is not within the family of spacecraft 

integrated on Proton. 

3.4.1.1 Ground Loads 

During ground transportation and handling operations, the spacecraft will be subjected to low frequency loads. Table 

3.4.1.1-1 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 Factor 

X 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 3-10



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 produce 

the worst case loading are liftoff and 1st/2nd stage separation. Figure 3.4.1.2-1 provides the bounding quasistatic loads 

in 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 3σ 

values are as provided in Figure 3.4.1.2-2. Flight loads shall be evaluated by Coupled Loads Analysis using a dynamic 

model incorporating the spacecraft and launch vehicle. 

Figure 3.4.1.2-1: Quasi-Static Design Load Factors 

(Axial) 

Londitudinal, g 

5 

4 

3 

2 

1 

0 

-3 -2 -1 0 1 2 3 

-1 

-2 

-3 

-4 

Lateral, g 

Block DM 

Breeze M 

Page 3-11



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 -2.3 

0.3 2.3 -2.3 

Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.2 

1st/2nd stages before separation 4.3 1.4 -1.4 

1st/2nd stages after separation (max 

compression) 3 1.4 -1.4 


tension) -2.8 1.4 -1.4 

2nd/3rd stages separation 3 0.3 -0.3 

3rd/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 -1.35 

0.3 1.35 -1.35 

Maximum Dynamic Pressure (Qmax) 2.2 1.2 -1.2 

1st/2nd stages before separation 4.3 0.9 -0.9 

1st/2nd stages after separation (max 3 0.9 -0.9 


tension) -3 0.9 -0.9 

2nd/3rd stages separation 3 0.3 -0.3 

3rd/4th stages separation 2.8 0.3 -0.3 

Page 3-12



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

(Axial) Longitudinal,g 

5 

4 

3 

2 

1 

0 

-3 -2 -1 0 1 2 3 

-1 

-2 

-3 

Transverse,g 

Event Longitudinal Longitudinal Transverse 

Static g Dynamic g Dynamic g 

Liftoff 

Wind and Blast 

1.5 1.5 -1.5 1.1 -1.1 

(Qmax) 2.2 0.5 -0.5 

Separation 1st/2nd 

Before 1st stage 

booster separation 3.6 0.9 -0.9 0.9 -0.9 

stages 

After 1st stage 

booster separation 1 2 -2.8 0.9 -0.9 

2nd stage engine cutoff 2.7 0.3 -0.3 0.3 -0.3 

3rd stage engine cutoff 2.5 0.3 -0.3 0.5 -0.5 

Note: Used only to evaluate measured flight loads. 

Page 3-13



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 propellant 

in 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 and 

unequal 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 are 

magnified 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 loads 

generated by the jetstream. Launch Vehicle longitudinal flexible oscillations are produced at 10-12 Hz by the natural 

random pulsation of the engine thrust. There is no pogo phenomenon. The maximum value of these oscillations based 

on 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. Because 

of the uneven thrust of the four engines, lateral reaction forces are generated, causing lateral flexible oscillations of the 

Launch Vehicle body. These oscillations are influenced additionally by the first stage engines reacting to control system 

commands. The first stage cutoff is characterized by an abrupt decay from 90% to 20% within 0.03 sec which causes 

significant flexible longitudinal oscillations of the Launch Vehicle second stage, driven by the preliminary thrust of its 

own engines. The oscillations are additionally magnified due to the increase in thrust to 100%. These oscillations damp 

out within about 3 seconds. 

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

preceding events. 

The above dynamic load environment can be represented by the quasi-sinusoidal vibration environment applied at the 

SC/LV interface plane shown in Figure 3.4.2-1. This can be considered a flight limit load environment. 

Page 3-14



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

Lateral Axes 

Level (g's) 

1.2 

1 

0.8 

0.6 

0.4 

0.2 

0 

Lateral Axes 

0 20 40 60 80 100 120 

Longitudinal Axes 

Level (g’s) 

1.2 

1 

0.8 

0.6 

0.4 

0.2 

0 

Frequency (hz) 

Longitudinal Axes 

0 20 40 60 80 100 120 


Frequency (hz) Level (g's) 

5 10 0.3 

10 20 0.4 

20 40 0.6 

40 100 0.7 

Frequency (hz) Level (g's) 

5 8 0.3 

8 20 0.8 

20 40 0.6 

40 100 0.9 

Page 3-15



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

different transport configurations are as shown in the following Figures 3.4.2-2, 3.4.2-3, 3.4.2-4 and 3.4.2-5. 

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

Ascent Unit 

Power Spectral Density (G2/hz) 

3.50E-03 

3.00E-03 

2.50E-03 

2.00E-03 

1.50E-03 

1.00E-03 

5.00E-04 

0.00E+00 

1 10 

Frequency (Hz) 

100 

Frequency (hz) PSD (G 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 

2 /hz) 

8 0.002000 0.003200 0.000660 

10 0.000600 0.003200 0.000800 

14 0.000280 0.000833 0.000330 

20 0.000275 0.000150 0.000320 

25 0.000275 0.000150 0.000310 

30 0.000275 0.000150 0.000300 *Durations are as follows: 

- 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 

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 minutes 

Duration 

(minutes) * * * 

X-X 

Y-Y 

Z-Z 

Page 3-16



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

Transportation of Ascent Unit 

y 

y 

y 

x 

x 

x 

Page 3-17



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

Power Spectral Density (G2 /hz) 

2.00E-05 

1.80E-05 

1.60E-05 

1.40E-05 

1.20E-05 

1.00E-05 

8.00E-06 

6.00E-06 

4.00E-06 

2.00E-06 

0.00E+00 

1 10 100 



PSD (G 

X-X Y-Y Z-Z 

2 0.000002 0.000002 0.000002 

4 0.000002 0.000002 0.000002 

8 0.000002 0.000002 0.000002 

10 0.000002 0.000002 0.000002 

14 0.000002 0.000020 0.000002 

20 0.000010 0.000001 0.000010 

25 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 

Duration 

0.000001 0.000001 0.000001 Levels are at SC to L/V interface 

(minutes) 10 10 10 

2 /hz) 

X-X 

Y-Y 

Z-Z 

Page 3-18



Figure 3.4.2-4: Transportation of Integrated Proton LV 

3.4.3 Acoustic Loads 

The launch acoustic loads arise from acoustic sound waves generated by the supersonic jets from the first-stage engine 

nozzles 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 acoustic 

loads 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. 

y 

x 

Page 3-19



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

Level (dB) 

140 

130 

120 

110 

100 

10 100 1000 10000 

Frequncy (Hz) 

1/3 Octave Band Center Acoustic Levels on 


Spacecraft (dB) 

25 117 

31.5 123 

40 127 

50 126.5 

63 127.8 

80 131.6 

100 132.4 

125 131.3 

160 132.1 

200 132.1 

250 132.1 

315 129.9 

400 129 

500 127 

630 124 

800 121 

1000 119 

1250 117 

1600 114.5 

2000 112.5 

2500 111 

3150 109 

4000 108 

5000 107 

6300 105.5 

8000 104 

10000 103.5 

OASPL 141.4 

Page 3-20



3.4.4 Shock Loads 

Worst case shock levels are introduced into the spacecraft during the firing of the SC/adapter separation system. The 

level is dependent on the type of clampband and the clampband tension. For the existing standard adapter 

configurations, three specific levels may be encountered as indicated in Figure 3.4.4-1. The 1194 separation systems 

use either a 26.6 kN or a 40 kN preload and the shock levels differ accordingly. The 1666 separation systems use a 30 

kN 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 Table 

3.4.4-1. 

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 3-21



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

g’s 

10000 

1000 

100 

100 1000 10000 


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 

Q= 10 

40 kN 1194 Sep Syst. 

26.6 kN 1194 Sep Syst. 

30 kN 1666 Sep Syst. 

Page 3-22



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 must 

demonstrate the capability of the structure to withstand the worst case combination of quasi-static loads shown in 

Table 3.4.1.1-1 and Figure 3.4.1.2-1 or obtained from coupled loads analysis. Qualification margins for structure 

static testing of 1.1 to yield and 1.25 to ultimate must be assumed. For ground handling lift points, qualification 

margin to ultimate shall be 1.5 minimum. 

Demonstration of the spacecraft secondary structures ability to withstand dynamic loads induced by flight events and 

ground 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 sweep 

protoqualification 

Levels in Figure 3.4.2-1 x 1.25 4 

Sine sweep acceptance Levels in Figure 3.4.2-1 4 

Page 3-23



Notching should be minimized, and as a test goal should not allow base inputs to decrease below the base equivalent 

level produced by the Final Coupled Loads Analysis multiplied by the appropriate Factor of Safety and should not 

allow spacecraft response to go below the worst case response predicted by the Final Coupled Loads Analysis 

multiplied by the appropriate Factor of Safety. 

For Spacecraft secondary structure which does not attain the worst case level predicted from the Final Coupled Loads 

Analysis multiplied by the appropriate Factor of Safety, an analysis must show the capability to sustain the CLA result 

multiplied 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’s 

facility in conjunction with a mechanical/electrical fitcheck for first of a kind spacecraft and the first follow-on 

spacecraft in a series. For this test the flight band will be tensioned to the same level as will be used for the flight 

installation per the manufacturers test procedure. Shock levels will be measured at a location on the adapter 120 mm 

above the SC separation plane. 

3.5 ELECTROMAGNETIC COMPATIBILITY 

Electromagnetic Interference (EMI) emissions and susceptibility of the SC and the LV shall be individually controlled 

to 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-fire 

Threshold) 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 be 

compatible with these emissions. 

The launch vehicle generated and launch base spurious EMI sources shall not exceed the levels of Figures 3.5.2-1a and 

3.5.2-1b 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 and 

parallel to the SC/LV interface plane. 

Page 3-24



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

Parameters First Stage 

Transmitter 1 

First Stage 

Transmitter 2 

Second Stage 

Transmitter 

Third Stage 

Transmitter 1 

Third Stage 

Transmitter 2 

Block DM 

Transmitter 

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

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

Block DM 

Receiver 

Modulation type and 

characteristics 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 the 

carrier freq (dbW), Nom. -137 

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

Antenna description and 

polarization 

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 Stage 

Transmitter 1 

First Stage 

Transmitter 2 

Second Stage 

Transmitter 

Third Stage 

Transmitter 1 

Third Stage 

Transmitter 2 

Carrier Frequency (MHz) 192 137 240 232 132 3410 

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


characteristics PM APM/FM PM PM APM/FM PM 


(dbW) Max (at carrier freq.) 20.8 11.8 20.8 20.8 11.8 9.0 


carrier freq (dbW), Nom. 

Tracking 

Antenna gain, dB -10 - 10 - 10 - 10 - 10 -9 within angles of 

+70 degrees 

-10 within angles of 

+85 degrees 


polarization 

Omni-directional, circular limited view omni 

antenna 

Operating on pad? Yes Yes Yes Yes Yes No 

Operating in flight? Yes Yes Yes Yes Yes Yes 

Page 3-25



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

Parameters Breeze M 

Telemetry 

Breeze M 

Telemetry 

Breeze M 

Telemetry 

Breeze M 

Receiver 

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 


characteristics PM FM FM PM PM 


(dbW) Max (at carrier freq.) 9.0 11.8 11.8 


carrier freq (dbW), Nom. - 137 - 137 

Antenna gain , dB 21 3 to 10 

(Parallel to 

Xsc axis) 


polarization 

right circular 

horn 

limited view 

omni antenna 

3 to 10 

(Parallel to 

Xsc axis) 

limited view 

omni antenna 

-7 3 

limited view 

omni antenna 

limited view 

omni antenna 

Operating on pad? No Yes Yes No Yes 

Operating in flight? Yes Yes Yes Yes Yes 

Page 3-26



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

dBm V/m 

140 

120 

100 

80 

60 

40 

20 

0 

100 1000 10000 100000 

Frequency (MHz) 

Frequency (MHz) dBmV/m 3dB Band width (MHz) 

132 98 3 

137 74 3 

192 84 0.256 

232 100 0.256 

240 87 0.256 

769 25 0.5 

923 120 0.5 

2000 70 N/A 

100000 15 N/A 

Page 3-27



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

dBm V/m 

140 

120 

100 

80 

60 

40 

20 

0 

100 1000 10000 100000 

Frequency (MHz) 

Frequency (MHz) dBmV/m 3dB Band width (MHz) 

132 98 3 

137 74 3 

192 84 0.256 

232 100 0.256 

240 87 0.256 

1000 127 0.5 

2000 70 N/A 

3400 123 0.5 

15150 135 0.5 

30000 15 N/A 

Page 3-28



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

dBmV/m 

130 

110 

90 

70 

50 

30 

10 

1.00E-01 1.00E+00 1.00E+01 1.00E+02 

Frequency (GHz) 

Frequency (GHz) dBmV/m 

0.01 120 

0.1 120 

0.75 60 

1 129 

20 80 

3.5.3 RF Transmitter/Receiver Systems EMC 

Standard RF system compatibility analyses will be performed which shall insure integrated system EMC of 

simultaneously operated SC and LV transmitters and receivers during time frames where such operations are necessary. 

Page 3-29



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 in 

Appendix 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. The 

clearance between the flanges of the SC and the adapter prior to clampband tensioning shall not exceed 0.6 mm. The 

geometry of the spacecraft flange is provided in Appendix D for a temperature of 21° C. The surface flatness of the SC 

interface 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 mode 

frequencies shall be greater than 10 Hz and 25 Hz respectively as cantilevered from a rigid interface. The SC/LV 

interface is assumed to behave linearly under all loading conditions. 

Page 4-1



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 separation 

system. 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 umbilical 

connectors. The adapter spring forces and the SC/LV electrical umbilical connector forces are provided in the adapter 

Appendix 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 door 

locations 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 of 

the 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 and 

4.1.3.1-6 provide the layouts for the Proton/Breeze M Standard and Long Commercial fairing generic useable 

volumes. These generic useable volumes do not take into account any specific adapter configuration. Specific adapters 

will alter the bottom portion of the useable volume in order to take into account required adapter clearances for 

installation and required flight clearances with the adapter structure. Specific useable volumes tailored to individual 

adapter 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, including 

manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the useable volume, except 

where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements due to ground or flight 

loads and deviations caused by an imperfect installation of the spacecraft on the Fourth Stage may protrude beyond the 

boundaries of this useable volume. An assumption made is that spacecraft dynamic displacements will not exceed 50 

mm. This must be verified by coupled loads analysis. 

Page 4-2



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 are 

dimensioned as shown in Figure 4.1.3.1-1. These doors are nominally used for access to the Block DM. The Customer 

may use these doors for access to spacecraft related interface equipment. These access requirements need to be 

coordinated and agreed upon with ILS in the mission specific ICD. Up to 2 access doors may be provided in the 

fairing in the locations shown in Figures 4.1.3.1-1, 4.1.3.1-3 and 4.1.3.1-5 for the Block DM and Breeze M vehicle 

versions. These access locations may be accessed at times coordinated with ILS from the time of fairing encapsulation 

up to the beginning of Launch Vehicle fueling on the launch pad. 

Page 4-3



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

Page 4-4



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

Page 4-5



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

Page 4-6



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

Page 4-7



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

Page 4-8



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

Page 4-9



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

Page 4-10



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

Page 4-11



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

Page 4-12



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

Page 4-13



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

Page 4-14



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

Page 4-15



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

Page 4-16



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

Page 4-17



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

Page 4-18



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

Page 4-19



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

Page 4-20



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

Page 4-21



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

Page 4-22



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 pneumatic 

fittings. GN2 can be provided via Customer provided gas bottles during operations in the PPF and on the Launch Pad 

up to Mobile Service Tower rollback. At this time, the line can be connected to a dry air source running through the 

LV 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 Mobile 

Service Tower rollback (including during transportation operations as mutually agreed upon 

between ILS and Customer) 

Operational gas Gaseous nitrogen or air 

Gas Content 

Particulate Size



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 power 

supplies located in the Vault under the launch pad. They are also used for hardline telemetry and command links 

between the spacecraft and the Customer ground support equipment located in the Bunker. Additionally, the Customer 

has an option to provide spacecraft measurements to the Launch Vehicle telemetry system on the Fourth Stage via 

these umbilicals. 

4.2.1.1 Electrical Connectors 

Two 37 or 61 umbilical connectors are provided at the Spacecraft interface with the Launch Vehicle adapter. The 

connectors 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 and 

mechanical configuration for these connectors. 

Umbilical connector brackets provide +/- 2 mm adjustment in longitudinal direction and +/- 4 mm adjustment in 

the lateral axes. 

4.2.1.2 Separation Verification 

Two diametrically opposed separation microswitches are provided on the top adapter interface flange. Refer to 

Appendix D for specific locations and mounting configuration for each specific adapter. At separation, the 

microswitches 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, the 

umbilical connectors will disengage, thereby opening these circuits and providing a redundant indication of separation 

to 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 there 

is 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 until 

stage three/four separation. Their positions are illustrated in Section 4.2.1.7. Three accelerometers measure 

longitudinal loads and two measure lateral loads. Refer to Section 4.2.1.7 for characteristics of these telemetry 

channels 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 to 

Appendix D for angular locations and for cross-section views. 

Page 4-24



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 of 

relay closures for initiating spacecraft commands during flight. The command for closure will be issued after launch 

and before SC/LV separation. Timing and signal characteristic requirements need to be provided by the Customer no 

later than at L-12 months. Characteristics of this command are as follows: 


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 interface 

contact 

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 

PO2A 

K7 K13 K8 K14 

K11 

K1 

3 

K7 

K11 

K14 K8 

K9 

K12 

K1 

5 

2 6 1 5 11 15 12 16 

J1 J2 

K15 K10 

K9 

K12 

K1 

6 

K1 

6 

K10 

SC/LV Interface 

Page 4-25



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 provides 

the characteristics and location of each flight telemetry sensor. Finally, Figures 4.2.1.7-1 to 4.2.1.7-9 show the 

locations 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 spacecraft 

telemetry points through the umbilicals to the Third and Fourth Stage telemetry system. Characteristics are described 

below: 


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 to 

spacecraft TM point (ohms) to maintain measurement error 

within +/- 7% 

Minimum isolation required between signal line and spacecraft 

structural ground (Megohms) 

2 kohm (any combination of resistance, impedance and 

inductance) 

Data recorded for these TM points will be provided to the Customer as part of the Post-Launch Report submission 

and in electronic format. 

TBD 

Page 4-26



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

Accelerations 

Accelerometer Location and 

Measurement Directions 

Amplitude 

Measurement 

Dynamic 

Range, G’s 

Frequency 

Measurement 

Range, Hz 

Transport by Motor Vehicle or Location: In the area of the attachment of the container 

Rail, Spacecraft Mounted in on shock pallet to the Transport Vehicle 

Shipping Container on Shock Vertical 

1.0 (TBX1) Up to 50 Hz 

Pallet 

Lateral 

1.0 (TBY1) 

Longitudinal 

1.0 (TBZ1) 

Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface Ring 

Mounted on LV Fourth Stage Vertical 

1.0 (TBX3) Up to 50 Hz 

(Fourth Stage and Fairing Lateral 

1.0 (TBY3) 

Only) 

Longitudinal 

Support Point of Fairing Assembly at Cylinder-Nose 

Cone Transition 

1.0 (TBZ3) 

Vertical 

1.0 (TBX4) Up to 50 Hz 

Lateral 

1.0 (TBY4) 

Longitudinal 

Spacecraft to SCA Separation Joint Interface 

1.0 (TBZ4) 

Vertical 

1.0 (TBX) Up to 50 Hz 

Lateral 

1.0 (TBY) 

Longitudinal 

1.0 (TBZ) 

Transport by Rail, Spacecraft Support Point of Fourth Stage Aft Interface Ring 

Mounted on Proton Launch Vertical 

1.0 (TBX6) Up to 50 Hz 

Vehicle Assembly 

Lateral 

1.0 (TBY6) 

Longitudinal 

Support Point of Proton First Stage at Aft Ring 

1.0 (TBZ6) 

Vertical 

1.0 (TBX5) Up to 50 Hz 

Lateral 

1.0 (TBY5) 

Longitudinal 

Spacecraft to SCA Separation Joint Interface 

1.0 (TBZ5) 

Vertical 

1.0 (TBX) Up to 50 Hz 

Lateral 

1.0 (TBY) 

Longitudinal 

1.0 (TBZ) 

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 conditioning 

car 

0-50 °C 

On Pad Temperature on Adapter 

-50 to 80 °C (TA5, TA6) 

Temperature of Air Supplied to Fairing 

-50 to 80 °C (T20, T21) 

Page 4-27



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 car 

0-80% 

Contamination Contamination Sensors Location Measurement Range, 

All ground events Particulate size at inlet/exit from air conditioning car (see 0.5 microns/5 microns and 

facility drwg) 

higher 

All ground events Witness Plates (2) located inside fairing on boat-tail section 

On Pad Access to fairing air supply for manual reading of 0.5 microns/5 microns and 

contamination levels 

higher (SENSOR LABEL**) 

Page 4-28



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

Purpose Quantity, measurement 

range 

Temperature sensors 

Environment 

under fairing 

Ext. surface 

thermal 

insulation 

External 

surfaces of the 

cellular 

construction 

skin and 

thermal 

insulation 

1 0...150C 

2 -40...100C 

2 -40...100C 

1 -40...600C 

1 -40...500C 

1 -40...200C 

1 -40...200C 

1 -40...200C 

2 -40...200C 

1 -40...200C 

1 -40...200C 

1 -40...200C 

1 -40...200C 

2 -40...200C 

1 -40...200C 

1 -40...200C 

1 -40...100C 

1 -40...100C 

2 -40...100C 

1 -40...100C 

1 -40...100C 

Sensor designation and location Time of operation Sensor 

sampling 

frequency 

T1 Nose internal surface 

T22 (T20) Cylinder, IV pl. 

T23 (T21) Cylinder, II pl. 

T26 30 degree cone 


T27 Cylinder 

External skin 



T5,T4 Cylinder I pl. 

T6 Cylinder III pl. 

T7 Inverse cone I pl. 

Internal Skin 

T8 30 deg. cone I pl. 





Thermal insulation 




T18 Cylinder III pl. 


Panel liquid thermoregulating system 

heater (LTS) 



1 -40...100C 

1 -40...100C 

Construction 4 -10...77C TA1 

of adapters 

TA2 adapter 

TA3 

TA4 

upper frame 

2 -90...100C TA5 

TA6 adapter middle 

4 -90...100C TA7 

TA8 adapter bottom 

TA9 

TA10 

frame 

Spacecraft separation sensors 

Adapter 2 

Separation 

plane 

J1, J2 

2 

From liftoff command 

to nose fairing separation 

(Temperature sensors in 

parenthesis are used on 

the ground only). 


to spacecraft separation 

To liftoff command 

(ground sensors) 


to spacecraft separation 

To spacecraft 

separation 

0.3 Hz 

1.6 Hz 

continuous 

1.6 Hz 

50-100 Hz 

Page 4-29



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


range 

Nose fairing separation sensors 

Nose fairing 2 

1-2 

4 

5-6 

7-8 

Pressure sensors 

4 

1-2 

3-4 

Nose fairing 4 0...780 1 20 degree cone, 

mm Hg 2 

3 

4 

external surface 

1: 0...400 mm Hg 5 Inverse cone, 

1: 0...780 mm Hg 6 external surface 

1: 0...780 mm Hg 1 20 degree cone, 

1: 0...250 mm Hg 2 internal surface 

1: 0...780 mm Hg 3 Inverse cone, 

1: 0...50 mm Hg 4 internal surface 

4: 0...780 mm Hg 1 

6 

8 

10 

Above vents 

7: 0...780 mm Hg 2 

3 

4 

5 

7 

9 

11 

Below vents 

Acoustic loads sensors 

Nose fairing 

internal and 

external 

acoustic 

pressure 

measurement 

2 places 

2 places 

30-2000 Hz 

120-155 dB 

30-2000 Hz 

125-165 dB 


sampling 

frequency 

AB-1 In the internal 

AB-2 volume between 

SC and nose 

fairing. AB-1 on KhSC 

adapter, AB-2 on fairing 

AH-1 On the nose 

AH-2 fairing external 

To nose fairing 

separation 

From liftoff command to 

nose fairing separation 


first stage separation 

50-100 Hz 

50 Hz 

8000 Hz 

Page 4-30



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


range 

Vibration loads sensors (low frequency) 

Low frequency 

vibration 

measurement 

along X, Y, and 

Z axes 

3 places along X 

up to 32 Hz 

from -2...+4 g 

2 places along Y & Z 

up to 32 Hz 

±0.6 g 

Vibration loads sensors (high frequency) 

High 

frequency 

vibration 

measurement 

along X and Y 

axes 

2 places 

15...2000 Hz 

up to 8 g 

up to 8 g 


sampling 

frequency 

KX-1 On 

KX-2 adapter near 

KX-3 SC interface 

KY-4 

KZ-5 

KX-1 

KX-2 

KX-3 

KY-4 

KZ-5 

BX-CT 

BY-CT 

On adapter 


third stage separation 

At time of each stage 

separation 


third stage separation 

200 Hz 

200 Hz 

200 Hz 

200 Hz 

200 Hz 

400 Hz 

200 Hz 

200 Hz 

100 Hz 

100 Hz 

8000 Hz 

Page 4-31



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

Page 4-32



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

TBS 

Page 4-33



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

Page 4-34



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

TBS 

Page 4-35



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) 

Thermal Rail Car 

Air Temperature 

Relative Humidity 

Particle Count 

Conditioned Air Inlet 

TBX1, TBY1, TBZ1 

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

Conditioned Air Exhaust 

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

Area 31 only) 







Exhaust to Atmosphere (2) 

T21 

(internal to container 

opposite to T20) 

T20 

(internal to container 

opposite to T21) 

y 

y 

x 

Page 4-36 

x



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. 


Air Temperature (inlet, exit) 

Relative Humidity (inlet, exit) 

Particle Count (inlet, exit) 


TBX, TBY, TBZ 


TA5, TA6 

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

T20, T21 



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

Exhaust to Atmosphere 

TBX, TBY, TBZ Conditioned Air Inlet (2) 

TA5, TA6 

T20, T21 

Witness Plates 

y 

x 




y 

Page 4-37 

x



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. 


TBX, TBY,TBZ 

Filter Block 

T20, T21 



Air Temperature (inlet, exit) 

Relative Humidity (inlet, exit) 

Particle Count (inlet,exit) 



TA5, TA6 

Page 4-38



Figure 4.2.1.7-9: Instrumentation During On-Pad Operations 

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

Temperature inside 

Fairing, T20 and T21 

(see table for description, 

see fairing schematic 

for exact location). 

Temperature on Adapter, 

TA5 and TA6 

(see table for description, 

see fairing schematic 

for exact location). 

Fairing Air Exhaust to Atmosphere 

(two at 420 x 500) 


(access door on Fairing boat tail) 

Conditioned Air Inlet to Fairing 

(Air sample access: continuous 

monitoring of air used to measure 

temp., humidity, and contamination 

levels) 

AIR 

THERM 

AL CONTROL 

SYSTEM 

(ATCS) 

Temperature 

Humidity 


Page 4-39



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 adapter 

which is connected to the LV flight umbilical harness running the length of the vehicle to an interface connector Ø06 at 

the bottom of the first stage. From here, ground cabling connects the umbilical to an interface panel in the Vault under 

the launch pad where the Customer electrical interface equipment is located. As can be seen from Figure 4.2.2-1, there 

are test access connectors on the Fourth Stage which permit access to the umbilical from the Mobile Service Tower up 

to 8 hours prior to launch. These can be used to interface Customer battery charging power supplies on the Mobile 

Service Tower with the SC. They can also be used to connect with wiring in the Mobile Service Tower to provide a 

parallel path with the flight LV umbilical to reduce overall resistance drop from the SC to Customer GSE for high 

current power lines. 

The Launch Pad interfaces include connections from the base of the Proton LV (and connections at station 43.85 on 

the Mobile Service Tower if required) to ground wiring interfacing with SC EGSE. ILS provides all necessary 

electrical harnesses and cables between the SC/LV IFD’s and the SC EGSE interface enables in the Vault and on the 

MST. Figure 4.2.2-1 provides a block diagram of the electrical interfaces available between the payload, Launch 

Vehicle and the ground systems. 

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

O 1A 

O 06 

Spacecraft 

37 or 61 pin Deutsch 

Connectors 

Mobile Service 

Tower 

Mission Specific 

Wiring 

Level 6 

Junction 

Box X1 

Mission Specific Wiring 

Vault 

Customer 

MPS 

Notes: 1) Cable length from level 6 of MST to Vault: 400 meters 

2) Cable length from SC connectors to O 06: 50 meters 

3) Cable length from O 06 to Vault: TBD meters 

Bunker Rm 250 

Customer 

STE 

Page 4-40



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 also 

needs 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 and 

airborne systems in the event of a short circuit . SC STE shall shut off power within 0.2 sec if I max is exceeded by 50%. 

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

flight mating. 

All SC EGSE electrical interface circuits shall be constrained at least 5 minutes prior to liftoff such that there will be 

no 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 choose 

between one of the 4 links defined in Tables 4.2.3-1a, 4.2.3-1b, 4.2.3-1c and 4.2.3-1d. A block diagram of the SC to 

bunker 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 hour 

session. 

d) The SC contractor shall provide to Khrunichev the measured coefficient values for TT &C signals via the RF 

window obtained during the RF channel checkup in the integration facility following the Ascent Unit 

encapsulation. 

e) Prior to installation of the LV+Ascent Unit on the pad and following the delivery of the STE to the bunker, the 

SC manufacturer shall verify continuity between command RF link and STE and issue to Khrunichev the 

Certificate 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, in 

conjunction with Khrunichev, shall check out the RF link between the Ascent Unit and STE. Such check out shall 

be performed 20 minutes after the mating of the LV aft section. SC contractor to confirm functionality of the RF 

link within 45 minutes. 

g) At L-6 months, the SC manufacturer shall provide to Khrunichev two connectors for installation by Khrunichev 

on the existing RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructions 

on cable dressing and cable performances. 

Page 4-41



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

Telemetry Link 

Reference Value Note 

Frequency range, GHz 3.95 ± 0.2 

Bandwidth, MHz > 250 

Signal polarization 

Radio link output signal power 

Left-hand circular 

maximum. dBm minus 0.0 with Service Tower rolled back 

minus 6.2 with Service Tower in place 

minimum, dBm minus 30.0 with Service Tower rolled back 


Radio link gain factor, dB minus 31.0 ±5 with Service Tower rolled back 

minus 42.2 ± 2 with Service Tower in place 

Gain factor adjustment limit for 

30 

radio link input, dB 

Radio link output signal-to-noise 

64 

ratio, dB*Hz 

Command Link 






Right-hand circular 

maximum, dBW/m2 minus 22.2 with Service Tower rolled back 


minimum, dBW/m2 minus 72.2 with Service Tower rolled back 


Radio link gain factor, dB minus 43.2 with Service Tower rolled back 


Gain factor adjustment limit for radio 

link input, dB 

30 

Radio link output signal-to-noise ratio, 

dB*Hz 

70 

Page 4-42



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





Signal polarization Linear vertical 

Radio link output signal power with SC antenna input signal power of 0 dBW 

maximum, dBm minus 31 with Service Tower rolled back 

minus 37 with Service Tower in place 

minimum, dBm minus 41 with Service Tower rolled back 


Radio link gain factor, dB minus 78.3 ± 5 with Service Tower rolled back 

minus 80.0 ± 2 with Service Tower in place 



30 


dB*Hz 

118 

Command Link 




Signal polarization Linear, horizontal 

Radio link output signal power with SCS antenna input signal power of 

3 dBW 







Gain factor adjustment limits for radio from -65 to -41 


from -67 to -43 


dB*Hz 

123 

Page 4-43



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





Signal polarization left-hand circular 

Radio link output signal power with SC antenna input signal power of 0 dBW 

maximum, dBm 2 with Service Tower rolled back 


minimum, dBm minus 8 with Service Tower rolled back 


Radio link gain factor, dB minus 35 ± 5 with Service Tower rolled back 

minus 45 ± 2 with Service Tower in place 



30 


dB*Hz 

118 

Command Link 




Signal polarization Right-hand circular 

Radio link output signal power with SCS antenna input signal power of 3 dBW 

maximum, dBW/m2 minus 31 with Service Tower rolled back 


minimum, dBW/m2 minus 41 with Service Tower rolled back 




Gain factor adjustment limits for radio from -50 to -80 


from -52 to -82 


dB*Hz 

123 

Page 4-44



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







Linear horizontal 

maximum, dBm minus 8.0 with Service Tower rolled back 


minimum, dBm minus 44.0 with Service Tower rolled back 






30 


dB*Hz 

118 

Command Link 






Linear vertical 







Gain factor adjustment limit for radio link 

input, dB 

30 


dB*Hz 

155 

Page 4-45



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

A048 

Launc h Pad 

SC 

Underground Vault 

Interconnect 

Panel 

4.2.3.1 RF Window Requirements 

RF link 

Interconnect 

Panel 

Interconnect 

Panel 

RF TM 

RF TC 

Bunker 

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. 

Khrunichev 

RF 

Rack 

There are 2 RF window positions in the fairing to take into account the possible view angles required at each of the 2 

Proton launch pads. When the launch pad is designated, 1 out of the 2 windows will be covered with conductive enamel 

leaving 1 active window for transmission of the spacecraft T and C signal between the spacecraft and the Bunker. The 

RF link between the spacecraft and the ground RF equipment is required before and after Mobile Service Tower 

rollback. Prior to rollback, the signals will be transmitted via a repeater on the Mobile Service Tower. Following 

rollback, signals will be transmitted directly between the spacecraft and the Bunker antenna. During rollback, there 

could be a maximum 20 minute outage of signal as the tower rolls through the line of site between the spacecraft and 

the bunker antennas. 

Page 4-46



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 with 

earth referenced steel ground busses with equipment attach points (threaded studs). The resistance between any point 

on these bars and the building earth ground is less than 4 ohms. The floor surfaces in the payload and hazardous 

payload processing areas is anti-static and connected to the facility grounding system. The SC contractor shall provide 

all 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 10 

milliamps. This may be accomplished either by conductive surface contact between the spacecraft and adapter 

interface ring (1666 adapters) or by the use of 2 bonding straps which incorporate a friction contact connector that 

releases upon spacecraft separation with a separation force of 40 ± 5 N. Signal and power grounds from the spacecraft 

are passed through the umbilical without connecting them to Launch Vehicle structure. Likewise umbilical shield 

grounds 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 and 

indirect hits. Augmentation of the standard provisions for any necessary SC individual circuit protection shall be 

provided 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 SC 

surface 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 SC 

manufacturer’s facility for first of a kind spacecraft and the first follow-on spacecraft in a series. 

Page 4-47



5. MISSION INTEGRATION AND MANAGEMENT 

ILS provides the Customer a SOW which define the management approach for a Customer Launch Service 

Agreement, the deliverables provided to the Customer during the course of the LSA and a schedule for all Mission 

Integration 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 Managers 

who 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 to 

perform 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 the 

Customer's technical requirements for the launch of their spacecraft and characteristics and constraints of the Launch 

Vehicle 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 interface 

milestones necessary for successful completion of the contract. A typical mission integration schedule is as shown in 

Figure 5.1.3-1a and 5.3.1-1b for a non-recurring and recurring program respectively. It is based on the meeting 

schedule in Section 5.1.5.1 and the deliverable milestones provided in Sections 6 and 7. This integrated program 

schedule for a particular program will be presented and agreed upon between all Parties at the Kickoff Meeting and 

further changes shall be made as necessary and agreed upon at subsequent Technical Interface Meetings (TIM). In case 

of changes to internal schedules, the other party will be promptly informed. 

Page 5-1



Figure 5.1.3-1a: Baseline Integration Schedule (Non-Recurring Program) 

ID # Activity Resp 

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 

31 

32 

33 

34 

35 

36 

37 

38 

39 

Kickoff 

Preliminary ICD Review 

Preliminary Design Review (PDR) 

Launch Site Visits 

Operations Review 

ICD Sign-Off 

Fitcheck 

Shock Test 

Critical Design Review (CDR) 

Launch Vehicle Preshipment Rev 

Spacecraft Preshipment Review 

Launch Site Acceptance Review 

Launch Vehicle Quality Review 

Update LV Roll-out Authorization Review 

Board LV Loading Authorization Review 

Board 

IRD (Amendment for specific SC) 

Preliminary ICD 

ICD prior to meeting to sign 

Preliminary SC Dynamic Model SCC 

Annex [x] of PDR package, CLA ILS 

Report SC Thermal Model 

SCC 

PDR Package, all except Annex [x] ILS 

Inputs to Final Analysis (except SCC 

CLA) Final SC Dynamic Model (test SCC 

verified) Annex [x] of PDR package; CLA ILS 

Report CDR Package; all except Annex [x] ILS 

SC Environmental Test Plan SCC 

Inputs to Fitcheck & Shock Test PlanSCC 

Fitcheck & Shock Test Plan ILS 

Fitcheck & Shock Test Report ILS 

SC Environmental Test Results SCC 

Safety Submissions 

All Certificates 

SC Launch Operations Plan SCC 

(Preliminary) 

SC Launch Operations Plan (Final) SCC 

Participants List w/ Passport Info SCC 

SC Orbital Data 

ILS 

Launch Evaluation Report 

ILS 

Printed: 31AUG98 

Management Report 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

SCC 

ILS 

ILS 

SCC 

All 

ILS 

STANDARD INTEGRATION SCHEDULE-NON-RECURRING 

Revised: 08/21/98 

L-24 L-21 L-18 L-15 L-12 L-9 L-6 L-3 LAUNCH L+2 

MEETINGS 

Kickoff 

ICD Review 

PDR 

Site Visit No.1 Site Visit No.2 

Operations Review #1 Operations Review #2 

ICD Sign-Off 

Fitcheck 

Shock Test 

CDR 

LV Preship 

SC Preship 

Site Review 

LVQR Update 

LV Roll-out Auth 

LV Loading Auth 

ICD DEVELOPEMENT 


Prelim ICD 

Signed ICD 

PRELIMINARY DESIGN 

Preliminary SC Dynamic Model 

CLA Report (if required) 

SC Thermal Model 

PDR Package 

CRITICAL DESIGN 

Inputs to Final Analysis (except CLA) 

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 



LAUNCH CAMPAIGN 

SC Launch Ops Plan (Prelim) 

SC Launch Ops Plan (final) 

Participants List w/ Passport Info 

SC Orbital Data 

Launch Eval Report 

MANAGEMENT 

Page 5-2



Figure 5.1.3-1b: Baseline Integration Schedule (Recurring Program) 

ID # Activity Resp 

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 


Fitcheck 

Shock Test 

LV Preshipment (Quality) Review 

SC Preshipment Readiness Review 



Launch Vehicle Quality Review 

Update 

LV Roll-out Authorization Review 

Board 

LV Loading Authorization Review 

Board 


ICD prior to meeting to sign 

revision 

Inputs to Final Analysis (except 

CLA) 

Final SC Dynamic Model (Test 

Verified) 

Annex [X] of CDR package; CLA 

Report 

CDR Package; all except Annex [x] 


Inputs to Fitcheck & Shock Test PlanSCC 


Fitcheck & Shock Test Report 




SC Launch Operations Plan (final) 

Participants List w/ Passport Info 

Orbital Data 

Printed: 31AUG98 

Launch Evaluation Report 

Management Report 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

All 

SCC 

ILS 

SCC 

SCC 

ILS 

ILS 

SCC 

ILS 

All 

SCC 

SCC 

ILS 

SCC 

SCC 

ILS 

ILS 

ILS 

MEETINGS 

Kickoff 

ICD DEVELOPEMENT 

IRD 


ICD 

Inputs to Final Analysis 


ICD Sign-Off 


Inputs to Fitcheck & Shock Test Plan 



LAUNCH CAMPAIGN 

MANAGEMENT 

STANDARD INTEGRATION SCHEDULE - RECURRING 

CDR 

Fitcheck 

Shock Test 

Final SC Dynamic Model 

CDR Package 

SC Launch Ops Plan (final) 

Fitcheck & Shock Report 

CLA Report 

LV Preship Rvw 

SC Preship Readiness Rvw 

Ops Review 

Site Review 




Participants & Info List 

Revised: 08/21/98 

L-12 L-10 L-8 L-6 L-4 L-2 LAUNCH L+2 

LVQR Update 

LV Roll-out Auth 

LV Loading Auth 

Orbital Data 

Launch Eval Report 

Page 5-3



5.1.4 Documentation Control and Delivery 

ILS maintains an internal documentation and configuration control system for all LSAs. Deliverable documentation 

shall be maintained under this configuration control system. 

All technical correspondence between ILS and the Customer relating to work on the LSA is strictly between the 

Customer 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 to 

launch preparation of each satellite. A preliminary meeting schedule is defined in Section 5.1.5.1 and meeting 

schedules will be updated through the course of the contract as part of the interface activities milestone schedule 

generated by ILS. Exact dates, locations, agendas, and participation are agreed upon in advance, on a case-by-case 

basis, 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 end 

of 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 recurring 

program respectively. A non-recurring program is one with a first of a kind SC which requires 2 analysis cycles. A 

recurring program is one with a similar SC which requires only one analysis cycle and no significant changes to the 

launch vehicle and the launch site. 

Page 5-4



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 5-5



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 well 

as management organization and preliminary program schedules. The Interface Requirements Document will be 

discussed 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 with 

ICD requirements. 


Review of requirements and corresponding implemention for launch base operations. 

Page 5-6



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's 

facility. 

Shock Test 

This is an actuation of the flight type separation system with the flight spacecraft at the manufacturer's facility. It is 

done in conjunction with the fitcheck. 


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 all 

Launch 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 the 

Launch Site. 


This meeting is held at the Launch Site prior to SC arrival to confirm the readiness of the Launch Site to begin the 

launch 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 into 

account all subsequent activities up to L-7 days before launch (TBC). 

Page 5-7



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 to 

launch. 

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 interchange 

involving 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 progress 

including: subcontract status, technical status, mission integration schedule, production schedule and status for all 

major hardware, action item status, contract deliverable status. Major program issues are summarized and resolution 

plans 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 the 

Customer's spacecraft. 

The Customer will obtain permits and approvals required to import and export the spacecraft and associated 

equipment 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 in 

Table 5.2-1. 

Page 5-8



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 


PDR package PDR -10 days 


CDR package CDR -10 days CDR -10 days 


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 will 

contain 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’s 

launch mission, including those analyses identified below. 

The following analyses are conducted during the mission integration effort for each satellite launch mission. For first 

of a kind spacecraft, one preliminary and one final analysis cycle will normally be conducted during each satellite 

integration effort. For follow-on spacecraft, one analysis cycle will be performed in most cases. Two analysis cycles 

will be performed for following on SC where spacecraft and/or launch vehicle relevant parameters have changed 

significantly. 

Page 5-9



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 previous 

vehicles. Emphasis is on specificities in adapter and payload compartment 

design to meet specific spacecraft payload requirements. 

Mission Design The flight design including maneuvers and maneuver sequence, orbit 

parameters and dispersions, collision avoidance 

Thermal Analysis Integrated thermal analysis of combined operations (ground and flight) for 

spacecraft and launch vehicle hardware to ensure thermal compatibility. 

The spacecraft mathematical model is provided by the Customer per the 

thermal model specification provided by ILS 

Separation Analysis Analysis of spacecraft separation including presentation of pertinent 

kinematic parameters and their dispersions during the separation event 

CLA/Acoustic/ Shock Loads 

Environment 

1) Dynamic coupled-loads analysis. The spacecraft mathematical model is 

furnished by the Customer according to the ILS provided dynamic model 

specification. 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 verification 

coupled loads analysis will be conducted unless significant launch vehicle 

configuration changes have occurred (e.g. if first launch occurs on Block 

DM and next launch occurs with Breeze M). 

2) Presentation of other load environments including acoustic, shock and 

ground transportation loads 

Contamination Analysis of ground and flight contamination sources and effect on 

spacecraft payload 

RF Link and EMC Analysis of the RF link between the bunker and the pad and EMC analysis 

verifying compatibility between the spacecraft and LV systems 

Clearance Analysis Clearance analysis between the spacecraft and the LV during flight to verify 

sufficient dynamic clearances 

Venting Analysis Analysis of fairing depressurization during flight 

Operations Detailed description of how Khrunichev will meet operational 

requirements specified by the Customer in the ICD 

Page 5-10



One hardcopy (and a soft copy where possible) of all design documentation will be provided to Customer 10 days prior 

to review 

Reports will be provided documenting the results of the above analyses. These reports will be provided for each analysis 

cycle and include the following topics: summary of results, detail of analyses performed, comparison of analysis results 

with 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 actions 

for each of the participants including Khrunichev, ILS, the SC manufacturer and SAAB (if applicable). ILS will 

provide 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 the 

spacecraft 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 is 

provided in Tables 5.3-1. 

Page 5-11



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 SC Dynamic Model L-24 

SC Thermal Model L-19 


Final SC Dynamic Model (test verified model) L-7 L-7 

Inputs to Final Analyses (except SC Dynamic Model) L-12 L-11 


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 5-12



5.3.1 ICD Development 

The Customer will provide an IRD to ILS with interface requirements describing all pertinent design information on 

the spacecraft characteristics, mechanical and electrical interfaces, and constraints necessary to define the integration 

tasks 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 second 

analysis 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 SC 

manufacturer to demonstrate compatibility with the Proton ground and flight environments. A summary of the results 

from 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 report 

following 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 for 

ground operations and flight. 

Page 5-13



5.3.4.3 SC Readiness Certificate 

The Customer will provide a certificate to ILS following spacecraft environmental testing to certify that the spacecraft 

is capable of withstanding the loads environments imposed on it during ground processing at the Launch Site and 

during 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 internationally 

recognized 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 required 

passport 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’s 

Guide. 

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 required 

to meet the contracted launch date. 

Page 5-14



5.4.3 Final Spacecraft Data 

For missions that employ the Block DM, the Customer shall also provide final spacecraft estimated dry and wet mass 

two days prior to Block DM fueling. Prior to encapsulation, the Customer will supply ILS the actual satellite dry and 

wet 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 the 

start of Combined Operations; prior to rollout to the pad; and prior to fueling of the Launch Vehicle. These dates will 

be 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 its 

ground 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 launch 

necessary to assist ILS in evaluating the performance of the launch vehicle and associated services provided under each 

LSA. 

5.4.7 Spacecraft Propellants 

The Customer will procure spacecraft propellants to support the launch campaign and is responsible for shipment of 

these propellants to the Port of Entry into Russia (typically St. Petersberg) and through Customs. After the launch 

campaign, the Customer will be responsible for removal of the propellants and associated equipment from the Port of 

Entry. 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 will 

be used for the assembly of flight and test harnesses. 

Page 5-15



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 launch 

site) 

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 technology 

necessary for analysis of SC propellants. The Customer must plan for shipment and analysis of samples outside the 

Russian 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 design 

which 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 requested 

by 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 by 

changes to the spacecraft design which are not in any way attributable to ILS and not required by the terms of the 

LSA and its' Exhibits 

Page 5-16



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). These 

facilities 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 same 

facilities 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 SC 

and GSE, as well as charter flights with campaign personnel. It is an internationally rated airport with a single 4.5 km 

long, 84 m wide landing strip oriented 60 degrees/240 degrees relative to North. The airport has an elevation of 

approximately 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 equipped 

with stationary spotlights for use in night operations. 

Page 6-1



Figure 6.1-1: Baikonur Facilities Map 

92A-50 

Facility 

Bldg 92-1 

Railline: 

Road: 

Hotel 

Complex 

Proton Launch Complex 81 

(Pads 23 & 24) 

Proton Launch 

Complex 200 

(Pads 39 & 40) 

Yubileini Airfield 

Area 254 

Area 31 

Propellant Storage Facility Oxygen and Nitrogen Plant 

Tyura-Tam Rail Station 

Krainy Airfield 

Syr dar'ya River 

Hwy To Tashkent 

BAIKONUR TOWN 

(LENINSK) 

Saturn Measurement Station 

approximately 90 km- figure not to scale 

Page 6-2



6.1.2 Building 92A-50 

Building 92A-50 contains all facilities necessary for processing a SC from its arrival through its integration with the 

Fourth Stage and encapsulation. The SC and GSE arrive at Hall 102, where the containers are cleaned. The SC 

container is transported on the railcar or on its own wheels into Hall 101, where the SC is normally removed from the 

container. The SC is then transported into Hall 103A on the SC transporter where it is installed onto its fueling/test 

stand. The SC remains in this hall for all subsequent testing and fueling operations. Following fueling, the SC is 

transported back to Hall 101 on a special transport dolly, where it is integrated with the Fourth Stage and 

encapsulated. 

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 also 

used for integrating the SC with the flight adapter and mating the SC to the Fourth Stage. Final encapsulation of 

the 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 as 

the 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, and 

personnel/office areas: 

a) Room 119 is a Class 100,000 cleanroom used for SC testing prior to loading. SC pressurization may occur in this 

room 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. Most 

electrical 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. 

Page 6-3



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 into 

three main halls. 

a) Room 1: Propellant Loading Area. A Class 100,000 clean tent is installed in this room to receive the SC for fuel 

and oxidizer loading operations. Oxidizer propellant cylinders are thermally conditioned in this room in special 

chambers prior to loading. 

b) Room 2: SC Removal from Container and Container Storage Operations. This room is used for agitation of 

propellant 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 to 

loading. 

In addition to the areas described above, a control room on the second floor (Room 58) houses SC Customer electrical 

ground support equipment that is used to monitor spacecraft telemetry, charge SC batteries, and communicate with 

the propellant-load team. 

Fuel and oxidizer decontamination rooms are available for cleaning the GSE. These are located at either end of the 

building 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 the 

Launch Vehicle. For the Breeze M configuration, the Launch Vehicle will return to Building 92A-50, Hall 111 for 

final 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 the 

Bunker, which is approximately 1 km from the launch pad, Room 250 is used for installation of SC customer electrical 

GSE 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 to 

house 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 required 

operations 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 of 

Kazakhstan in Central Asia, approximately 2,000 km southeast of Moscow. The annual temperature averages 13 o C, 

ranging from -40 o C in winter to 45 o C in summer. Figure 1.3-1 depicts the overall layout of the Cosmodrome, showing 

the facilities that are used for ILS launch campaigns. The specific SC processing areas at Cosmodrome described 

include: 

Page 6-4



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. Specialized 

equipment 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 of 

the 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 fueling 

and pneumatics pressurization are accomplished in either Area 92, (Building 92A-50) or Area 31, (Building 44) the 

Spacecraft Fueling Hall). The LV fourth stage with the mated spacecraft and fairing are transferred to the LV 

Technical Zone in Area 92, (Building 92-1), where they are horizontally mated to the assembled LV Stages 1-3. The 

integrated LV is then transported to the Proton Launch Zone (Area 81) for erection, checkout, and launch (see Section 

6.3: 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 the 

specific needs of SC Customers, and provides the capability to perform required operations in one conveniently located 

area. 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 and 

services available to SC Customers for pre-launch payload processing. 

Page 6-5



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 encapsulation 

in a single building. The halls/rooms, facility systems, and equipment are sized to accommodate SCs of up to 

approximately 4.5 m diameter, 10.0 m height, and loaded weights of up to 8,700 kilograms. Figure 6.2.1.1-1 depicts 

the 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 the 

Receiving Area. In Hall 101, the container is removed from the railcar and placed on the floor. After the railcar is 

removed and the environment reestablished, the SC is removed from the shipping container and placed on a 

transporter to be moved into the Processing and Fueling Hall (Hall 103A). Once there the SC is placed on the fueling 

island and require no further movement in order to complete all necessary standalone assembly, checkout, propellant 

loading, 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 commercial 

programs. 

The Receiving Area (Hall 102) is the primary entrance for the SC and associated equipment, and is located on the east 

side 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 the 

building, near the Change Room Area. An additional entrance on the north side of Hall 103A is used for Control 

Room equipment delivery. 

Page 6-6



Figure 6.2.1.1-1: Building 92A-50 General Arrangement 

Page 6-7



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 its 

transport railcar. External access for the SC and its GSE is provided by rail through two locally controlled, exterior 

sliding doors located in the hall’s east wall. An area is provided outside of Hall 102 for initial wash-down of the railcar 

and 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 and 

offloading. 

Container cleaning is accomplished in Hall 102. An area of approximately 240 square meters is provided for non-Class 

100,000 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, and 

the height of the overhead 50 T crane hook is 18.01 m. The SC unloading area is approximately 8.85 m wide and 34.1 

m long. 

When ready, the SC Container on the railcar is moved, via railcar, from Hall 102 to the inside of Hall 101 where it is 

transferred to an air pallet for the move into the Processing and Fueling Area (Hall 103A). Alternatively, the SC may 

be 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 is 

a Class 100,000 cleanroom. The Integration Hall is used to assemble the Ascent Unit, which involves the following 

operations: 

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 as 

transferring the integrated Ascent Unit from the rollover fixture to a railcar for delivery to the Integration Facility 

92-1. 

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, thermal 

insulation, and an anti-static floor coating. 

Page 6-8



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 SC 

processing, including loading propellants and servicing pneumatics in the SC (Figure 6.2.1.4-1). Access to Hall 103A 

is made available from Hall 103 through two sliding doors with a clear opening that is 9.5 m wide by 11.95 m high. A 

15 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. It 

is surrounded by a grating-covered trench, which drains any fuel or oxidizer spills into separate waste tanks. The 

grating 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/cm 2 ) per truck axle. All 

finishes 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 per 

square 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. The 

SCAPE Shower Areas (Rooms 4121 and 4122) have showers for post-operation clean-up. A parking lot for 

ambulances and fire trucks is located next to the rapid egress routes. The Pressurization Airlock (Room 4110) and the 

space between the double doors between Hall 103 and 103A are pressurized with clean air in order to isolate Hall 103A 

during propellant loading operations. 

The clear dimensions for the SC Processing and Fueling Hall 103A are 16.5 m wide by 22 m long. 

Page 6-9



Figure 6.2.1.4-1: Building 92A-50 Spacecraft Processing and Fueling Area 

. 

BUILDING 92A-50 

SPACECRAFT PROCESSING & 

FUELING AREA 

Page 6-10



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, N 2H 4) 

and thermal conditioning of the fuel before loading. Room 4105, the Oxidizer Conditioning Room, is used for 

temporary storage of the campaign oxidizer (e.g.N2O4) and thermal conditioning of the oxidizer before loading (Plate 

3.1-6). Both rooms have the capability to collect and dispose of propellant spills. Room 4112 contains no materials 

that 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/cm 2 ) per truck axle. The 

floor 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, the 

Oxidizer Equipment Decontamination Room, is used to decontaminate oxidizer loading equipment. Both rooms have 

the capability to collect and dispose of propellant spills. Room 4111 contains no materials that react with fuel, and 

Room 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 Hall 

103A, 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 monitoring 

all processing and fueling operations. The wall between 103A and the Control Room is a welded, reinforced steel 

structure 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 entry 

vestibule, with inner and outer doors 2.9 m wide and 2.9 m high, is provided to facilitate equipment movement into 

the Control Room. 

The floors of the Control Room and all associated access corridors are designed for wheeled dollies. Forklifts may be 

used to bring equipment into the vestibule of the room, but they are not permitted to operate in the Control Room 

itself. Temporary ramps are available to aide moving items from the entrance vestibule into the Control Room. 

Page 6-11



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’s 

restrooms, change and shower areas, a storage and issue room for cleanroom garments, a Personal Protective 

Equipment storage and donning room, and a corridor with an air shower. Clean passage is available from the Change 

Rooms 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. During 

propellant loading operations in Hall 103A, the Airlock is pressurized slightly more that Hall 103A to prevent vapor 

migration from Hall 103A. SCAPE suited personnel use the airlock to access the SCAPE Showers and Doffing 

Rooms, and the Pressurization Airlock and Corridor to the Airshower and Change Rooms can be used as an 

emergency 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 doffing 

Personal Protective Equipment (PPE) for a propellant loading operation. As necessary, SCAPE Showers, Rooms 4121 

for Fuel and 4122 for Oxidizer, are available to decontaminate the PPE suits before doffing. The dedicated showers 

are 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 m 

by 3.4 m and Room 4122 is 1.2 m by 3.4 m. 

Page 6-12



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 provides 

a 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 cubic 

meter, 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 ceiling 

height is 15.0 m. At heights greater than 3 m above the floor, the width of Hall 103 restricted by HVAC ducting to 

about 16 m. 

6.2.1.13 Ordnance Storage 

Khrunichev provides limited storage of ordnance required to support a launch campaign. The Ordnance Storage Room 

may 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 cm 

may 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 of 

Transportation-approved shipping and storage containers 

c) The SC Customer must provide a certificate of conformance to the Hazard of Electromagnetic Radiation to 

Ordnance (HERO) Specification (MIL-I-23659) 

Page 6-13



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 (Figure 

6.2.1.14). 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 via 

stairs from the “street” entrance to the Change Room Area. Only essential personnel are permitted in the 

Office/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 the 

Khrunichev area of the building. A 2,000 kg capacity freight elevator, with a 1.8 m wide by 2.2 m high door opening 

and 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 and 

outfitted with the specific needs of SC manufacturers and customers in mind and, along with Building 44, provide the 

capability 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 relative 

to the Propellant Filling Hall (Building 44). 

Page 6-14



Figure 6.2.2-1: Spacecraft Processing Zone (Area 31) 

Building 44 

Propellant 

Filling Hall 

Rm.1 

P A S - 5 - 

0 9 1 

SC fueling area 

300-500 m 

Area C 

H a l l 1 0 0 

Building 40 

Common Hal l 

Building 40D 

Preparation Hall 

A scent Uni t 

Int egrati onA r ea 

This section describes the specific areas of Buildings 40/40D as well as the equipment and services available to SC 

Customers 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 to 

enable customers, and potential customers, to make detailed plans for payload processing activities. It also serves as a 

useful 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 perform 

the 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 

Area A 

Area B 

c) Ascent Unit integration (i.e., mating the Fourth Stage, the LV adapter system, the SC, and the Fairing, followed 

by 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. 

Rm. 

119 

SC 

Area 

Rm. 

120 

Rm. 

121 

Page 6-15



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 rail 

in the manufacturer’s shipping container to Area C and transferred into environmentally controlled Class 100,000 

clean 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 is 

approximately 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 for 

vertical SC integration and the necessary mechanical ground support equipment (GSE) for integration operations with 

the Fourth Stage, the LV adapter system, the Fairing, and the assembled Ascent Unit. Figure 6.2.2.2-1 shows the 

hall’s general arrangement. 

Figure 6.2.2.2-1: Building 40, Hall 100 (Common Hall) Layout 

30 m 

Thermal 

Control Car 

Locomotive 

Rooms: 

100 = Common Hall 

119 = Payload Preparation Hall 

120 = STE Lowbay 

121 = Storage 

6.2.2.2.1 Area A 

Railway 

Transporter 

with SC 

Area C 

120 m 

Room 100 

Ascent Unit 

Integration 

Stand 

Area A 

Area B 

Cleanroom for 

Ascent Unit 

Two Traveling Cranes Assembly 

Hook Height = 14.5 m 

Load Carrying Capacity 10/50 Tons 

Traveling 

Crane Above 

Area A, which measures approximately 11 m by 47 m, is a Class 100,000 clean area used for final preparation of the 

Ascent Unit and its components. It also is used for integration of the SC with the SC flight adapter and then with the 

Fourth Stage. Final encapsulation of the spacecraft takes place in this zone. It is separated from Area B by an 8 m high 

partition that has an upper opening approximately 18 m long and 4 m high used for transferring cargo by crane 

between 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 material 

from 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 strictly 

environmentally controlled and is used for loading/unloading SC and support equipment, as well as the Fourth Stage 

and Ascent Unit components. 

Room 

119 

Room 

120 

Room 

121 

Page 6-16



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 rooms 

are environmentally controlled, Class 100,000 clean areas. Details of this layout are provided in Figure 6.2.2.3-1. The 

total usable area of Rooms 119, 120, and 121 is approximately 400 square meters. Change rooms and restrooms also 

are located on the first floor. 

Figure 6.2.2.3-1: Building 40D, First Floor Layout 

12 m 

PA S- 5 - 0 2 2 

5.5 m 

(H = 7 m) 

Room 119 

Room 101 

Room 

122 

5. 5 m 

2 m 

3 m 

13 m 

114 

Room 

111 

Room 

112 

Room 

104 

Room 

109 

Room 110 

Main Entrance 

Room 

108 

4 m 

(H = 5 m) 

Room 120 

H = 4.5 m 

Room 121 

12 m 

9 m 

16 m 

Rooms: 

101 - Personnel changing 

room 

104 - Air shower 

108 - Security Office 

109 - Medical Office 

112 - Security Checkpoint 

116 - Restroom 

120 - Control Room 

121 - Storage 

122 - ILS Storage 

The normal personnel entrance into Building 40D is through an outside door past a security checkpoint. This building 

is the only one completely controlled by U.S. security for the duration of a launch campaign. Security video monitors 

are located in the first floor Security Office (Room 108). Access to cleanroom areas (Room 119, 120, 121) is through 

an air shower (Room 104) into Room 121. 

Page 6-17



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 an 

environmentally controlled Class 100,000 cleanroom; details of its layout are provided in Figure 6.2.2.3-1. The room 

is approximately 15.8 m long, 13.8 m wide, and 10 m high. A 4.85 m wide door that is 7 m high provides access to 

Hall 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 100,000 

cleanroom that is approximately 12 m long, 9 m wide, and 10 m high (Figure 6.2.2.3-1). The room has a 3.4 m wide, 

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, and 

10 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 and 

shower rooms are faced with ceramic tiles throughout their entire height. The floors of the support rooms are cast-inplace 

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 near 

Building 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 the 

physical arrangements of these areas. 

Page 6-18



Figure 6.2.2.4-1: Building 40D, Second Floor Layout 

PA S- 5 -023 

Room 201 

Room 202 

Room 203 

Room Rm Room Rm 

207 207 206 

Figure 6.2.2.4-2: Building 40D, Third Floor Layout 

P A S- 5 - 0 2 4 

Room 301 

Room 

303 

Room 

31 3 

Room Rm 

Room 

312 31 

311 

1 

Room 

305 

Room 

205 

Room 3 04 

Room 203 

Room 

308 

Room 204 

Rooms: 

201 - ILS Office 

203 - 50 Hz Power 

Conditioning 

Equipment 

204 - Security Office 

206/7 Restroom 

208 - Khrunichev 

Storage 

Room 305 

Room 307 Room 306 

Rooms: 301/307 = SC Customer Offices 

304 = Automatic Telephone Exchange 

305 = SC Customer Breakroom 

306 = Conference Hall 

311/312 = Restroom 

Page 6-19



The second floor of Building 40D is used by the ILS team. Room 201 (48 m 2 ) is used as the ILS Office Area, and 

Room 204 (32 m 2 ) is used for ILS Security. Room 203 houses 50 Hz power racks and is reserved for Khrunichev 

engineers. 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 m 2 ) 

and 307 (15 m 2 ). Room 305 (42 m 2 ) is normally used as a breakroom. Room 306 (42 m 2 ) is used as a conference 

room. 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. This 

facility provides the capability to safely perform SC pre-filling operations, propellant thermal conditioning, and SC 

fueling in a clean environment. The fueling station has been upgraded to support the loading of the SC with 

appropriate propellants. (Propellants are supplied by the SC Customer and are loaded using the SC Customer’s 

equipment.) 

This section describes the specific areas of Building 44, as well as the equipment and services, available to SC 

Customers 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 three 

equal 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 is 

placed on the Customer's fueling stand in a clean tent and secured to a tray platform. Room 1 includes a 5.5 m wide by 

7 m high sliding door and two emergency exits. Two emergency exits are equipped with panic hardware (each lock 

normally 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 ground 

support equipment to monitor the SC. Fuel and oxidizer decontamination rooms are available for decontamination of 

support equipment following loading. Areas for propellant thermal conditioning also are available. 

Page 6-20



Figure 6.2.3.1-1: Building 44 (Filling Hall) Floor Plan 

Page 6-21



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 clean 

tent consists of two parts: the SC Zone, which measures approximately 9.2 m by 7.6 m; and the Ground Support 

Equipment (GSE) Zone, which measures approximately 4.6 m by 6.6 m. Two configuration options are available to 

the SC Customer for the clean tent: a pass-through corridor may be installed between the SC and GSE zones to 

facilitate personnel passage between the two areas; or a blast shield may be installed between the SC and GSE zones to 

prevent 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 compatible 

with 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 SC 

adapter. The aluminum floor has channels that are approximately 2 m apart and 3 to 8 cm deep that provide a pathway 

for liquid spills to drain to one end of the tent and to the spill tanks located underneath the floor of Building 44. The 

oxidizer 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 a 

Class 100,000 cleanroom. The air flows into the tent and leaves through outlets at the bottom of the walls. The tent is 

pressurized to between 9.8 and 28.6 Pa relative to the outside of the tent. HEPA filters are installed at the inlets of the 

stainless steel air ducts. Lighting in the clean tent is provided at an illumination of 100 dekalux. The top and front top 

wall 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 m 

long and has a window that is covered with explosion-proof glass to facilitate observation of fueling operations. An 

umbilical 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 SC 

propulsion GSE after propellant loading operations. Room 26, the fuel decontamination room, is approximately 5 m 

wide and 8 m long and is located adjacent to Filling Hall Room 1. Room 27, the oxidizer decontamination room, is 

approximately 9 m wide and 9 m long and is located adjacent to Filling Hall Room 3. Access to these rooms is 

provided through exterior entrance doors only. 

Each decontamination room is supplied with gaseous nitrogen, which may be used to purge SC ground support 

equipment. Floor drains in each room are connected to the facility’s appropriate contaminated water sump tanks. 

Page 6-22



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 launch 

vehicle 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 steel 

that measures approximately 50 m wide and 120 m long (Figure 6.2.5.1-1). The building includes the 

integration/assembly room and two laboratory annexes adjoining the assembly room. The building is provided with 

heating, ventilation, and fire-fighting systems, fire and security alarms, and special lighting. An overhead crane is 

available 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 the 

LV 

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 rail 

tracks with a gauge of 1,524 mm. The central track and one of the lateral tracks are throughways, while the third track 

terminates in the Processing Room. Along the building’s central axis, there are three double-leaf gates measuring 4.7 

by 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 27 o C 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 by 

using the Thermal Conditioning Railcar, if necessary. 

Page 6-23



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 the 

Proton Launch Zone, Area 81, for erection, checkout, and launch. This section describes the facilities that are used for 

erection, checkout, and launch of the Proton LV at the Baikonur Cosmodrome. In particular, areas available at the 

launch complex for Spacecraft (SC) equipment and personnel are described. These areas may be used by the SC 

Customer 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 erection 

and launch. Figure 6.3.1-1 shows a plan view of the launch complex. The launch area includes the following physical 

facilities, 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 6-24



Figure 6.3.1-1: Proton Launch Zone, Area 81 

PUG968 

Launch 

Pad 23 

600 m 

Bunker 

Rooms 250/251 

Launch 

Pad 24 

340 m 

Page 6-25



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 provide 

electric, pneumatic, and hydraulic links between the ground system testing equipment and on-board LV hardware via 

the LV transit cables and pipes. The launch structure is designed to withstand LV first-stage engine plume 

impingement. 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, and 

communications services. All launch campaign operations requiring the presence of personnel in the vault must be 

completed prior to the start of LV fueling, which occurs approximately 7 hours prior to launch. All personnel are 

required to leave the vault by this time, and from then on, all vault equipment must be controlled remotely from the 

bunker (Room 250). Additional details about Room 64 are provided in Drawing 4.1-1. Plate 4.1-1 provides a 

photographic 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 is 

equipped with 60 Hz electrical power, grounding, and communications services. All launch campaign operations 

requiring the presence of personnel in the vault must be completed prior to the start of LV fueling, which occurs 

approximately 7 hours prior to launch. All personnel are required to leave the vault by this time, and from then on, all 

vault 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-launch 

processing and launch. The Mobile Service Tower includes service platforms, a gallery, service fixtures, two cargopassenger 

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 6-26



Figure 6.3.2.2-1: Proton Mobile Service Tower 

30.515 m 

30.400 m 

Page 6-27



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 and 

provides protection for personnel and equipment during the launch. The bunker houses the LV and SC system test 

equipment and GSE required for pre-launch operations and monitoring of SC readiness. Air temperature and 

humidity 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 6-28



7. LAUNCH CAMPAIGN 

7.1 ORGANIZATIONAL RESPONSIBILITIES 

In any given launch campaign several organizations are involved. What follows is a brief overview of the responsibilities 

of these organizations. 

7.1.1 Khrunichev 

a) Overall responsibility for coordinating work performed at the launch complex by the Strategic Rocket Forces (and 

NPO Energia if Block DM used) 

b) Engineering support and quality inspection for all testing performed on Stages 1-3 of the Launch Vehicle, as well 

as the adapters and fairing. For launches with the Breeze M Khrunichev is also responsible for Breeze M 

engineering, 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 7-1



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 assets 

while 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 SC 

Customer and ILS must be dealt with by the SC Customer and ILS. Matters affecting the terms of the subcontract 

between ILS and Khrunichev must be dealt with by ILS and Khrunichev. In particular, any issues involving possible 

additional 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 in 

conformance 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 launch 

campaign. 

Page 7-2



Figure 7.2.1-1: Organization During Launch Campaign 

Customer MM 

SC Subcontractor 

ILS MM 

SC Ops Manager ILS Ops Manager 

7.2.2 Organization During Combined Operations 

Khrunichev 

Program Director 

Khrunichev 

Ops Manager 

Russian Strategic 

Rocket Forces 

Other Subcontractors 

Combined operations are those operations involving some combination of SC Customer organization, ILS, and 

Khrunichev personnel (e.g., Khrunichev adapter mating, Fourth Stage to payload mating, encapsulation, Ascent Unit 

checkout, Ascent Unit integration to Stages 1-3, any operations on the pad which require payload faring access). For 

each such combined operation, one operation leader is assigned, either from Khrunichev or the SC Customer. This 

individual directs the operation and ensures that it is carried out in conformance with the mutually agreed upon 

procedures. 

For each operation, one person from the SC Customer organization, ILS and Khrunichev is designated as team leader 

for their respective organizations. Agreements among organizations can only be reached among these three team 

leaders. 

Security personnel from either or both ILS and the Strategic Rocket Forces may be present during any operation if 

required by the Security Plan. 

ILS provides at least one interpreter for each combined operation. Special training is conducted with the SC Customer 

and Russian personnel for joint crane operations to ensure reliable communications between English and Russianspeaking 

personnel. 

Either Khrunichev or the SC Customer provides a Quality Assurance Representative for each operation who 

documents any test discrepancies on a Quality Assurance Report. 

Strategic Rocket Forces personnel implement many of the operations at the Baikonur Cosmodrome. The Strategic 

Rocket Forces is a Khrunichev subcontractor and, as such, coordinates directly with Khrunichev and NOT with the SC 

Customer or ILS personnel. 

Page 7-3



7.2.3 Planning Meetings 

ILS maintains the master schedule for combined operations planning and reviews it with the SC Customer and 

Khrunichev at a daily scheduling meeting. At this meeting, the current operations for the day are reviewed and the next 

3 day’s operations are agreed upon. Following each meeting, ILS revises the master schedule for the following day. At 

a 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. These 

critical 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 7-4



7.3 COUNTDOWN ORGANIZATION 

The countdown organization is illustrated in Figure 7.3-1. 

Figure 7.3-1: Countdown Organization 

SC Customer 

Mission 

Director 

ILS 

Mission 

Manager 

Energia 

Mission 

Director 

(if Block DM 

used) 

Khrunichev 

Mission 

Director 

KBOM* 

Readiness 

Review Board 

Authorization Strategic Rocket 

Forces Launch 

Commander 

Launch 

Authorization 

SC Ready 

L - 10 min 

Spacecraft 

Director 

(SC Customer) 

Automatic 

Computer-Controlled 

Sequencer 

LV Ready 

L - 5 min 

Stages 1-3 

Director 

(Khrunichev) 

LIFT-OFF 

COMMAND 

Fourth Stage 

Ready 

L - 2 min 

Fourth Stage 

Director 

(Energia) if 

Block DM 

used) 

*Design Bureau for General Machine Engineering (Launch Site) Note: Launch Commander has sole authority to abort launch 

Page 7-5



The Strategic Rocket Forces directs the countdown, which follows a pre-approved script. The Launch Commander 

receives authorization to launch from the Readiness Review Board, which consists primarily of the five entities shown 

in Figure 7.3-1. 

Certain organizations have pre-assigned abort capability. Each of these organizations is asked to acknowledge the 

readiness of their subsystems on Launch Day according to the Launch Day Script. These subsystem readiness checks 

are 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 each 

representative is vested with abort authority over the launch sequencer for their respective area of responsibility. For 

example, 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-board 

power to the first three stages and begins the sequence for first stage engine ignition. Normally, readiness signals are 

given 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 Start 

Command to be issued automatically by the Start Timer Mechanism. This command is issued at 2.5 seconds prior to 

liftoff, at which time the final relay closes, sending a command transferring to on-board power and initiating the First 

Stage 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 opened 

following issuance of the Launch Vehicle Start Command, the launch can no longer be aborted. The SC Customer can 

abort 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 capability 

up to 2.5 seconds prior to lift-off contact. 

Page 7-6



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 provides 

details 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 and 

Customer requirements. 

Page 7-7



Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50 

Page 7-8



Figure 7.5-1: Typical SC Campaign Operations Assuming Use of Building 92A-50 (Continued) 

Page 7-9



Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31 

Page 7-10



Figure 7.5-2: Typical SC Launch Operations Assuming Use of Area 31 (Continued) 

Page 7-11



7.5.1 Launch Vehicle Processing 

The Proton LV stages and fairings are built in Moscow by Khrunichev and transported by rail to the Baikonur 

Cosmodrome typically well in advance of the SC delivery date. After transportation of the Proton’s stages and fairing 

by rail, LV build-up takes place in an integration and test facility (Building 92-1) capable of supporting four 

simultaneous Proton assembly and checkout operations. The Fairing is moved either to Building 40 or to Building 

92A-50, depending upon which facility is to be used for SC integration, prior to SC arrival, there it is stored and 

cleaned in preparation for encapsulation. The Fourth Stage is delivered to Area 254 for pre-launch checkout and 

testing. The Fourth Stage is then delivered to the Building 44 in Area 31, the propellant fueling hall, where either 

kerosene is loaded (Block DM) or MMH and N 2O 4 (Breeze M). It is then moved to either Building 40 or 92A-50 for 

integration with the SC. Payload adapters typically are delivered shortly before the processing cycle and cleaned and 

prepared 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 to 

commodities 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. They 

are 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 containers 

are 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 mobile 

cranes, K-loader, and 5 and 15-ton forklifts as required. After the SC container is placed on a railcar, it may be 

connected to a thermal control railcar via two air duct flanges (inlet and outlet air flow) to provide thermal 

conditioning during transport. Some spacecraft containers are completely self-contained thermally and 

environmentally and do not require this support option. The thermal car also provides a dynamic load monitoring 

system, so use of the car may be limited to use of this system. SC Customer personnel effects may be transported 

directly 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 is 

approximately 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. The 

Building 40 bridge crane may be used for offloading large GSE such as the SC container. Spacecraft GSE and 

container handling proceed as required for each unique spacecraft operations flow. 

Page 7-12



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 brief 

exposures 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 SC 

maintains its vertical orientation, as the Thermal Transport Container is designed to interface to the SC adapter 

mating surface. This container is used to transfer the SC to the Filling Hall for propellant fueling. Once on the 

container 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. Use 

of this temperature control system may be waived by the SC Customer since the nominal time for transportation to the 

Filling Hall is of short duration (less than 1 hour). The Thermal Transport Container has been specifically designed to 

provide 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 is 

removed in Room 2 and placed on support stands. The railcar is then rolled into Room 1 and the SC is hoisted from 

the Thermal Transport Container base onto the SC loading stand already installed in the clean tent. The clean tent 

ceiling 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 vent 

and 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 cold 

GN2 purge is available as required. 

Following fueling, the process is reversed and the SC is returned to the Thermal Transport Container and then moved 

back to Building 40, Room 100 for further testing and processing. Electrical checkouts are made on the adapter 

electrical 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 integration 

stand commonly referred to as the “Gallows”. The Gallows allows taller payloads to be installed onto the Fourth 

Stage without structurally modifying Building 40D. The flight clamp band is installed and tightened to the correct 

flight-specific tension. 

Following integration, an end-to-end electrical checkout is performed. The composite Ascent Unit is rotated to the 

horizontal for encapsulation with the two fairing halves. Following encapsulation, an RF GO/NO GO test is 

performed through the RF window. The Ascent Unit is then transferred to its railcar for transport to Area 92. During 

transport, the Ascent Unit is thermally conditioned using the thermal conditioning railcar; input air to the Ascent Unit 

is provided at the base of the Fourth Stage and output occurs at the Fairing nose. 

Page 7-13



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 in 

this area prior to moving the SC to the Integration Hall (Hall 101). Hall 101 is a Class 100,000 cleanroom, and the 

SC container cover may be removed here or in Hall 103A as required by the unique mission-specific SC processing 

flow. 

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 by 

means of an external door which opens directly into the control room loading area. This is a small buffer zone between 

two sets of double doors with a concrete floor. 

After container removal, SC electrical testing, pneumatic testing, and propellant fueling occurs in Hall 103A. Passthroughs 

from the control room are available for cabling. These cable feeds are verified to be leak-tight prior to 

propellant operations. The 60 Hz, 120 Vac power source is provided by an UPS. Typically the UPS is activated the 

week prior to SC arrival and not deactivated until the SC leaves the facility and all parties agree that no further 

requirement 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 vapor 

detection system which alarms locally, in the control room, and at the Security Command Post. Breathing air is 

supplied by a single source which is sampled prior to operations. GN2, water, and shop air are provided on demand. A 

fire suppression system which will arm but not release on alarm is also active in Hall 103A. The command to activate 

the suppression system deluge is made in the control room, and is not an automatic function of the alarm system. LN2 

is available with 24 hr call-up. 

Page 7-14



The loaded SC is transported back to Hall 101 using the Transport Dolly. The SC is lifted from the transporter and 

mated to the SC adapter using the 50T bridge crane. The SC/adapter unit is then lifted and mated to the Fourth Stage 

which 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 with 

the SC mated to the Fourth Stage. 

After SC integration, final closeout operations and photographs are performed. The combined Fourth Stage/SC is 

rotated to the horizontal position on the integration stand and is encapsulated by the bi-conic payload Fairing. After 

the upper fairing half is emplaced, an RF Go/NoGo test is performed to ensure that the SC link has not been disturbed 

and that the RF window is transparent to RF. This is performed as soon as the fairing half is mechanically emplaced 

and before continuation of encapsulation sealing operations. If any anomaly is found, the fairing may be removed 

relatively 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 Building 

92-1 (Integration Hall) for integration with the Proton Launch Vehicle. At this point, the processing flow for all SC 

becomes a common flow, independent of SC Processing Facility. 

7.6 LAUNCH VEHICLE INTEGRATION (BUILDING 92-1) THROUGH LAUNCH PAD 

OPERATIONS 

The Ascent Unit is transported to Building 92-1, where it is disconnected from the thermal conditioning car and lifted 

onto the integration dolly. Here, it is brought together horizontally with the mating surface of Stage Three of the 

integrated Proton launch vehicle (LV). An end-to-end electrical check is performed on the SC/LV cables. The 

integrated LV is then transferred to the transportation/erection fixture and a thermal blanket is installed over the 

Fairing to protect the payload from temperature extremes during periods when there is no active thermal control. A 

typical launch flow requires 4 days in the Integration Hall. Activities in the Integration Hall are controlled by the 

Strategic 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 the 

Breeze M, the LV will be transported first to 92A-50 for final integrated Breeze M tests. This will be followed by 

transportation to the Breeze M Filling Station for top off of the low pressure MMH and N2O4 reservoirs on its way to 

Launch Pad 24. 

The first of two Inter-governmental Commission Meetings is held on Day L-6, prior to vehicle roll-out to the pad, to 

ensure all agencies are go for pad roll-out. All LV agencies, including the SC Manufacturer will be called upon to 

provide a launch readiness statement. 

The LV along with the thermal control railcar is moved to the launch pad. The vehicle is erected, and the liquid 

thermal conditioning system is activated. An RF check of the SC telemetry and command links is performed from the 

Bunker (Room 250) prior to Mobile Service Tower rollup. Once the Mobile Service Tower is in place, the Ascent Unit 

air-conditioning system is activated and the liquid system is turned off. 

Page 7-15



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 Vehicle 

Transporter. 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 by 

Khrunichev with SC/Customer input and should include all pad access requirements and requirements for RF 

radiation and commanding the SC. Operator “7” is the Khrunichev Program Director while Operator “701” is the SC 

Point of Contact. Information required from the SC/Customer: RF radiation, battery charging, SC commanding, pad 

access. Note that ILS functions such as pad access will also be coordinated through this script. Figure 7.7.1 provides a 

typical detailed on-pad operations flow. 

The SC Customer is expected to participate in a Launch Countdown Rehearsal on Day 3 on pad. This countdown 

rehearsal is supported by a full booster vehicle launch crew countdown and requires the SC Customer to indicate SC 

readiness to go at the required time. The rehearsal also includes a planned abort on a SC No-Go condition. SC full 

fidelity countdown rehearsal is not required for this exercise, simply the operation of the readiness switch at the 

planned 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 launch 

prior 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 tower 

removal. At T-2 hours, all personnel are cleared from the hotel areas and should be in their final positions for launch 

i.e. Bunker, Viewing Area, and Comm Center. Note that personnel in the Bunker and the Comm center should be 

limited to essential personnel only. 

The SC Customer participates in the final countdown by sending a SC readiness to launch signal at T-10 minutes, as 

noted in the Countdown Organization discussion. 

Active commanding of the SC is prohibited during critical booster vehicle functions. Propellant fueling of the booster 

vehicle, which starts at T- 8 hours, is one such time-frame. Other typical RF silence and no-command times are 

shown in Figure 7.7-2. 

Page 7-16



Figure 7.7-1: Typical Launch Pad Operations Timeline 

Page 7-17



8. PROTON LAUNCH SYSTEM ENHANCEMENTS 

The Proton launch system is supported by the Russian space industry leaders Khrunichev State Research and 

Production Space Center and Russian Space Complex Energia. Based on the long, successful history of the Proton 

launch system, ILS, Khrunichev, and Energia have continued to assess and conceptually study hardware and launch 

services enhancements that would improve the competitiveness and capability of the Proton launch system. 

Page 8-1



Figure 8-1: Proton Evolution Options 

Years of operat io n 

LEO capability 

GSO capability 

2000- 

22 .0 M T 

4.5 M T 

Proton-KM 

Page 8-2



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 launch 

system. 

8.1 HIGH VOLUME PAYLOAD FAIRINGS 

With the quantity of LEO, GTO, and planetary spacecraft on the Proton launch record, a number of payload fairing 

configurations have been developed in the past to uniquely meet the needs of the diverse Russian government 

programs. Most notable in recent developments has been the launch of large low earth orbit payloads such as the MIR 

space-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 booster 

vehicle. Unlike the Block DM configuration vehicle, the design of the Breeze M enables the stage to be completely 

encapsulated inside the fairing. For the Breeze M configuration, the payload fairing is attached directly to the third 

stage of Proton enabling large volume/higher mass fairings to be successfully carried with the limits of the Proton 

design. A number of payload fairing options have been assessed for development with the commercial Proton 

M/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. Usable 

cylindrical volume lengths of up to 15,000-mm have been assessed and determined feasible with the existing control 

authority capability of the Proton launch system. 

ILS and KhSC are willing to develop these concepts with the award of firm launch services contracts. These fairing 

options can be fielded in 30-36 months of authority to proceed and will support Proton M and Proton M/Breeze M 

vehicle launches as early as mid-2001. 

Page 8-3



Figure 8.1-1: Proton Larger Payload Fairing Concepts 

TBS 

Page 8-4



8.2 TANDEM LAUNCH SYSTEM 

In addition to large volume payload fairing options, ILS and KhSC have studied and conceptually designed a tandem 

launch system payload carrier concept. The tandem launch system concept may offer attractive opportunities for 

launch of multiple payload constellations to low and intermediate altitude orbits. Figure 8.2-1 illustrates possible 

interfaces available with the tandem launch system concept. 

Page 8-5



Figure 8.2-1: Breeze-M Launch Configuration With Tandem Launch Systems (TLS) 

5 

0 

30 

6 

4 

39 

45 

Φ 4350 

Φ 2490 

Φ 4100 

30 

20 

21 

1 

4 

1 6 

1 

6 

2 

6 

19 

Page 8-6



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 third 

stage of Proton and use a modified upper stage avionics system for control of the entire launch system from liftoff to 

separation of the payload in the targeted orbit. Removal of the avionics system increases typical performance to 

geosynchronous 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 significantly 

enhance launch system performance to high energy transfer orbits. Payload performance gains to GSO and GTO of 

almost 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 is 

now 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 capability 

to meet near term commercial launch services needs. 

Page 8-7



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 Soviet 

academician, V. N. Chelomey. The first launch took place in July 1965. The two-stage 8K82 (D) version, last flown 

in 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 the 

four-stage Proton K/Block DM (D-l-e, SL-12) model, both of which are currently in use. An improved version of 

the Proton (Proton M) is now in development, incorporating changes to the first three stages, as well as the 

development of a new storable propellantBreeze M upper stage. Figure A.1-1 shows the major variants of the Proton 

launch vehicle family. 

Figure A.1-1: Proton Launch Vehicle Family 

Years of operation 

LEO capability 

GTO capability 

GSO capability 

8K82 

(Proton D) 

1965-1967 

12.2 MT 

- 

- 

Proton K 

(Proton D-1) 

1967 

19.76 MT 

- 

- 

Proton K/Block DM 

(Proton D-1-e) 

1967- Present 

- 

4.9 MT 

1.9 MT 

Proton-M 

Breeze-M 

1999- Present 

21.0 MT 

5.5 MT 

2.92 MT 

Page A-1



The Block DM fourth stage of the Proton was developed independently during the 1960s as the fifth stage of the 

Russian manned lunar launch vehicle, the N2-L3. It was originally known in Russia as the Block D ("block" is the 

common translation of the Russian word denoting a rocket "stage", while "D" is the fifth letter in the Russian 

alphabet). 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 the 

Department of Defense. 

Proton has flown more than 235 missions, and has carried the Salyut series space stations and the Mir space station 

modules. It has launched the Ekran, Raduga, and Gorizont series of geostationary communications satellites (which 

provided telephone, telegraph, and television service within Russia and between member states of the Intersputnik 

Organization), 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 geostationary 

and interplanetary missions are launched on Proton. Approximately 90% of all Proton launches have been the fourstage 

version. 

The Proton launch vehicle has a long history of outstanding reliability. From its first operational launch in 1970 to the 

present 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 Table 

A.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 percent 

reliability record over the last 50 launches. 

Summary launch data by year are shown in Table A.2-1. 

Page A-2



Table A.2-1: Proton Launch Record Summary (1970-1998) 

Year Number of 

Launches 

Number of Launches by Version Total Launches on 

Accrual Basis 

4-Stage Version 3-Stage Version 

Failures 

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/Block 

DM’s 

1979 6 5 1 59 

1980 5 5 - 64 

1981 7 6 1 71 

1982 10 9 1 81 2 Proton K/Block 

DM’s 

1983 12 11 1 93 

1984 13 13 _ 106 

1985 10 9 1 116 

1986 8 7 1 124 1 Proton K 

1987 13 11 2 137 2 Proton K/Block 

DM’s 

1988 12 13 _ 150 2 Proton K/Block 

DM’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 

1996 8 7 1 224 2 Proton K/Block 

DM’s 

1997 8 8 - 232 1 Proton K/Block DM 

1998 6 5 1 238 

Page A-3



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 in 

December 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 w Cosmos Failed to orbit Command abort 

8 Aug 1970 w Unknown Failed to orbit No data 

12 Sep 1970 w Luna-16 Escape 

20 Oct 1970 w Zond-8 Escape 

10 Nov 1970 w Luna-17 Escape 

24 Nov 1970 w Cosmos-379 192 km x 1210 km at 51.9 deg 

2 Dec 1970 w Cosmos-382 2464 km x 5189 km at 51.9 deg 

26 Feb 1971 w Cosmos-398 201 km x 7250 km at 51.6 deg 

19 Apr 1971 w Salyut-1 200 km x 210 km at 51.6 deg 

10 May 1971 w Cosmos-419 145 km x 159 km at 51.5 deg 

19 May 1971 w Mars-2 Escape 

28 May 1971 w Mars-3 Escape 



14 Feb 1972 w Luna-20 Escape 

29 Jul 1972 w Salyut Failed to orbit 

8 Jan 1973 w Luna-21 Escape 

3 Apr 1973 w Salyut-2 207 km x 248 km at 51.6 deg No data 


21 Jul 1973 w Mars-4 Escape 

25 Jul 1973 w Mars-5 Escape 

5 Aug 1973 w Mars-6 Escape 

9 Aug 1973 w Mars-7 Escape 

26 Mar 1974 w Cosmos-637 LEO 

29 May 1974 w Luna-22 Escape 

24 Jun 1974 w Salyut-3 LEO 

29 Jul 1974 w Molniya-1S Elliptical orbit 

28 Oct 1974 w Luna-23 Escape 

26 Dec 1974 w Salyut-4 LEO 

6 Jun 1975 w Venera-9 Earth escape 

14 Jun 1975 w Venera-10 Earth escape 

8 Oct 1975 w Cosmos-775 LEO 

16 Oct 1975 w Luna Escape 

22 Dec 1975 w Raduga-1 GSO 

22 Jun 1976 w Salyut-5 LEO 

9 Aug 1976 w Luna-24 Escape 

11 Sep 1976 w Raduga-2 GSO 

26 Oct 1976 w Ekran-1 GSO 

15 Dec 1976 w Cosmos-881 and 882 LEO 

Page A-4



Table A.3-1a: Proton Launch History (Continued) 



17 Jul 1977 w Cosmos-929 301 km x 308 km at 51.5 deg 

23 Jul 1977 w Raduga-3 GSO 

4 Aug 1977 F Cosmos Failed to orbit 

20 Sep 1977 w Ekran-2 GSO 

29 Sep 1977 w 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 w Venera-l l Escape 

14 Sep 1978 w Venera-12 Escape 

17 Oct 1978 F Ekran Failed to orbit Second stage failure 

19 Dec 1978 w Gorizont-1 20,600 km x 50,960 km at 14.3 deg Failure? 

21 Feb 1979 w Ekran-3 GSO 

25 Apr 1979 w Raduga-5 GSO 

22 May 1979 w Cosmos- 1100 and 1101 193 km x 223 km at 51.6 deg 

5 Jul 1979 w Gorizont-2 GSO 


28 Dec 1979 w Gorizont-3 GSO 

2 Feb 1980 w Raduga-6 GSO 

14 Jun 1980 w Gorizont-4 GSO 

15 Jul 1980 w Ekran-5 GSO 

5 Oct 1980 w Raduga-7 GSO 

26 Dec 1980 w Ekran-6 GSO 

18 Mar 1981 w Raduga-8 GSO 

25 Apr 1981 w Cosmos-1267 240 km x 278 km at 51.5 deg 

26 Jun 1981 w Ekran-7 GSO 



30 Oct 1981 w Venera-13 Escape 

4 Nov 1981 w Venera-14 Escape 

5 Feb 1982 w Ekran-8 GSO 

15 Mar 1982 w Gorizont-5 GSO 

19 Apr 1982 w Salyut-7 473 km x 474 km at 51.6 deg 

17 May 1982 w Cosmos-1366 GSO 

23 Jul 1982 F Ekran Failed to orbit First stage failure 


12 Oct 1982 w Cosmos-1413 and 1415 19,000 km x 19,000 km at 64.7 deg 

20 Oct 1982 w Gorizont-6 GSO 

26 Nov 1982 w Raduga-11 GSO 

24 Dec 1982 F Raduga Failed to orbit Second stage failure 

Page A-5



Table A.3-1b: Proton Launch History (Continued) 



2 Mar 1983 w Cosmos-1443 324 km x 327 km at 51.6 deg 

12 Mar 1983 w Ekran-10 GSO 

23 Mar 1983 w Astron-1 1,950 km x 201,100 km at 51.09 deg 


2 Jun 1983 w Venera-15 Escape 

6 Jun 1983 w Venera-16 Escape 


10 Aug 1983 w Cosmos-1490 and 1492 19,000 km x 19,000 km at 64.8 deg 

25 Aug 1983 w Raduga-13 GSO 

29 Sep 1983 w Ekran-II GSO 

30 Nov 1983 w Gorizont-8 GSO 

29 Dec 1983 w Cosmos-1519 and 1521 19,000 km x 19,000 km at 64.8 deg 


2 Mar 1984 w Cosmos-1540 GSO 


29 Mar 1984 w Cosmos-1546 GSO 

22 Apr 1984 w Gorizont-9 GSO 

19 May 1984 w Cosmos-1554 and 1556 19,000 km x 19,000 km at 64.8 deg 

22 Jun 1984 w Raduga-15 GSO 

1 Aug 1984 Gorizont-10 GSO 

24 Aug 1984 w Ekran-13 GSO 

4 Sep 1984 w Cosmos-1593 and 1595 19,000 km x 19,000 km at 64.8 deg 

28 Sep 1984 w Cosmos-1603 836 km x 864 km at 71 deg 

15 Dec 1984 w Vega-1 Escape 

21 Dec 1984 w Vega-2 Escape 

18 Jan 1985 w Gorizont-ll GSO 

21 Feb 1985 w Cosmos-1629 GSO 


17 May 1985 w Cosmos-1650 and 1652 19,000 km x 19,000 km at 64.8 deg 


8 Aug 1985 w Raduga-16 GSO 

27 Sep 1985 w Cosmos-1686 291 km x 312 km at 51.6 deg 

25 Oct 1985 w Cosmos-1700 GSO 

15 Nov 1985 w Raduga-17 GSO 

24 Dec 1985 w Cosmos-1710 and 1712 19,000 km x 19,000 km at 64.8 deg 

17 Jan 1986 w Raduga- 18 GSO 

19 Feb 1986 w Mir 335 km x 358 km at 51.6 deg 

4 Apr 1986 w Cosmos-1738 GSO 

24 May 1986 w Ekran-15 GSO 

10 Jun 1986 Gorizont-12 GSO 

Page A-6



Table A.3-1c: Proton Launch History (Continued) 




25 Oct 1986 w Raduga-I9 GSO 


30 Jan 1987 F Cosmos-1817 188 km x 210 km at 51.6 deg Fourth stage failed to 

ignite 


31 Mar 1987 w 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 early 

shutdown 

11 May 1987 w Gorizont-14 GSO 

25 Jul 1987 Cosmos-1870 237 km x 249 km at 71.9 deg 





26 Nov 1987 w Cosmos-1897 GSO 


27 Dec 1987 w 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 w Gorizont-15 GSO 


6 May 1988 w Ekran-18 GSO 

21 May 1988 w Cosmos 1946-1948 19,000 km x19,000 km at 64.9 deg 

7 Jul 1988 w Phobos-1 Escape 

12 Jul 1988 w Phobos-2 Escape 

1 Aug 1988 w Cosmos-1961 GSO 

18 Aug 1988 w Gorizont-16 GSO 

16 Sep 1988 w Cosmos-1970P1972 19,000 km x 19,000 km at 64.8 deg 


10 Dec 1988 w Ekran-I9 GSO 

10 Jan 1989 w Cosmos-1987P1989 19,000 km x 19,000 km at 64.9 deg 

26 Jan 1989 w Gorizont-17 GSO 


31 May 1989 w Cosmos-2022P2024 19,000 km x 19,000 km at 64.8 deg 

21 Jun 1989 w Raduga-l-1 GSO 


28 Sep 1989 w Gorizont- 19 GSO 

26 Nov 1989 w Kvant-2 215 km x 321 km at 51.6 deg 

1 Dec 1989 w Granat 1957 km x 201,700 km at 52.1 deg 

15 Dec 1989 Raduga-24 GSO 

27 Dec 1989 w Cosmos-2054 Unknown 

Page A-7



Table A.3-1d: Proton Launch History (Continued) 




19 May 1990 w Cosmos-2079P81 19,000 km x19,000 km at 65 deg 

31 May 1990 w Kristall 383 km x 481 km at 51.6 deg 

20Jun 1990 w Gorizont-20 GSO 

18 Jul 1990 w Cosmos-2085 GSO 

9 Aug 1990 F Unknown Did not achieve orbit 



8 Dec 1990 w Cosmos-2109P11 19,000 km x 19,000 km at 64.8 deg 



14 Feb 1991 w Cosmos-2133 GSO 


31 Mar 1991 w Almaz-1 268 km x 281 km at 72.7 deg 

4 Apr 1991 w Cosmos-2139P41 19,000 km x 19,000 km at 64.9 deg 


13 Sep 1991 w Cosmos-2155 GSO 

23 Oct 1991 w Gorizont-24 GSO 

22 Nov 1991 w Cosmos-2172 GSO 


29 Jan 1992 w Cosmos-2177P79 19,000 km x 19,000 km at 64.8 deg 

2 Apr 1992 w Gorizon-25 GSO 

14Ju11992 w Gorizont-26 GSO 

30 Jul 1992 w Cosmos-2204-06 19,000 km x 19,000 km at 64.8 deg 

10 Sep 1992 w Cosmos-2209 GSO 


27Nov 1992 w Gorizont-27 GSO 

17 Dec 1992 w Cosmos-2224 GSO 

17 Feb 1993 w Cosmos-223?P3? 19,000 km x 19,000 km at 64.8 deg 


27 May 1993 F Gorizont Did not achieve orbit 2 nd and 3 rd stage 

propulsion failure 

30 Sep 1993 w Gorizont GSO 

28 Oct 1993 w Gorizont GSO 

18 Nov 1993 w Gorizont GSO 

23 Dec 1993 w Gorizont GSO 

20 Jan 1994 w GALS GSO 

5 Feb 1994 w Raduga - 30 GSO 

18 Feb 1994 w Raduga- 31 GSO 

11 Apr 1994 w Glonass l9,000 km x l9,000 km at64.8° 

20 May 1994 w Gorizant GSO 

Page A-8



Table A.3-1e: Proton Launch History (Continued) 



7 Jul 1994 w Cosmos GSO 

11 Aug 1994 w Glonass 19,000 km x 19,000 km at 64.8° 

21 Sep 1994 w Cosmos - 2291 GSO 

13 Oct 1994 w Express GSO 

31 Oct 1994 w Electro GSO 

20 Nov 1994 w Glonass 19,000 km x 19,999 km at 64.8° 

16 Dec 1994 w Luch GSO 

28 Dec 1994 w F Raduga - 32 GSO 

7 Mar 1995 w Glonass 19,000 km x 19,000 km at 64.8° 

20 May 1995 w Spektr 335 km x 358 km at 51.6° 

24 Jul 1995 w Glonass 19,000 km x 19,000 km at 64.8° 

31 Aug 1995 w Gazer GSO 

11 Oct 1995 w Looch - 1 GSO 

17 Nov 1995 w GALS GSO 

14 Dec l995 w F Glonass 19,140 km x l9,l00 km at 64.8° 

25 Jan 1996 w Gorizant GSO 

19 Feb 1996 F Raduga GSO Block DM propulsion 

failure 

9 Apr 1996 w Astra 1F Hi-GTO Commercial 

23 Apr 1996 w Priroda 214 km x 328 km at 51.6 deg 

25 May 1996 w Gorizant GSO 

6 Sep 1996 w Inmarsat 3 F2 GSO Commercial 

26 Sep 1996 w Express GSO 

16 Nov 1996 F Mars 96 Did not achieve escape trajectory Failure of Mars 96 control 

system to initiate Block 

D2 engine ignition 

24 May 1997 w Telstar-5 Hi-GTO Commercial 

6 June 1997 w Arak GSO 

18 June 1997 w Iridium LEO Commercial 

14 Aug 1997 w Cosmos-2345 GSO 

28 Aug 1997 w PanAmSat-5 Hi-GTO Commercial 

15 Sep 1997 w Iridium LEO Commercial 

3 Dec 1997 w Astra-1G Hi-GTO Commercial 

25 Dec 1997 F AsiaSat-3 GTO Block DM Engine Failure 

7 Apr 1998 w Iridium LEO Commercial 


8 May 1998 w Echostar-IV Hi-GTO Commercial 

30 Aug 1998 w Astra 2A Hi-GTO Commercial 

04 Nov 1998 w PanAmSat-8 Hi-GTO Commercial 

20 Nov 1998 w Zarya (FGB) LEO RSA/NASA 

Page A-9



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, 1 st stage engine cutoff due to false alarm from the launch vehicle safety system 

activated by the engine pressure gage. Manufacturing defect. Additional check of gages introduced at point of 

installation. 

b) 1972: After 181.9 seconds of flight, 2 nd stage automated stabilization system failure due to a relay short circuit in 

the "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 4 th stage oxidizer booster pump. Manufacturing/design defect. Cryogen-helium condensate 

freezing. Booster pump blowing introduced. 

d) 1977: After 40.13 seconds of flight, spontaneous cutoff of 1 st stage steering motor, loss of stability and engine 

cutoff at 53.68 seconds into the flight safety system command. Steering most failure due to spool-an-sleeve pair 

manufacturing defect (faulty liner) which caused penetration of hard particles under liner rim and resulted in 

spool-and-sleeve seizure. 

e) 1978: After 87 seconds of flight, loss of stability commenced due to error of 1st stage second combustion chamber 

steering gear. High temperature impact on cables due to heptyl leak into second block engine compartment. Leak 

likely 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 system 

electric circuit failure in rear compartment of 2nd stage caused by hot gases leaking from second engine gas inlet 

due to faulty sealing of pressure gage. Gage attaching point upgraded. 

g) 1978: After 235.62 seconds of flight, 2 nd stage engine shutoff and loss of stability caused by a turbine part igniting 

in turbo pump gas tract followed by gas inlet destruction and hot air ejection into 2 nd rear section. Engine design 

upgraded. 

h) 1982: At 45.15 seconds into the flight, major malfunctioning of 1 st stage engine fifth chamber. Flight terminated 

by launch vehicle safety system command. Failure caused by steering motor malfunctioning: first stage of 

hydraulic booster got out of balance coupled with booster dynamic excitation at resonance frequencies. Hydraulic 

booster design redefined. 

i) 1982: 2 nd 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 included 

introduction of self-latching action capability for program power distributor shaft. 

k) 1987: 4 th stage control system failure due to component (relay) defect. Manufacturing defect. Remedial program 

introduced at supplier's factory. Inspection made more stringent. 

l) 1987: 4 th stage control system failure due to control system instrument defect. Manufacturing defect. Device 

manufactured 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-10



m) 1988: 3 rd stage engine failure caused by destruction of fuel line leading to mixer. Unique manufacturing defect. 

Inventory rechecked. 

n) 1988: 4 th stage engine failure due to temperature rise in combustion chamber caused by penetration of foreign 

particles from the fuel tank. Manufacturing defect. Remedial program introduced at point of manufacture to 

prevent penetration of foreign particles into tanks. No recurring failures recorded. 

o) 1990: 2 nd 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 launch 

vehicle. 

p) 1993: 2 nd and 3 rd stage engine failures. Multiple engine combustion chamber burn-through caused by propellant 

contaminants. Remedial program introduced to modify propellant specifications and testing procedures. All 

launch 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 prevent 

two possible causes. The first involved introduction of redundant lockers, revised installation procedures, and 

increased factory inspections to prevent a loosening of a tube joint causing a leak that would prevent engine 

ignition. The second involved additional contamination control procedures to further precule particulate 

contamination of the hypergolic start system. 

r) 1996: Block D2 4 th stage engine failure during second burn due to malfunction of Mars 96 spacecraft control 

system, and associated improper engine command sequences. Unique configuration of spacecraft and 4 th stage. 

Remedial program includes stringent adherence to established integration and test procedures. 

s) 1997: Block DM 4 th stage engine failure resulting from improperly coated turbopump seal. Remedial program 

includes removal of unnecessary (for < 4 burn missions) coating. 

Page A-11



B1. QUALITY SYSTEM 

TBS 

Page B1-1



Introduction 

This section describes the data required from the spacecraft customer to determine the compatibility of the spacecraft 

with the Proton launch vehicle. Providing this data in full constitutes providing an Interface Requirements Document 

which is a contractual document provided by the Customer at the beginning of a mission integration cycle. For 

preliminary 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 to 

simplify 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 C1-1



C2. INPUT DOCUMENTS 


Spacecraft Dynamic 

Model 

Spacecraft Thermal 

Model 

Spacecraft Physical 

Model 

Per specification LKES-9704-0207 

Per specification LKET 9704-0206 

Per specification LKEB 9801-039 

Launch Operations Plan Reference Section 5.3.5 of this Mission Planner’s Guide 

Environmental Test 

Plan 

SC Interface Control 

Drawing 

Reference Section 5.3.3 of this Mission Planner’s Guide 

Drawing based on physical model above which shows: 

Locations of all critical surfaces (surfaces closer than 50 mm from 

useable volume envelope). Show on drawing and in table in SC 

coordinates 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 to 

125 mm above interface plane including stiffness characteristics 

Umbilical connector definition including position relative to 

separation plane 

Pneumatic connector interfaces (if applicable) 

Access envelopes for accessing SC and description of how access is to 

be made 

Drawing to show handling fixtures and how attached to SC (hoists, 

slings, crossbars) 

Page C2-2



C3. INTERFACES 

C3.1 MECHANICAL INTERFACES 


Adapter system Specify which standard adapter system. Reference Appendix D of this 

Mission Planner’s Guide. 

# pushoff springs (if known). Reference Appendix D of this Mission 

Planner’s Guide. 

Specific requirements for mechanical interfaces not provided by above 

standard adapter system 

Coordinate system Drawing showing spacecraft coordinate system 

SC Logo Provide design 

SC interface flange Cross section properties 

I xx 

Drawing showing desired orientation of spacecraft coordinate system 

relative to adapter coordinate system 

Description of constraints on spacecraft orientation 

I yy 

L s = 25 mm 

S 

Scribe mark location 

SC stiffness Minimum fundamental lateral and axial mode frequencies (must be 

greater than 10 and 25 Hz respectively 

SC interface loads Confirm SC lifting device and structure can lift SC+adapter mass = 

200 kg 

SC mass properties Fill in tables in attached Table C3.1-1 and C3.1-2. 

mm 4 

mm 4 

mm 

mm 2 

Hz 

Yes or no 

Fairing access doors Location required for access SC 

coordinates 

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 per 

Figure C3.1-1 

SC Slosh model Provide slosh model of SC during ballistic flight and at separation by 

providing parameters in Figure C3.1-2. 

Propellant tank Provide general propellant tank geometry per Figure C3.1-3 

Page C3-3



C3.2 ELECTRICAL INTERFACES 


Electrical connector Confirm type of connectors 

LVIJ1 

LVIJ2 

Confirm location and keying 

LVIJ1 

LVIJ2 

Confirm quantities to be supplied 

LVIJ1 

LVIJ2 

LVIP1 

LVIP2 

Provide pin locations of SC separation jumper loops 

LVIJ1 pins: 

LVIJ2 pins: 

Current flow 20 seconds or less prior to separation. mA (must 

be less 

than 100) 

Dry loop commands Dry loop commands required? Yes/no 

Characteristics if yes: 

# commands 

Desired pin locations 

LVIJ1 pins: 

LVIJ2 pins: 

Current 

Voltage 

Pulse duration 

Time during flight 

mA 

V 

ms 

sec prior 

to 

separation 

SC telemetry processing LV processing of SC telemetry required? Yes/no 

Characteristics of if yes: 

Desired pin locations 

LVIJ1 pins: 

Current through 

umbilicals at liftoff 

LVIJ2 pins: 

Voltage 

Current 

Data rate 

V 

mA 

Hz 

Current flow 5 minutes or less prior to liftoff. mA (must 

be less 

than 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, -3db 

beamwidth 

Umbilical wiring Fill out Table C3.2-2 with pin assignments and desired line 

characteristics 

Page C3-4



C3.3 ENVIRONMENTAL INTERFACES 


Thermal requirements Provide any particular ground thermal requirements 

Provide any particular thermal requirements during ascent or parking 

orbit 

Provide any particular thermal requirements during transfer orbit 

including: 

sun angle vs. time 

Provide any particular thermal requirements during final injection 

orbit 

Venting analysis Provide archimedes volume of SC in launch configuration (for 

venting analysis) 

SC Testing Provide confirmation of compliance with Planner’s Guide test 

requirements for acoustic, sine, static testing and for the 

fitcheck/shocktest 

EMC Provide confirmation of compliance with EM Susceptibility curve in 

LV in Planner’s Guide 

Provide confirmation of acceptability of LV radiated emissions curve 

in Planner’s Guide 

Humidity Provide humidity requirements for ground transportation 

Provide humidity requirements for processing facility 

Page C3-5



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 

Launch Window Define constraints on launch window 

Provide target launch date 

Provide launch window at perigee passage for one year covering the 

contractual launch date. Include open and close times in GMT for 

each day. 

Separation Define type of separation 3 axis 

stabilized, 

spinning 

or 

transverse 

spin 

Define desired separation attitude with a diagram showing the 

pointing vector in SC coordinates and the pointing vector relative to 

relative right ascension and declination as defined in the Planner’s 

Guide. 

Define desired spin rate about each SC axis. 

Define desired spin rate tolerance about each SC axis 

kg 

km 

km 

deg. 

deg. 

deg. 

deg. 

deg. 

Define desired separation velocity m/sec. 

Page C3-6



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 and 

gases 

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 C3-7



C3.6 OPERATIONS (Continued) 


Transportation Provide description of all items including propellant to be shipped 

including: 

Item name 

Quantity 

Weight 

Tiedown method 

Storage requirements 

Handling Provide description of items which require physical handling at the 

launch site including: 

Equipment name/location 

Dimensions 

Weight 

Handling method 

Communications Define number and type of international lines required. Up to 3 64 

kbps lines can be provided. 

Ground Electrical 

Interfaces 

If multiplexer provided, provide characteristics of and desired 

location. 

Provide locations of all intersite data transmissions including data type 

and rate and interface (analog modem, V.35 or RS232 interface) 

Provide requirements for hardline data transmission between Area 81 

Vault and Bunker, Bldg 44 Rm 58 and Bldg 40D Rm 120 and/or 

between Bldg 40D Rm 301 and Room 120. 

Provide block diagrams of desired umbilical interfaces between SC 

and EGSE while being processed and while on the pad 

Feedthroughs Provide description of feedthroughs required between Control Rm of 

92A-50 and Hall 103A or between Rm 58 of Bldg 44 and Hall 1, 

including: 

Feedthrough designation 

Cable designation 

Cable connector dia. 

Cable dia. 

kg 

m 

kg 

mm 

mm 

Page C3-8



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 to 

be 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-M 2 ) 

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-M 2 ) 

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 manufacturers 

final mass properties report 

d) Above data based on the SC manufacturers Mass Properties Report dated TBD. 

Page C3-9



Table C3.1-1c: SC Mass Properties(Dry Spacecraft) 




c.g.(kg-M 2 ) 

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 C3-10



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 


Moments of inertia relative to propellant c.g. 

(kg-M 2 ) 

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-M 2 ) 



+/- TBD TBD TBD TBD TBD % TBD % TBD % TBD % TBD % TBD % 



Note: Z coordinate relative to separation plane 

c) 1g (During Transportation) 

Mass 

(kg) 




(kg-M 2 ) 






Note: Z coordinate relative to separation plane 

Page C3-11



Figure C3.1-1: SC Pendulum Model 

1. 0g 

2. 1g 

L 

m s 

a s 

x s 

Pendulum Model Parameters: 

Tank m s, Kg L, m x s, m a s, m , % m 0, Kg x 0, m 

F 

O 

Tank m s, Kg L, m x s, m a s, m , % m 0, Kg x 0, m 

F 

O 

Filling Parameters: (t = 20 O C) 

Tank Mass, Kg Level, m Filling, % Density, Kg/m 3 Kinematic viscosity, m 2 /s 

F 

O 

a s - Pivot point of the pendulum from the tank bottom, m 

L - Length of the pendulum, m; 

x s - Slosh mass location from the tank bottom, m; 

m s - Mass of the pendulum, Kg; 

- Damping, %. 

m 0 - Mass of fixed liquid, kg; 

x 0 - Coordinate of mass m 0 from the tank bottom, m. 

Page C3-12



Figure C3.1-2: SC Slosh Properties During Ballistic Flight and at Separation 

dω 

I g + M g C × 

dt 

2 

dω d l 

- M g C × + M g 2 

dt dt 

2 

d ai 

dω 

= - a 

2 

i 

dt dt 

2 

d l 

A i = + 

2 

dt 

g 

dt 

s 

∑ 

2 

d l 

T - × 

2 

dt 

ω ( I ω ) ( × F ) 

g 

= F - × 

a 

× - 2 ω × 

dt 

dω 

d i 

× r + × 

oi 

M = M + m ( 1 − G ) 

i 

M g C = s s r M + mi ( 1 − Gi 

) 

I = I - m ( 1 − G ) 

g 

s 

( r ) 

x 

i 

S = S y 

oi ⎢ i ⎥ 

⎣ z 

i ⎦ 

i 

i 

i 

∑ 

ω ( × C) 

i 

R oi CT ⋅i 

∑ 

ω M g - F CT ⋅i 

i 

- ω × ( ω × a i ) - i A + F / mi 

ω ( ω × r ) 

∑ r oi 

∑ 

i 

i i ( r ) oi 

⎡ 

⎤ 

2( 

1− 

Ki 

) 

Gi = 

1+ 

( 1− 

K ) 

1 

ri= 

1 3 1− 

K 1 

d 

i 

− 

i ÷ 4 R 

i 

i 

oi 

S S( 

r ) 

0 z ⎡ i 

− y 

i ⎤ 

= − z ⎢ i 

0 x 

i ⎥ 

⎣ y 

i 

− x 

i 

0 ⎦ 

oi 

MINIMUM 

G i CT ⋅i 

ai 

R = r + 

oi oi 

G 

F CT ⋅ i = 

• β i ( a) 

β i ( ) 

3H 

i 

( )( ) a ∈[ 

−0. 

5; 

+ 0. 

5] 

β i( 

a) 

= −d0i 

d0i 

( Ri 

− ri 

) 

β ( a) 

i 

i 

⎧ 

0 

⎨ 

⎩ 

− k pr d i − k demV 

− k pi shearV 

ti 

1 − K 

K 

a i K i 

= − 

1 − K 

⎡1⎤ 

+ aH 

⎢ ⎥ 

⎢ 

0 

⎥ 

⎢⎣ 

0⎥⎦ 

a i 

i 

⎡ 

⎢1 

⎢⎣ 

= 

i 

−3 

⎜ 

⎛ 

⎝ 

i 

i 

d i • 

dt 

da 

i 

V Pi = • d i I V ti 

= −V 

Pi 

i • i 

dt 

k shear 

d 

da 

d 

1 − K 

= 170229N 

/ m k 16256 N ( m / sec) 

shear = k 354. 3 N ( m / sec) 

shear = 

⎛ 

⎜ 

⎟ 

⎞⎜ 

i ⎠⎜ 

⎜ 

⎝ 

3H 

1 + 

4R 

r < R 

i 

i 

i 

r ≥ R 

⎞ 

⎟ 

i 

⎟ 

⎟ 

i ⎠ 

⎤ 

⎥ 

⎥⎦ 

i 

Page C3-13



NOMENCLATURE 

ω 

the spacecraft angular rate vector; 

l the vector determining the position of the coordinate origin fixed in the spacecraft body in the inertial 

coordinate system; 

C 

T 

F- 

a i 

V pi,V ti 

M s, I s 

R i 

r s 

m i 

r oi 

R oi 

A i 

K i 

H i 

the vector determining the spacecraft mass center position; 

the sum of external moments with respect to the selected origin of coordinates; 

the sum of external forces acting on the spacecraft; 

position vector of fluid center of mass of i-th tank about center of this tank; 

dai the radial and tangential components of velocity of the liquid mass center in the i - tank; 

dt 

the mass and matrix of inertia of the dry spacecraft (without liquid components); 

the i - tank radius; 

the vector determining the dry spacecraft mass center position in the i - tank with respect to the origin of 

coordinates; 

the liquid mass in the i - tank; 

the vector drawn from the origin of coordinates to the i - tank center; 

the vector determining the liquid mass center position in the i - tank with respect to the origin of coordinates; 

the vector of inertial acceleration of the i - tank; 

the i - tank filling factor 

K i=V filling 

V total 

the i - tank cylinder length 

H i=0 corresponds to spherical tank 

where: V filling is the volume of poured fuel component; 

V total is the total volume of the tank. 

Page C3-14



Figure C3.1-3: Propellant Tank Geometry Required Data 

TBD 

rTBD 

TBD 

-X 

SC/Upper Stage Separation Plane 

sc 

TBD 

TB 

D 

TBD 

Liquid Level 

Page C3-15



Table C3.2-1: SC RF Characteristics 

Transmitter 1 Transmitter 2 Receiver 1 Receiver 2 

Description TM Transmitter 1 TM Transmitter 2 Command Receiver 

1A 

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 


characteristics 

TBD TBD TBD TBD 

Intermediate frequency (Mhz) TBD TBD 

Local Oscillator frequencies 

(Ghz) 

Transmit antenna output power 

EIRP (dbW) 

Max (on-orbit setting) 

Max (normal on pad setting) 

Nom 

Min 

Receive antenna receive flux 

density (dbW/M 2 ) 

Max 

Nom 

Min 

Antenna description, polarization, 

pattern 

Antenna location 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD TBD 

TBD 

TBD 

TBD 

Command Receiver 

1B 

TBD 

TBD 

TBD 

TBD(c) TBD(c) TBD(c) TBD(c) 

Operating on pad? TBD (e) TBD (e) TBD TBD 

Operating in flight? TBD TBD TBD TBD 

Ground equipment output power 

(dbm) 

Max 

Nom 

Min 

Ground equipment receive power 

(dbm) 

Max (see note f) 

Nom 

Min 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

TBD 

Page C3-16



Notes: [Sample] 

a) TBD RF command links and TBD RF telemetry links are required. The SC Checkout Station will have 1 physical command 

interface 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 window 

obtained 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 SC 

manufacturer shall verify continuity between command RF link and STE and issue to Khrunichev the Certificate of Launch 

Pad Readiness to accommodate the LV and Ascent Unit. The spectrum analyzer to be provided by the SC manufacturer shall 

be 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 with 

Khrunichev, shall checkup the RF link between the Ascent Unit and STE. Such checkup shall be performed 20 minutes after 

the 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 the 

existing RF cables in the bunker, two spare connectors and two corresponding jacks, as well as instructions on cable dressing 

and cable performances. (If required) 

Page C3-17



Table C3.2-2a: LVIJ1 Umbilical Pin Assignments 

Connector 

Pin 

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 

31 

32 

33 

34 

35 

36 

37 

38 

39 

40 

41 

42 

43 

44 

45 

Function Max. 

Voltage 

(V)* 

Max. 

Current 

(A) 

Max. 

Voltage at 

Liftoff (V) 

Max. 

Current at 

Liftoff (A) 

Line 

Resistance 

(ohms) (4) 

Time of 

Usage 

Page C3-18



Connector 

Pin 

Notes: 

46 

47 

48 

49 

50 

51 

52 

53 

54 

55 

56 

57 

58 

59 

60 

61 

Function Max. 

Voltage 

(V)* 

Max. 

Current 

(A) 

Max. 

Voltage at 

Liftoff (V) 

Max. 

Current at 

Liftoff (A) 

Line 

Resistance 

(ohms) (4) 

Time of 

Usage 

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 provides 

continuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power through 

the 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 the 

vault 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 C3-19



Table C3.2.2b: LVIJ2 Umbilical Pin Assignments 

Connector 

Pin 

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 

31 

32 

33 

34 

35 

36 

37 

38 

39 

40 

41 

42 

43 

44 

45 

Function Max. 

Voltage 

(V)* 

Max. 

Current 

(A) 

Max. 

Voltage at 

Liftoff (V) 

Max. 

Current at 

Liftoff (A) 

Line 

Resistance 

(ohms) (4) 

Time of 

Usage 

Page C3-20



Connector 

Pin 

Notes: 

46 

47 

48 

49 

50 

51 

52 

53 

54 

55 

56 

57 

58 

59 

60 

61 

Function Max. 

Voltage 

(V)* 

Max. 

Current 

(A) 

Max. 

Voltage at 

Liftoff (V) 

Max. 

Current at 

Liftoff (A) 

Line 

Resistance 

(ohms) (4) 

Time of 

Usage 

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 provides 

continuous monitoring and recording of this bus voltage on pins TBD and TBD of umbilicals P1 and P2 . SC power through 

the 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 the 

vault 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 C3-21



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 C3-22



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 C3-23



Table C3.5.1-1: SC Contractor Electrical Ground Support Equipment (Continued) 

Connectors 


Bldg 44 Rm 58 

BLDG. 92-1 

LAUNCH 

COMPLEX 81 

Bunker (Rm. 250) 

Bunker (Rm 251) 

Bunker (Rm 244) 

Vault (Rm. 64 or 76)) 

Page C3-24



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 Hall 

Fueling Hall 

demineralized water See Launch Campaign Guide for max available KhSC/ILS Decontamination 

Area 

Nitrogen GOST-92-93-74, 

technical grade 1 

Nitrogen GOST-92-93-74, 

technical grade 1 

See Launch Campaign Guide for max available. KhSC/ILS Decontamination 

Area 

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-27407 

Type 1, Grade A 

TBD K-bottles high pres (400 bar) TBD Kbottles 

low pres (135 bar) 

MMH TBD Cylinders - Tot weight ea TBD kg max, 

TBD kg max prop weight 

SC contractor Fueling Hall 


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 Grade 

GOST 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 C3-25



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 payload 

adapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interface 

characteristics, 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 and 

fabricated from aluminum alloy. Mechanical drawings for the 1194AX-500 adapter are presented in Section D1.7 of 

this 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 load 

path between the spacecraft and adapter during ground operations and flight. The outboard features of the combined 

cross section are designed to interface with the separation system clamp band. The cross section and material property 

characteristics for the spacecraft and adapter interface ring are presented in Table D1.2.1-1 and Figure 

D1.2.1-1. 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 mm 2 ± 15% 558 mm 2 

Moment Of Inertia (Ixx) 52000 mm 4 ± 15% 72000 mm 4 

Moment Of Inertia (Iyy) 13400 mm 4 ± 15% 30000 mm 4 

Young Modulus (E) 69 x 10 3 MPa 70 x 10 3 MPa 

Page D1-1



Figure D1.2.1-1: Spacecraft and Adapter Interface Ring Cross Section 

Spacecraft Ring 

y 

y 

Adapter Ring 

x 

x 

x 

x 

y 

y 

L 

Separation Plane 


Page D1-2



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 this 

Appendix. 

D1.2.2 Structural Capability 

The structural capability of each adapter system is based on the allowable line load obtained from testing. The 

relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1194AX-500 adapter system is 

presented in Figure D1.2.2-1. These structural capabilities assume the standard cylindrical interface ring stiffness 

characteristics presented in Table D1.2-1, the geometry presented in Section D1.7 of this Appendix and the quasistatic 

design load factors presented in Section 3.4.1. Additionally, the line loading at the interface has been calculated 

by classical plane section assumptions. Distortion of the interface plane producing peaking of line loading will reduce 

the allowable CG offset for a given spacecraft mass. This may result from spacecraft primary loads reacted as point 

loads through the spacecraft structure close to the interface. 

The structural capability presented should only be used as a guideline for assessment of interface structural 

compatibility. Coupled loads analysis performed early in the mission integration will verify margins for structural 

loading of the adapter and separation system. 

Page D1-3



Figure D1.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC) 

Longitudinal Offset of SC C.G. from 


3.50 

3.00 

2.50 

2.00 

1.50 

1.00 

0.50 

0.00 

26.6kN Clamp Band Tension 

40kN Clamp Band Tension 

2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600 

SC Mass (kg) 

Preliminary 

SC Mass 

Allowable C.G. Offset (m) 

(kg) 40kN Tension 26.6kN Tension 

2000 3.07 2.30 

2200 2.79 2.09 

2400 2.56 1.92 

2600 2.37 1.78 

2800 2.20 1.65 

3000 2.06 1.55 

3200 1.93 1.45 

3400 1.82 1.37 

3600 1.72 1.29 

3800 1.63 1.23 

4000 1.55 1.17 

4200 1.48 1.11 

4400 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 

Page D1-4



D1.3 USABLE VOLUME 

The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-500 adapter system is 

presented in Figures D1.3-1a through D1.3-1e. The spacecraft static envelope (maximum dimensions of unloaded 

spacecraft, including manufacturing tolerances and expansion of thermal blankets) must not protrude beyond the 

useable volume, except where it is mutually agreed upon by ILS and Khrunichev. Spacecraft dynamic displacements 

due to ground or flight loads and deviations caused by an imperfect installation of the spacecraft on the Block DM 

may protrude beyond the boundaries of this useable volume. It is assumed that spacecraft dynamic displacements will 

not exceed 50 mm. This must be verified by coupled loads analysis. 

Page D1-5



Figure D1.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter 

Page D1-6



Figure D1.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 1 of 4) 

Page D1-7



Figure D1.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 2 of 4) 

Page D1-8



Figure D1.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 3 of 4) 

Page D1-9



Figure D1.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 500mm Adapter (Sheet 4 of 4) 

Page D1-10



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 band 

are preloaded using a precision hydraulic tensioning device and secured by two bolts. The nominal clamp band preload 

is 40.0 kN for the1194AX 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 

retention device moves the clamp band away from the interface ring. The retention device also secures the band to the 

adapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and 

adapter. 

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 an 

initial force of 1500 Newtons. The stroke for each spring is selectable to customize the total energy per spring required 

to provide the desired separation velocity and separation rates. Table D1.4-1 defines the maximum and minimum 

spring stroke range with the associated spring force characteristics. Spring sets can include any number between two 

and twelve. 

Table D1.4-1: Separation Spring Characteristics 

Stroke 

(mm) 

Initial Force 

(N) 

Final Force 

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 of 

electrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electrical 

disconnects and grounding connectors are provided as shown in Section D1.7 of this Appendix. This symmetrical 

arrangement is provided to minimize overturning moments at separation. The force profile for each electrical 

disconnects is shown in Figure D1.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resists 

separation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D1.4-2. 

(N) 

Page D1-11



Figure D1.4-1: Electrical Disconnect Force Profile for a Single 37 Pin Electrical Connector 

Force (N) 

200 

150 

100 

50 

0 

-50 

-100 

-150 

0 1 2 3 4 5 6 7 8 9 10 

Force Displacement 

(N) (mm) 

178 0 

156 3.4 

-115 3.4 

-115 6.7 

0 6.7 

Note: A positive force assists separation. 

Displacement (mm) 

Page D1-12



Figure D1.4-2: Purge Connector Force Diagram (Typical) 

Force (N) 

100 

80 

60 

40 

20 

0 

0.00 0.10 0.20 0.30 0.40 0.50 0.60 

Axial Travel (mm) 

Page D1-13



D1.5 ELECTRICAL INTERFACE 

In order to accommodate command and control signals to the spacecraft, two electrical connectors are provided. These 

electrical connectors provide a spacecraft dedicated umbilical from the spacecraft to the ground support equipment and 

the 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 electrical 

interfaces configurations are presented in Section D1.7 of this Appendix. Connectors conforming to MIL-C-81703 

have been used for both adapter system configurations. Part numbers for the standard 37 pin connectors are presented 

in 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 launch 

vehicle separation indicators. Loop backs on the launch vehicle side of the interface for indication of separation for the 

spacecraft 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 electrical 

connectors and GSE. This information includes available wire types, shielding, voltage requirements, current 

requirement, and resistance requirements. 

The requirement for maximum resistance across the separation interface is 10 milliohms to ensure electrical continuity 

across the separation interface. Electrical continuity across the separation interface is provided by two diametrically 

opposed 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. The 

standard configuration includes 5 accelerometers; 3 oriented to monitor longitudinal accelerations and 2 oriented to 

monitor transverse accelerations. The installation of the accelerometers on the adapter is presented in Section D1.7 of 

this Appendix. The characteristics of these adapter mounted accelerometers and for all of the telemetry channels for 

the Proton mission are presented in Section 4.2.1.7 of the Mission Planner’s Guide. 

Page D1-14



D1.7 1194AX-500 ADAPTER MECHANICAL DRAWINGS 

Page D1-15



1194AX-500 Adapter Mechanical Drawings (Continued 2 of 9) 

TBS 

Page D1-16




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



D2. 1194AX-625 ADAPTER SYSTEM 


The 1194AX-625 adapter system is comprised of the 1194 AX clamp band system, separation springs, a payload 

adapter and electrical rise off disconnects. This appendix defines the mechanical and electrical interface 

characteristics, 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 fabricated 

from aluminum alloy. Mechanical drawings for the 1194AX-625 adapter are presented in Section D2.7 of this 

Appendix. 









Height Of Effective Cross Section(L) 25 mm 25 mm 





Page D2-1





y 

y 

Adapter Ring 

x 

x 

x 

x 

y 

y 

L 



Page D2-2





Appendix. 



relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the 1194AX-625 adapter system is 










Page D2-3



Figure D2.2.2-1: Capability of 1194AX Adapter System - SC Mass vs Longitudinal C.G. (TBC) 



3.50 

3.00 

2.50 

2.00 

1.50 

1.00 

0.50 

0.00 

26.6kN Clamp Band Tension 

40kN Clamp Band Tension 

2000 2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600 

SC Mass (kg) 

Preliminary 

SC Mass 

Allowable C.G. Offset (m) 

(kg) 40kN Tension 26.6kN Tension 

2000 3.07 2.30 

2200 2.79 2.09 

2400 2.56 1.92 

2600 2.37 1.78 

2800 2.20 1.65 

3000 2.06 1.55 

3200 1.93 1.45 

3400 1.82 1.37 

3600 1.72 1.29 

3800 1.63 1.23 

4000 1.55 1.17 

4200 1.48 1.11 

4400 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 

Page D2-4




The usable volume for the spacecraft encapsulated in the Proton fairing with the 1194AX-625 adapter system is 

presented in Figures D2.3-1a through D2.3-3e for the Block DM and Breeze M versions of the LV. The spacecraft 

static envelope (maximum dimensions of unloaded spacecraft, including manufacturing tolerances and expansion of 

thermal blankets) must not protrude beyond the useable volume, except where it is mutually agreed upon by ILS and 

Khrunichev. Spacecraft dynamic displacements due to ground or flight loads and deviations caused by an imperfect 

installation of the spacecraft on the Fourth Stage may protrude beyond the boundaries of this useable volume. It is 

assumed that spacecraft dynamic displacements will not exceed 50 mm. This must be verified by coupled loads 

analysis. 

Page D2-5



Figure D2.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter 

Page D2-6



Figure D2.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) 

Page D2-7



Figure D2.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) 

Page D2-8



Figure D2.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) 

Page D2-9



Figure D2.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) 

Page D2-10



Figure D2.3-2a: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter 

Page D2-11



Figure D2.3-2b: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC) 

Page D2-12



Figure D2.3-2c: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC) 

Page D2-13



Figure D2.3-2d: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC) 

Page D2-14



Figure D2.3-2e: Usable Volume - Proton/Breeze M Standard Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC) 

Page D2-15

Proton Mission Planners Guide, LKEB-9812-1990 


Figure D2.3-3a: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter 

Page D2-16



Figure D2.3-3b: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 1 of 4) (TBC) 

Page D2-17



Figure D2.3-3c: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 2 of 4) (TBC) 

Page D2-18



Figure D2.3-3d: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 3 of 4) (TBC) 

Page D2-19



Figure D2.3-3e: Usable Volume - Proton/Breeze M Long Commercial Fairing with 1194AX X 625mm Adapter (Sheet 4 of 4) (TBC) 

Page D2-20






is 26.6 or 40.0 kN for the 1194AX 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 

retention device moves the clamp band away from the interface ring. The retention device also secures the band to the 

adapter and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and 

adapter. 

The shock environment for the separation event is provided in Section 3.4.4 of this Mission Planners Guide. 





and twelve. 


Stroke 

(mm) 

Initial Force 

(N) 

Final Force 




electrical disconnects and grounding connectors and, if provided as an option, the purge disconnect. The electrical 

disconnects and grounding connectors are provided as shown in Section D2.7 of this Appendix. This symmetrical 

arrangement is provided to minimize overturning moments at separation. The force profile for each electrical 

disconnects is shown in Figure D2.4-1. Each grounding connectors imparts a force of 40 ± 5 Newtons that resists 

separation. The purge fitting imparts a force that assists separation. This force profile is provided in Figure D2.4-2. 

(N) 

Page D2-21




Force (N) 

200 

150 

100 

50 

0 

-50 

-100 

-150 

0 1 2 3 4 5 6 7 8 9 10 


(N) (mm) 

178 0 

156 3.4 

-115 3.4 

-115 6.7 

0 6.7 



Page D2-22



Figure D2.4-2: Purge Connector Force Diagram (Typical) 

Force (N) 

100 

80 

60 

40 

20 

0 

0.00 0.10 0.20 0.30 0.40 0.50 0.60 

Axial Travel (mm) 

Page D2-23













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 electrical 


requirement, and resistance requirements. 

The requirement for maximum resistance across the separation interface is of 10 milliohms to ensure electrical 

continuity across the separation interface. Electrical continuity across the separation interface is provided by two 

diametrically opposed electrical grounding connector as presented in Section D2.7 of this Appendix. 






the Proton mission are presented in Section 4.2.1.7 of the Mission Planners Guide. 

Page D2-24



D2.7 1194AX-625 ADAPTER MECHANICAL DRAWINGS 

Page D2-25



1194AX-625 Adapter Mechanical Drawings(Sheet 2 of 9) 

TBS 

Page D2-26




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



D3. 1666V-1000 ADAPTER SYSTEM 


The 1666V-1000 adapter system is comprised of the 1666 V clamp band system, separation springs, a payload adapter 

and 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 fabricated 

from aluminum alloy. Mechanical drawings for the 1666V-1000 adapter are presented in Section D3.7 of this 

Appendix. 














Page D3-1





y 

y 

Adapter Ring 

x 

x 

x 

x 

y 

y 

L 



Page D3-2





Appendix. 



relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666V-1000 adapter system is 










Page D3-3



Figure D3.2.2-1: Capability of 1666V Adapter System - SC Mass vs Longitudinal C.G. (TBC) 



3.50 

3.00 

2.50 

2.00 

1.50 

1.00 

0.50 

0.00 

2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600 

SC Mass (kg) 

Preliminary 

Allowable C.G. 

SC Mass Offset (m) 

(kg) 30.0kN Tension 

2000 3.39 

2200 3.09 

2400 2.83 

2600 2.62 

2800 2.44 

3000 2.28 

3200 2.14 

3400 2.02 

3600 1.89 

3800 1.77 

4000 1.66 Note: A positive force assists separation. 

4200 1.56 

4400 1.47 

4600 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 N c = 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 

Page D3-4




The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666V-1000 adapter system is 







Page D3-5



Figure D3.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter 

Page D3-6



Figure D3.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 1 of 4) 

Page D3-7



Figure D3.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 2 of 4) 

Page D3-8



Figure D3.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 3 of 4) 

Page D3-9



Figure D3.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1666V X 1000mm Adapter (Sheet 4 of 4) 

Page D3-10






is 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 retention 

device moves the clamp band away from the interface ring. The retention device also secures the band to the adapter 

and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter. 






and twelve. 


Stroke 

(mm) 

Initial Force 

(N) 

Final Force 




electrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects and grounding 

connectors are provided pairs as shown in Section D3.7 of this Appendix. This symmetrical arrangement is provided to 

minimize overturning moments at separation. The force profile for each electrical disconnects is shown in Figure 

D3.4-1. 

(N) 

Page D3-11




Force (N) 

400 

300 

200 

100 

0 

-100 

-200 

0 1 2 3 4 5 6 7 8 9 10 


(N) (mm) 

300 0 

281 3.4 

-156 3.4 

-156 6.7 

0 6.7 



Page D3-12








interfaces configurations are presented in Section D3.7 of this Appendix. Connectors conforming to MIL-C-81703 are 

used for both adapter system configurations. Part numbers for the standard 61 pin connectors are presented in Table 

D3.5-1. 



Spacecraft Side LVIJ1 61 MS3424E61-50S 

LVIJ2 61 MS3424E61-50S 

Launch Vehicle Side LVIP1 61 MS3446E61-50P 

LVIP2 61 MS3446E61-50P 






requirement and resistance requirements. 


across the separation interface. Electrical continuity across the separation interface is provided by conductive coatings 

on both the spacecraft and adapter interface flanges. 







Page D3-13



D3.7 1666V-1000 ADAPTER MECHANICAL DRAWINGS 

Page D3-14



1666V-1000 Adapter Mechanical Drawings(Sheet 2 of 9) 

TBS 

Page D3-15




TBS 

Page D3-16




TBS 

Page D3-17




TBS 

Page D3-18




TBS 

Page D3-19




TBS 

Page D3-20




TBS 

Page D3-21




TBS 

Page D3-22



D4. 1666A-1150 ADAPTER SYSTEM 


The 1666A-1150 adapter system is comprised of the 1666 A clamp band system, separation springs, a payload adapter 

and 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 fabricated 

from aluminum alloy. Mechanical drawings for the 1666A-1150adapter are presented in Section D4.7 of this 

Appendix. 





characteristics for the spacecraft and adapter interface ring are presented in Table D4.2.1-1 and Figure D4.2.1-1. The 

dimensions of the spacecraft interface are presented in Section D4.7 of this Appendix. 








Page D4-1





y 

y 

Adapter Ring 

x 

x 

x 

x 

y 

y 

L 



Page D4-2





Appendix. 



relationship between spacecraft mass and longitudinal center of gravity (C.G.) for the1666A-1150 adapter system is 










Page D4-3



Figure D4.2.2-1: Capability of 1666A Adapter System - SC Mass vs Longitudinal C.G. (TBC) 



3.50 

3.00 

2.50 

2.00 

1.50 

1.00 

0.50 

0.00 

2200 2400 2600 2800 3000 3200 3400 3600 3800 4000 4200 4400 4600 4800 5000 5200 5400 5600 

SC Mass (kg) 

Preliminary 

Allowable C.G. 

SC Mass Offset (m) 

(kg) 30.0kN Tension 

2000 3.39 

2200 3.09 

2400 2.83 

2600 2.62 

2800 2.44 

3000 2.28 

3200 2.14 

3400 2.02 

3600 1.89 

3800 1.77 

4000 1.66 Note: A positive force assists separation. 

4200 1.56 

4400 1.47 

4600 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 N c = 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 

Page D4-4




The usable volume for the spacecraft encapsulated in the Proton fairing with the 1666A-1150adapter system is 







Page D4-5



Figure D4.3-1a: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter 

Page D4-6



Figure D4.3-1b: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 1 of 4) 

Page D4-7



Figure D4.3-1c: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 2 of 4) 

Page D4-8



Figure D4.3-1d: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 3 of 4) 

Page D4-9



Figure D4.3-1e: Usable Volume - Proton/Block DM Commercial Fairing with 1166A X 1150 Adapter (Sheet 4 of 4) 

Page D4-10






is 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 retention 

device moves the clamp band away from the interface ring. The retention device also secures the band to the adapter 

and prevents rebound or recontact. Two separation indicators verify separation between the spacecraft and adapter. 






and twelve. 


Stroke 

(mm) 

Initial Force 

(N) 

Final Force 




electrical disconnects and, if provided as an option, the purge disconnect. The electrical disconnects connectors are 

provided pairs as shown in Section D4.7 of this Appendix. This symmetrical arrangement is provided to minimize 

overturning moments at separation. The force profile for each electrical disconnect is shown in Figure D4.4-1. 

(N) 

Page D4-11




Force (N) 

400 

300 

200 

100 

0 

-100 

-200 

0 1 2 3 4 5 6 7 8 9 10 


(N) (mm) 

300 0 

281 3.4 

-156 3.4 

-156 6.7 

0 6.7 



Page D4-12













Spacecraft Side LVIJ1 61 MS3424E61-50S 

LVIJ2 61 MS3424E61-50S 

Launch Vehicle Side LVIP1 61 MS3446E61-50P 

LVIP2 61 MS3446E61-50P 






requirement and resistance requirements. 


across the separation interface. Electrical continuity across the separation interface is provided by conductive coatings 

on the spacecraft and adapter interface flanges. 







Page D4-13



D4.7 1666A-1150 ADAPTER MECHANICAL DRAWINGS 

Page D4-14



1666A-1150 Adapter Mechanical Drawings(Sheet 2 of 9) 

TBS 

Page D4-15




TBS 

Page D4-16




TBS 

Page D4-17




TBS 

Page D4-18




TBS 

Page D4-19




TBS 

Page D4-20




TBS 

Page D4-21




TBS 

Page D4-22

proton launch vehicle mission planner's guide

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